51 results about "Circular orbit" patented technology

The invention relates to a method for designing a transfer orbit from a DRO orbit to a lunar low-orbit circular orbit. The method comprises the following steps: establishing a mapping data table and an orbital transfer speed increment three-dimensional data pair according to a preset DRO orbit size parameter and a DRO aircraft initial moment phase angle; dividing the orbital transfer speed increment three-dimensional data pair into two groups according to positive and negative conditions of an orbital transfer speed increment elevation angle, acquiring ascending node longitude and transfer flight time corresponding to the data pair of the speed increment in the minimum transfer process of each group to calculate orbital transfer moments, and selecting a minimum value in the two orbital transfer moments as a selected orbital transfer moment; correcting the phase at the selected orbital transfer moment, and calculating an orbital transfer speed increment projection according to the corrected phase; and adopting a Newton iteration method to take the projection of the selected orbital transfer moment and orbital transfer speed increment as an iteration initial value to carry out iteration solution, and obtaining the transfer orbital transfer moment and the transfer orbital transfer speed increment. By adopting the method, the transfer track design precision can be improved.

To provide a transfer device which efficiently and surely implements transfer of an electronic component.SOLUTION: A transfer device comprises: a first rotary transfer section which moves a first holding section sucking and holding an electronic component along a circular orbit; a second rotary transfer section which moves a second holding section sucking and holding the electronic component along a circular orbit; and a position correction section which is provided between the first rotary transfer section and the second rotary transfer section on a transfer path of the electronic component and delivers the electronic component received from the first holding section to the second holding section after correcting a position. The position correction section includes: a sucking and holding section which includes a sucking face sucking and holding the electronic component and delivers the electronic component with the first holding section and the second holding section; a rotary transfer table which moves the sucking and holding section along a circular orbit; and a parallel driving section which moves the rotary transfer table in parallel and corrects a plane position of the electronic component sucked and held by the sucking and holding section. The sucking face of the sucking and holding section and a sucking face of the second holding section, which are opposed at a second delivery position, are parallel.SELECTED DRAWING: Figure 1

A satellite constellation (100) includes a plurality of artificial satellites (111 to 116) that number a multiple of 6. Each of the plurality of artificial satellites orbits in an inclined circular orbit a plurality of times a day. Normals to a plurality of orbital planes corresponding to the plurality of artificial satellites are shifted by an equal angle from each other in an azimuth direction. The plurality of orbital planes make up one or more orbital plane sets each consisting of six orbital planes. Timings at which the six artificial satellites orbit on the six orbital planes of each orbital plane set are synchronized with each other.

The present invention discloses an industrial CT three-dimensional image reconstruction method based on data rearrangement and conjugate rays. The method of the present invention first obtains the projection data of the sample to be measured through circular orbit scanning of industrial cone-beam CT; then, the projection data is rearranged by converting the planar detector into a cylindrical detector, and four-neighbor bilinear interpolation is used to improve the accuracy of the projection data; then, the rearranged projection data is cosine weighted, and a three-dimensional empirical weighting function is constructed by a method based on conjugate rays for secondary weighting; then, the weighted projection data is filtered by a one-dimensional ramp filter, and the reconstruction result is obtained through back-projection. The present invention has strong robustness, can eliminate the reconstruction artifacts of projection data with a large cone angle, and thus improve the accuracy of industrial CT measurement.

The invention discloses a synchronous orbit electric propulsion satellite short-range game control method based on differential game, which comprises the following steps: under an LVLH coordinate system, establishing a near-circular orbit relative motion dynamic model of spacecrafts of both pursuit and escape parties, and further determining a state space form in a spacecraft pursuit and escape game process; based on a state space form in a spacecraft pursuit game process, introducing a Hamiltonian function, and constructing a short-distance pursuit game model with optimal time; introducing a co-state equation, and describing a short-range pursuit game process as a two-point boundary value problem; and converting the problem into a parameter optimization problem, and carrying out optimization solution on the problem by utilizing an improved difference method so as to realize short-range game control.

The present invention proposes a satellite Ka-band variable coding modulation data transmission system and method, which can be applied to the Ka-band satellite-to-ground data transmission system of non-geostationary circular orbit remote sensing satellites such as low orbit or medium orbit. It includes: a navigation receiver, a data transmission integrated controller, a data processor, a Ka variable coding modulator, a filter, a line amplifier input switch matrix, a traveling wave tube amplifier, a line amplifier output switch matrix, and a ground data transmission antenna; the navigation receiver calculates the position / speed / time (PVT) information of the satellite and transmits it to the data transmission integrated controller; the data transmission integrated controller completes the modulation parameter selection of the modulation coding mode (MODCOD) and the line amplifier input power back-off (IBO) according to the information and the stored parameters, and sends the MODCOD / IBO setting instruction to the Ka variable coding modulator when the modulation parameters change; after the data processor completes the payload data processing, it sends the baseband data to the Ka variable coding modulator.

