1623results about "Artificial satellites" patented technology

The invention provides a lunar satellite formation autonomous navigation method based on inter-satellite ranging information. The lunar satellite formation autonomous navigation method comprises three parts of orbit dynamics model establishment, collaborative observation model establishment and navigation filtering algorithm data fusion. The invention provides an autonomous navigation method based on inter-satellite distance measurement and angle measurement information, which establishes a high-precision dynamic model containing 165-order moon non-spherical perturbation for moon satellite formation. The system can operate independently of a ground station, and high-precision autonomous navigation of the lunar satellite formation is realized only by using on-orbit observation data.

The invention provides a lunar satellite formation orbit determination method and system. The method comprises the following steps: S1, establishing a lunar satellite high-precision dynamic model; s2, establishing a lunar satellite laser ranging and inter-satellite ranging mathematical model; s3, measuring the distance, namely the absolute position, of the A-type satellite relative to the ground station through laser ranging, and measuring the relative position of the B-type satellite relative to the A-type satellite through inter-satellite ranging; s4, preprocessing the absolute position and relative position data of the satellite measured by laser ranging and inter-satellite ranging, and generating a format data file required by orbit calculation; and S5, resolving the position coordinates of each satellite in the monthly formation through the format data file. According to the method, laser ranging and inter-satellite ranging are combined for orbit determination, precise orbit determination of the lunar satellite formation is achieved, compared with a traditional orbit determination mode, the precision can be improved from the hundred-meter level to the meter level, the requirements for the arrangement number of ground stations and set distribution are reduced, and the measurement and control pressure of existing ground stations is relieved.

The invention discloses a service spacecraft multi-orbit maneuvering strategy approaching observation method based on J2 perturbation compensation, solves the problem of how to quickly improve orbit maneuvering accuracy, and belongs to the field of service spacecraft multi-orbit maneuvering strategy approaching observation. Comprising the following steps: constructing a motion equation under J2 item perturbation compensation under a target orbit coordinate system; constructing a service spacecraft multi-orbit maneuvering strategy model through the relative motion propagation relational expression and the motion equation; inputting the termination moment position conforming to the termination moment distance constraint into the model to output a single speed increment, and reserving the single speed increment conforming to the termination moment illumination angle constraint and the single speed increment constraint at the same time to obtain a speed increment solution set; constructing a circular region which takes the space target as a center and has a radius meeting a termination moment distance constraint; when the service spacecraft arrives at the circular area at a specified moment, marking the task successfully, and obtaining a service spacecraft multi-orbit maneuvering strategy for approaching observation; according to the invention, accurate and rapid track maneuvering is realized.

The invention relates to a simulation-based satellite scheduling task dynamic planning method, device and equipment, and the method comprises the steps: carrying out the modeling according to various time-varying environment dynamic parameters in an environment where a satellite is located, and obtaining a dynamic environment model which comprises an environment factor model and an emergency triggering mechanism; the method comprises the following steps: constructing a satellite resource dynamic model according to various resources of satellites, resource consumption during task execution and resource supply, constructing a multi-satellite collaborative model according to an interaction relationship among a plurality of satellites and related information of each satellite, and constructing a multi-satellite collaborative model under a dynamic environment model, the satellite resource dynamic model and the multi-satellite collaborative model. According to the method, tasks in a to-be-executed task data set containing single-satellite and cooperative tasks are dynamically scheduled by adopting a task planning algorithm, a task execution planning table of satellites is obtained, satellite task planning can be carried out aiming at emergencies and multi-satellite cooperative tasks, and the planning is closer to the actual situation.

The invention provides a winding type flexible solar wing unfolding mechanism and a star body. The winding type flexible solar wing unfolding mechanism comprises a solar wing driving mechanism, a hinge, a rigid solar cell array, a flexible wing, a first pressing and releasing device and a second pressing and releasing device. The solar wing unfolding mechanism has the advantages of being small in folding size, light in weight, capable of being unfolded in orbit and large in folding-unfolding ratio. In addition, the structure is simple, the unfolding driving force source is simple, movement is reliable, unfolding driving kinetic energy is provided by releasing winding strain energy, and the unfolding rigidity of the solar cell array is provided after in-orbit unfolding.