The present application relates to a kind of gravity gradient satellite segmented optimal attitude tracking control method, the present application is aimed at satellite platform, coiled stretch arm and the gravity gradient system of end payload three compositions, adopt dumbbell model to describe its swing motion in coplanar orbit transfer, establish the attitude dynamics model of gravity gradient satellite;Orbit dynamics model is established based on electric propulsion coplanar orbit transfer.Segmented optimal trajectory of gravity gradient satellite under the action of electric propulsion is obtained by designing optimal controller, while realizing the orbit transfer between two coplanar circular orbits, the stability and robustness of system are enhanced.On this basis, based on finite time tracking error function, a kind of segmented adaptive terminal sliding mode control method is designed, the problem of finite time convergence and uncertainty boundary is solved.The present application is specific object in the gravity gradient satellite of coplanar orbit transfer, the control method proposed is conducive to the improvement of control performance, easy to realize.

Ensure stable operation of the equipment. [Solution] A part processing device according to one aspect of the present disclosure comprises a plurality of work holding sections, each of which is configured to hold a work; a support section that supports the plurality of work holding sections so that they are positioned on a circular orbit; a rotation drive section that rotates the support section around the central axis of the circular orbit; an advance / retract drive section that moves a first work holding section, which is one of the plurality of work holding sections, along a predetermined direction when the first work holding section is positioned at a predetermined position on the circular orbit; a processing unit that performs a predetermined process on the work together with the first work holding section driven by the advance / retract drive section at the predetermined position; a transport fixing mechanism configured to fix each of the plurality of work holding sections to the support section in a releasable state; and an advance / retract fixation mechanism configured to fix the first work holding section, when released from fixation by the transport fixing mechanism, to the advance / retract drive section.

The invention provides an on-satellite autonomous controllable transverse side-by-side multi-satellite formation method. The method comprises the following steps: establishing a multi-satellite relative motion model under a circular orbit; taking any group of satellites as master and slave satellites, and obtaining a direct corresponding relation between a formation system baseline requirement and an orbit parameter; preliminary design of formation configuration is carried out according to inter-satellite baseline requirements of a formation system; establishing an expression of a baseline angle between the master and slave satellites and satellite orbit parameters by using a mathematical expression of a relationship between baseline vectors and orbit six elements, and taking the expression as a constraint condition of non-common-view shielding of formation configuration design; and designing orbit parameters of a plurality of satellites of the formation satellite system according to the latitude distribution requirement of the base line required by the task, and forming the formation system meeting the task requirement.

This invention proposes a method and system for remote sensing satellite orbit design based on multi-parameter fuzzy constraints. The method includes: receiving input orbit boundary constraints and design constraints; calculating a series of information about the corresponding orbit based on the altitude range, pixel resolution, revisit constraints, and number of mosaic operations for the sun-synchronous return circular orbit; calculating relevant information that can be directly filtered by subsequent design constraints based on the candidate orbit parameters; filtering the directly filtered relevant information using imaging angle constraints, swath width constraints, and integration time constraints; and filtering the orbit using latitude constraints, marking results that meet and do not meet the constraints. This invention integrates and coordinates multi-domain knowledge related to orbit design and the fragmented steps of orbit design, significantly improving the overall efficiency of satellite design.

The present invention relates to a synchrotron continuous variable energy extraction method, comprising the following steps: using a synchrotron high-frequency system to adjust particle energy so that the particle energy passes through a predetermined variable energy range; during this process, by controlling the synchrotron high-frequency system and the strength of the dipole iron, quadrupole iron, and hexapole iron magnetic fields, the particles are allowed to stably run in a circular orbit of the synchrotron; during the entire process of the particle beam energy passing through the variable energy range, a radio frequency excitation device is used to continuously excite the particles to overflow from a transverse phase space stable region and enter an electrostatic deflector, and the particles are separated from the circular orbit of the synchrotron under the action of the electrostatic deflector, wherein the energy time interval and energy value interval of the particles overflowing from the transverse phase space stable region are sufficiently narrow so as to be considered to meet the requirements of continuous energy extraction.