According to the present invention, it is possible to improve the effects such as search tasks when searching for an object by a plurality of systems. 【Solution means】The present invention is a control system for searching for an object in a predetermined first area using a plurality of mobile bodies, and is obtained from a cooperation system capable of measuring a second area including at least a part of an area outside the first area. Based on the acquired information, an object information interpretation unit that interprets object information regarding an object existing in the second area, and based on the object information, an entry prediction information generation unit that generates entry prediction information regarding the predicted content of the object entering the first area from an area outside the first area, and the entry prediction information generated by the object movement prediction unit, or information regarding a mobile body movement command for some or all of the plurality of mobile bodies generated by the mobile body movement determination unit based on the entry prediction information, and an information output unit that performs display output or transmits to the outside.

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The invention discloses a modular micro-nano satellite integrated power system suitable for batch production and an assembly method thereof. The integrated power system is composed of a semi-open type electric box structure and a modular circuit board. The electricity box structure comprises an electricity box upper cover and an electricity box lower cover, and a plurality of threaded holes are formed in the two sides of the electricity box upper cover and the two sides of the electricity box lower cover and used for being connected with a satellite structure cabin plate; the heat dissipation plate bodies are distributed on the two sides of the modular circuit board; the modularized circuit board is provided with a plurality of circuit boards which are respectively provided with electronic modules in a stacking arrangement. A plurality of electronic modules are integrated into an electronic system by adopting a mechanical-electrical-thermal integrated design method, the system breaks through the function limit of a traditional satellite subsystem, the defects of information exclusion and resource waste are eliminated, a satellite is used as a whole to be subjected to unified design and resource planning tradeoff from the system level, and the system performance is improved. The weight of cables and connectors between electronic modules is greatly reduced, and the effective load ratio is improved.

The invention discloses a high-performance satellite platform based on a real-time operating system and a wireless bus and a working method of the high-performance satellite platform, and belongs to satellite communication technologies, future deep space exploration and low-orbit satellite constellations. On the basis of an original structure of a satellite platform, a real-time operating system is effectively combined with a wireless bus system, and the real-time operating system comprises a kernel module, a communication module, an equipment driving module, a file system module, an application interface module, a safety module and a fault management module. The wireless bus system comprises a physical layer, a data link layer, a network layer, a transmission layer and an application layer. According to the invention, efficient and reliable communication and task execution capabilities are provided for a satellite platform, the requirements of real-time performance and reliability of a satellite in a complex environment are effectively met through modular design and tight work of the system, and meanwhile, complex lines and interfaces brought by a wired system are avoided.

The invention discloses an ultra-low orbit guarantee method, system and equipment for a near-earth satellite. Comprising the following steps: determining a satellite orbit element at the current moment based on an atmospheric resistance perturbation item, a conventional perturbation item and a small thrust perturbation item, and inputting the satellite orbit element at the current moment into an orbit mechanical model; based on the orbit mechanical model and the satellite orbit elements, determining a first velocity position parameter of the satellite at the prediction moment under the ICRS coordinate system, and converting the first velocity position parameter into an ECEF coordinate system to obtain a second velocity position parameter at the prediction moment; inputting the time characteristic, the satellite state characteristic, the space environment characteristic and the second speed position parameter of the current moment into an orbit correction model to obtain a target error prediction value; adjusting the second speed position parameter based on the target error prediction value to obtain a target speed position parameter of the satellite at the prediction moment; and performing orbit attenuation risk early warning based on the target speed and position parameters of the satellite at the prediction moment.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

The invention relates to a modular multifunctional satellite configuration which comprises an integrated flat plate, a peripheral frame, a reinforcing frame, four bearing columns and externally-hung side cabins, the integrated flat plate and the reinforcing frame are arranged in the peripheral frame, the number of the externally-hung side cabins is 1-4, the number of the bearing columns is 4, multiple satellites can be stacked in a multi-layer mode through the bearing columns, and the satellites have two stacking modes. One is that four bearing columns between the upper and lower satellites are coaxially stacked; according to the second structure, two force bearing columns between the upper-layer satellite and the lower-layer satellite are coaxially stacked, the other two force bearing columns are provided with compensation columns between the upper-layer satellite and the lower-layer satellite, and the height of the compensation columns adapts to the stacking height design of the interlayer satellites. According to the invention, the engineering requirements of different satellite equipment mounting areas, fairing carrying states and a one-rocket multi-satellite launching mode can be met, a cable mounting bracket does not need to be additionally arranged, the cable network mounting process is simplified, and convenience is provided for batch final assembly of the cable network.

The present invention relates to an plant disease detection at onset stage. Provided is a computer-implemented method for determining an onset and / or onset time of a plant (12) disease in agriculture. The method comprises providing (S110) first data including field data (14) associated with the plant's cultivation and weather data (16) associated with a location where said plant is cultivated to a computer model (20). The method further comprises determining (S120), by using said computer model (20), a plant disease presence prediction for said plant and its infestation with said plant disease and determining, from said computer model (20) output including said plant disease presence prediction and second data (18) including one or more vegetation indices associated with said plant, the onset and / or onset time to which said plant disease is expected to onset at said plant, by using the plant disease presence prediction and a change in the one or more vegetation indices.

A modular satellite system is provided comprising a main body member, one or more controllers, a communication system, a datastore, a power unit, and one or more orbital cameras. The controller(s) may be operable to perform a mission instruction. The communication system may be in communication with the controller(s) and with a client terminal. The datastore may store data accessible by the controller(s). The orbital camera(s) may be operable to capture one or more image(s) associated with the mission instruction, which may be defined as captured image(s). The controller(s) may be operable to detect and identify one or more predetermined object(s) in the one or more captured image(s). The controller(s) may generate a mission analytics packet, and the controller(s) may be operable to transmit the mission analytics packet to the client terminal.

In accordance with the present invention, unmanned watercraft can be utilized to perform multiple missions. [Solution] The present invention is a control system for controlling the operation of a single or multiple unmanned boats equipped with a first measurement sensor, and includes a data collection request acquisition unit that acquires request information or plan information for an ocean data collection mission that uses the first measurement sensor to collect ocean data of the ocean area in which the unmanned boat is deployed, an other mission request acquisition unit that acquires request information or plan information for other missions different from the ocean data collection mission, a multiple mission execution feasibility determination unit that determines whether multiple missions including the ocean data collection mission and the other missions can be carried out in parallel, and an unmanned boat operation execution unit that controls the operation of carrying out the multiple missions in parallel when it is determined that the multiple missions can be carried out in parallel.

The invention provides a multi-mode satellite terminal power supply adaptive management method. The method comprises the following steps: S1, acquiring an operation state and resource sensing data of a satellite terminal power supply; obtaining business demand data; s2, determining a plurality of working modes corresponding to the satellite terminal power supply based on the operation state and the resource sensing data; s3, based on the plurality of working modes, determining total power consumption corresponding to the plurality of working modes; s4, determining a satellite terminal power management optimization strategy based on the total power consumption, the operation state, the resource sensing data and the service demand data; s5, adjusting a satellite terminal power supply control instruction in a corresponding working mode by adopting a satellite terminal power supply management optimization strategy; judging whether the system operation parameters of the satellite terminal power supply meet preset change conditions or not; and if yes, repeating the steps S1-S5. Signal abrupt change is responded in real time through a closed-loop feedback mechanism, the sensing and decision-making process is cyclically updated, and self-adaptive balance of energy efficiency, communication quality and battery life in a complex scene is achieved.

Systems, apparatus, and method for momentum management for a spacecraft are disclosed. An example spacecraft includes a body; an appendage moveably coupled to the body; a thruster carried by the body; machine-readable instructions; and processor circuitry to execute the machine-readable instructions to detect a change in angular momentum at the spacecraft during firing of the thruster; and in response to the detection, cause the appendage to move relative to the body to adjust a center of mass of the spacecraft relative to a net thrust vector associated with the thruster.