The invention provides a coplanar spacecraft relative motion configuration recognition method based on multiple orbit elements, belongs to the field of spacecraft relative motion configuration recognition, and solves the problem of large calculation amount during spacecraft relative motion configuration recognition. Comprising the following steps: constructing a relative motion configuration set when a target spacecraft is in a circular orbit and a tracking spacecraft and the target spacecraft are coplanar; obtaining constraint conditions corresponding to a linear motion polynomial and an elliptic motion polynomial in the relative motion equation according to the motion features of each relative motion configuration, and forming a multi-orbit element set by orbit elements in the constraint conditions; extracting input data required by a relative motion equation in the task data, outputting a numerical value of the multi-orbit element set through the relative motion equation, inputting the numerical value into a constraint condition to obtain an output result, and identifying a relative motion configuration corresponding to the constraint condition in the relative motion configuration set through the output result; according to the invention, the calculation amount is reduced and the identification efficiency is improved.

The invention provides a lunar back and bipolar landing area observation orbit design and task planning method, relates to the field of lunar exploration, and solves the problems that an existing lunar exploration satellite orbit is insufficient in adaptability, observation and data transmission planning is split, observation coverage of lunar back and bipolar areas is incomplete, data return is low in efficiency, and the satellite orbit adaptability is poor. And the requirements of long-term lunar exploration continuous observation and real-time data support cannot be met. According to the method, 14 key observation targets on the back of the moon and the two poles are screened, a 500 km height polar circle orbit is designed, an observation and data transmission joint model containing multiple constraints is constructed, a genetic algorithm-simulated annealing hybrid algorithm is adopted to optimize a task sequence and imaging parameters, and efficient observation and data return are achieved. According to the method, full coverage and high revisit of the lunar back and the two poles are achieved through the 500 km polar circle orbit, the communication pain point is solved through the combined planning model, the task efficiency and quality are optimized through the genetic algorithm-simulated annealing mixed algorithm, observation and data return can be efficiently completed through simulation verification, and stable support is provided for long-term lunar exploration.

The present invention discloses a hybrid method of neural network and parameter optimization for low-thrust rendezvous trajectory optimization. First, a fast optimization method for low-thrust rendezvous control law of near-circular orbit with a flight time of 1 circle is obtained; then two neural network proxy models are established, corresponding to the given [Δa, Δi, ΔΩ, Δe x ,Δe y ] Output whether there is a feasible solution and the ratio of the arc length of the output thrust start to the entire cycle, where Δa is the change in semi-major axis, Δi is the change in inclination, ΔΩ is the change in right ascension of the ascending node, and Δe x is the change in eccentricity on the x-axis of the inertial system, Δe y The present invention solves the problem of excessive variables and low efficiency in existing technologies.

A satellite system may have a constellation of communications satellites that provides services to users with electronic devices such as portable electronic devices and home and office equipment. A network operations center may use gateways to communicate with the satellite constellation. The satellite constellation may include sets of satellites with different orbits such as circular orbits with different inclinations, sets of satellites with elliptic orbits, sets of satellites with circular orbits of different altitudes including low earth orbits, medium earth orbits, and / or geosynchronous orbits, and / or sets of satellites with other orbits. The satellite orbits of the satellites in the satellite constellation may be selected to provide coverage to desired user population concentrations at different locations on the earth without using an excessive number of satellites.

The present invention may provide a motor including a shaft, a rotor coupled to the shaft, a stator disposed outside the rotor, a busbar disposed above the stator, and a power terminal unit coupled to the busbar connected to the coils, wherein the power terminal unit includes a first groove, the busbar includes a busbar body and a plurality of terminals disposed on the busbar body, the busbar body includes a plurality of first protrusions protruding from an upper surface of the busbar body and disposed in the first groove, and the plurality of first protrusions are disposed on the same circular orbit based on a center of the busbar.

An optimal rescue orbital elements online decision-making method based on RBFNN for launch vehicles under thrust drop fault includes establishing the flight dynamic equations of launch vehicles in the second-stage ascending phase in the geocentric inertial coordinate system, to construct a series of optimization problems of maximum semi-major axis of circular orbit under the thrust drop fault. The method further includes using the adaptive pseudo-spectrum method to solve the optimization problems of maximum semi-major, and using the maximum and minimum method to normalize the sample data to [−1, 1], using the orthogonal least square method to select the data center of the radial basis function neural network (RBFNN), where the Gaussian function is selected as the radial basis function, and the RBFNN is trained offline to establish a nonlinear mapping relationship from the fault states to the optimal rescue orbital elements.

The present application relates to a method for dynamic incremental greedy assignment of multi-satellite on-orbit services based on a resource balancing strategy. The method comprises: recursively assigning the service satellite and target satellite to the epoch time of the mission start; constructing a circular orbit rendezvous cost model to estimate the rendezvous cost matrix between the service satellite and the target satellite; constructing the objective function of the mission planning problem model based on the overall balanced allocation strategy and the rendezvous cost matrix, and designing the constraints of the mission planning problem model based on the mission start and end times and the information of each service satellite; constructing the mission planning problem model using the objective function and the constraints; designing a dynamic incremental greedy algorithm to solve the mission planning problem model to obtain the assignment result for each service satellite; and designing the access sequence and rendezvous time for each service satellite's assignment result based on the dynamic programming algorithm to obtain the final mission planning plan for each service satellite. This method can improve the efficiency of mission planning while ensuring the accuracy of the solution.

The present invention provides a coplanar spacecraft relative motion configuration recognition method based on multiple orbital elements, which belongs to the field of spacecraft relative motion configuration recognition and solves the problem of large computational complexity in spacecraft relative motion configuration recognition. The method comprises: constructing a relative motion configuration set when a target spacecraft is in a circular orbit and a tracking spacecraft is coplanar with the target spacecraft; obtaining constraint conditions corresponding to linear motion polynomials and elliptical motion polynomials in the relative motion equation from the motion characteristics of each relative motion configuration, and forming a multi-orbital element set from the orbital elements in the constraint conditions; extracting input data required for the relative motion equation in mission data, outputting numerical values ​​of the multi-orbital element set through the relative motion equation, inputting the numerical values ​​into the constraint conditions to obtain output results, and identifying the relative motion configuration corresponding to the constraint conditions in the relative motion configuration set through the output results. The present invention reduces the computational complexity and improves recognition efficiency.

A satellite constellation includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

The present invention relates to the field of aerospace measurement and control technologies, and provides a satellite deorbiting control method, device, equipment and medium based on a small thrust. It can determine different ignition strategies according to the control efficiency index values of each candidate ignition strategy, and first perform a perigee altitude lift simulation based on the ignition strategy, then perform a circular orbit control simulation, and finally execute an overall orbit altitude lift simulation to continuously lift the satellite to the graveyard orbit. Further, it evaluates the collision risk of the deorbiting process of the target satellite according to the simulation data, and when the collision risk assessment result shows that there is no collision risk between the target satellite and other satellites during the deorbiting process, it generates a deorbiting control strategy for the target satellite based on the different ignition strategies used to assist in performing a safe and efficient deorbiting control task for the satellite.

The invention discloses a lunar orbit rendezvous remote guidance rapid iterative guidance method. The method comprises the following steps: acquiring an initial orbit state corresponding to an initial moment and an orbit target state corresponding to a terminal moment; establishing a nonlinear equation set of a design variable and a terminal track state; based on the near-circular orbit element representation forms of the terminal orbit state and the orbit target state, the nonlinear equation set of the design variables and the terminal orbit state is converted; determining an iteration initial value and a convergence condition; and on the basis of the determined iteration initial value and the convergence condition, iterative solution is conducted on the converted design variable and the nonlinear equation set of the terminal track state till the convergence condition is met, the final design variable is obtained, and guidance is completed. The method is high in iteration efficiency, high in calculation speed and suitable for any orbital transfer frequency, the component of each orbital transfer can be specified at will, and the method is particularly suitable for meeting the requirements for on-orbit real-time generation of the remote guidance orbital transfer strategy of the moon orbit intersection and remote guidance strategy reconstruction under the fault condition.

To provide a technique for transmitting satellite information acquired by a monitoring system to ground equipment approximately in real time, and for controlling each satellite approximately in real time.SOLUTION: In a monitoring system, satellites constituting a first satellite constellation 831 conduct information exchange with ground equipment via a second satellite constellation 832. The first satellite constellation 831 orbits an inclined circular orbit multiple times per day. In a plurality of orbital planes formed by the plurality of satellites included in the first satellite constellation 831, their mutual normal lines deviate respectively at equal angles in an azimuth direction. In the ground equipment, flight positions of the respective orbital planes are synchronously controlled, and information is transmitted via the first satellite constellation 831 and the second satellite constellation 832, and the satellites constituting the first satellite constellation 831 and the second satellite constellation 832 respectively are controlled according to commands.SELECTED DRAWING: Figure 31