A design method of a trust region satellite-borne dual-base interference SAR system is provided, relating to the field of radar measurement. Firstly, error sources of the interference SAR system are classified and analyzed, phase errors caused by image decoherence are calculated according to wave positions, and an analytical expression of baseline-related errors is reserved; a maximum phase error caused by imaging processing and an electronic device are estimated; and a baseline error is estimated. Then, the phase errors and the baseline errors are converted into elevation errors through an elevation fuzzy number, thereby obtaining an analytical expression of the elevation errors and the baseline parameters (length and dip angle). With a relative height measurement accuracy index of the system as an optimization object, a feasible solution interval between the baseline length and the baseline dip angle of the system is solved through a trust region algorithm. A minimum value of the baseline length is corrected according to a flying-around safety distance of a dual-satellite formation. According to the method of the present disclosure, the design of the satellite-borne dual-base interference SAR system can be guided, so that the satellite-borne dual-base interference SAR system is ensured to carry out terrain elevation surveying and mapping tasks with high accuracy and high reliability.

A satellite stack includes satellites arranged in a stack, where each satellite in the stack has a planar portion with an associated support structure and a subsystems housing containing satellite subsystems extending along an edge of the planar portion. For each satellite, a height of the subsystems housing is greater than a combined height of the planar portion and the associated support structure. Each satellite is oriented at a nonzero rotational angle with respect to the satellite immediately beneath, such that no satellite has its subsystems housing positioned directly above the subsystems housing of the satellite immediately beneath.

The satellite-borne heterogeneous computing system comprises a CPU (central processing unit) module, an FPGA (field programmable gate array) module and an NPU (network processing unit) module, the CPU module is used for receiving instructions and telemetering polling of a satellite service computer, reading load scientific data received by the FPGA module, providing installation and operation environments for various types of universal application programs, and scheduling and controlling the NPU module; the FPGA module is used for providing external interface expansion, receiving load scientific data and sending the load scientific data to the CPU module, and is also used for providing a logic control function comprising SPI FLASH triple modular redundancy logic and barking reset; and the NPU module is used for performing high-performance calculation and providing a high-computing-power environment for the system. The system has high flexibility, reliability and universality, and can operate various computing software and neural network algorithms; the power consumption is low, and strict requirements of a satellite platform on power consumption can be met. And a plurality of systems are interconnected to form a distributed computing system, so that the performance is extremely high.

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

The invention provides a low-orbit navigation constellation autonomous orbit determination method and system under a weak inter-satellite topological structure, and relates to the technical field of satellite orbit determination, and the method comprises the steps: obtaining various observation data of an inter-satellite link on a predetermined orbit, and carrying out the preprocessing of the various observation data; establishing a synchronous observation equation of orbit and clock correction parameters by adopting a segmented linear polynomial satellite clock correction modeling method, generating orbit, clock correction and related error parameters combining satellite-borne GNSS and inter-satellite link data, and forecasting; based on an error analysis result, multiple key features are extracted from the generated orbit, clock error and related error parameters to serve as input of a parameter prediction model, and a parameter time sequence of four types of prediction including relative weight, inter-satellite link hardware delay, low-orbit satellite clock error and unmodeled error is obtained through output; and uploading the predicted parameter time sequence and the broadcast ephemeris to a main satellite together, and transmitting to each sub-satellite by the main satellite through an inter-satellite link to realize distributed autonomous orbit determination of the sub-satellites.

The invention discloses a method for constructing and screening orbital transfer speed increment required by on-orbit service of a pulse thrust spacecraft based on a CW equation, belongs to the field of on-orbit service of spacecrafts, and solves the problem of low precision of a maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft. Comprising the following steps: constructing a relative motion equation of a pulsed thrust spacecraft relative to space debris based on a CW equation; inputting the orbital transfer state transfer matrix and the speed transfer matrix into the relative motion equation at the same time to obtain an on-orbit service orbital transfer model; setting a termination state distance constraint, a termination state illumination angle constraint and an orbital transfer speed increment constraint according to task requirements; and inputting all the termination moment position vectors conforming to the termination state distance constraint into the on-orbit service orbital transfer model, and reserving orbital transfer speed increments conforming to the termination state illumination angle constraint and the orbital transfer speed increment constraint to obtain a screened orbital transfer speed increment solution set. According to the invention, the precision of the maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft is improved.

A method for automated spacecraft design includes receiving mission parameters provided by a client for spacecraft design, generating one or more concept designs based on the received mission parameters, a concept design including one or more components selected for each subsystem included in a designed spacecraft, for the concept design, automating a generation of a frame design that includes a plurality of thermal management components to be formed through a single metal-additive manufacturing process, performing a series of analyses or simulations to evaluate a performance of the frame design, and generating a digital twin representing a virtual representation of the designed spacecraft based on the generated frame design and the series of analyses and simulations.

The invention discloses an active orbital transfer method, device and product of an ultra-low orbit satellite, and relates to the technical field of satellite detection. Comprising the steps of inputting a first instantaneous orbit element and an attitude strategy to an orbit mechanical model under the condition that the first instantaneous orbit element of a satellite at a first moment and the attitude strategy of the satellite in an adjustment period are obtained, and obtaining a second instantaneous orbit element of the satellite at a second moment; converting the second instantaneous orbital element into an average orbital element, and making a difference between a semi-major axis parameter in the average orbital element and a target average orbital semi-major axis to obtain a predicted attenuation height of the satellite in the adjustment period; if the predicted attenuation height is smaller than zero, the next adjustment period is started at the third moment, and the interval duration between the first moment and the third moment is T / 2; if the predicted attenuation height is larger than zero and smaller than or equal to the first attenuation threshold value, the satellite is controlled to rise the predicted attenuation height between the first moment and the third moment, and the next adjustment period is started at the third moment.

Embodiments generally relate to small satellites. Example small satellites may include a satellite body having a rectangular tile shape, with a substantially flat front side having a length and width and being separated from a substantially parallel back side by a depth, wherein the depth is equal to or less than a third of the length and the width. A patch antenna array may be positioned on the front side and / or a large patch antenna may be positioned centrally on the front side. A solar panel may be positioned on or coupled to the back side.

The invention provides an adjustable air-breathing power ultra-low orbit satellite configuration. The adjustable air-breathing power ultra-low orbit satellite configuration comprises an adjustable closed head part, a body, a stable wing surface and a contraction tail part, the adjustable closing head and the contraction tail are connected to the two opposite ends of the body respectively, and the stable wing face is connected to the side, close to the contraction tail, of the body. The design that the adjustable closing head is an openable / closable head is adopted, the adjustable closing head is a spherical head after being closed when a satellite does not need to inhale, and resistance is reduced; and after being expanded, the air inlet channel can be contracted into an inwards-concave air inlet channel, so that active compression of air is achieved. Therefore, task requirements under different flight heights can be met through flexible switching of the inspiration working mode and the non-inspiration working mode, the aerodynamic resistance of the satellite in an ultra-low orbit is effectively reduced, and the hang time of the satellite is prolonged.

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

The invention relates to the technical field of spaceflight, in particular to a flat plate type satellite configuration. The flat plate type satellite configuration comprises a first mounting plate, a second mounting plate opposite to the first mounting plate and a frame clamped between the first mounting plate and the second mounting plate, and the first mounting plate and the second mounting plate are arranged on the two opposite sides of the frame; the frame comprises a plurality of edge beams connected end to end and a plurality of middle beams detachably fixed to the inner sides of the edge beams, and the interior of the frame is divided into a plurality of containing spaces used for containing satellite single machines through the middle beams. According to the flat plate type satellite configuration, the frame is clamped between the first mounting plate and the second mounting plate, the bearing capacity can be improved, the manufacturing difficulty and cost are reduced through detachable connection of the middle beams and the edge beams, and protection can be provided for the satellite single machine through the containing space.

The invention discloses a super-large plane array type space computing power center, and belongs to the field of space technologies and space infrastructures. The space computing power center is deployed in a morning and evening orbit and comprises a plurality of computing power satellites which are physically docked in orbit through a docking mechanism. Each computing power satellite comprises a satellite platform, and a computing power load, a condensation type energy system and a space pump drive fluid loop system which are arranged on the satellite platform, and a docking mechanism for inter-satellite in-orbit physical docking is also assembled on the satellite platform. A plurality of computing power satellites form a planar array type rigid whole body through the docking mechanism, and energy scheduling among the light condensation type energy systems is realized based on the docking mechanism. According to the space computing power center, through multi-satellite physical docking and energy collaborative scheduling, the energy utilization efficiency and the power supply reliability of the system are improved.