The present invention discloses a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, comprising: obtaining an initial orbital state corresponding to an initial moment and an orbital target state corresponding to a terminal moment; establishing a nonlinear equation group of design variables and terminal orbital states; transforming the nonlinear equation group of design variables and terminal orbital states based on the near-circular orbit element representation of the terminal orbital state and the orbital target state; determining iterative initial values ​​and convergence conditions; iteratively solving the nonlinear equation group of the transformed design variables and terminal orbital states based on the determined iterative initial values ​​and convergence conditions until the convergence conditions are met, obtaining the final design variables, and completing the guidance. The method of the present invention has high iterative efficiency and fast calculation speed, is applicable to any number of orbit changes, and the components of each orbit change can be arbitrarily specified. It is particularly applicable to the needs of on-orbit real-time generation of lunar orbit rendezvous and remote guidance orbit change strategies and reconstruction of remote guidance strategies under fault conditions.

Achieve both high-precision load management and high productivity. [Solution] A part processing device according to one aspect of the present disclosure comprises a plurality of work holding sections each configured to hold a workpiece, a support section that supports the plurality of work holding sections so that they are positioned on a circular orbit, a rotation drive section that rotates the support sections around the central axis of the circular orbit, a plurality of load sensors that are each provided on a corresponding work holding section and detect the load applied to the corresponding work holding section, an advance / retract drive section that moves a first work holding section, which is one of the plurality of work holding sections, along a predetermined direction when the first work holding section is positioned at a predetermined position on the circular orbit, and a processing unit that supplies a workpiece to the first work holding section driven by the advance / retract drive section at the predetermined position.

The present invention proposes a method and system for training an artificial intelligence model of a spacecraft. Based on the existing self-developed spacecraft simulation platform, the SpaceSim platform, a general-purpose platform, SpaceSimGYM, for the research and development and application of intelligent technologies for spacecraft attack and defense confrontation and gaming is developed; on a near-Earth circular orbit with a fixed confrontation time, the observation settings of the red visible light reconnaissance satellite are carried out; at the same time, the actions of the red visible light reconnaissance satellite are set; a reward return function is designed to encourage the intelligent agent to complete the orbital target interception, and the real-time state is updated and the termination conditions are set: The present invention supports simulating the confrontation and cooperation processes between multiple intelligent agents, and can support multiple types of tasks, realizing the optimal interception strategy for non-cooperative space targets based on reinforcement learning under continuous thrust. While supporting the construction of a spacecraft confrontation scenario using a visualization platform, the convenience of the scenario is achieved.

This invention discloses a dynamic obstacle avoidance and intelligent game-playing method for spacecraft swarms, comprising: S1: establishing a coordinate system with a virtual spacecraft operating in a circular orbit as the reference spacecraft; S2: performing relative motion dynamics modeling on the spacecraft swarm and the target in the coordinate system to obtain a relative motion dynamics model describing the virtual spacecraft and the spacecraft swarm; S3: acquiring the coordinate data of each sub-spacecraft in the spacecraft swarm in the coordinate system; S4: determining the distance between each sub-spacecraft and the obstacle based on the motion dynamics model and the coordinate data; S5: determining whether there is a distance greater than a preset threshold; if so, proceed to S6; otherwise, proceed to S7; S6: driving the spacecraft swarm to achieve an encirclement state based on the gravitational relationship between each sub-spacecraft and the target and the repulsive relationship between each sub-spacecraft and the obstacle determined by the artificial potential field method; S7: using an intelligent encirclement strategy to encircle the target obstacle.

A satellite constellation (110) includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

The utility model provides an annular track centripetal force analysis experiment device comprising a track assembly which comprises a release track, an annular track and a recovery track; a small ball runs on the track assembly; the releasing mechanism comprises a sliding base, an electromagnetic assembly and a guide rod unit arranged on a releasing track, the electromagnetic assembly is arranged on the sliding base, and in the initial state, the electromagnetic assembly generates magnetic force to adsorb and fix the small ball; a driving screw rod is arranged on the guide rod unit, a driving plate arranged below the sliding base is nested on the driving screw rod, and a supporting bolt arranged on the driving screw rod is screwed in to adjust the height of the sliding base. According to the utility model, the supporting bolt is screwed on the driving screw, so that the tedious disassembly and assembly process in the height adjustment process is simplified, the preparation and disassembly and assembly work before and after an experiment is reduced, the experiment time is shortened, and the experiment progress is accelerated.

The present application relates to the field of satellite communication and information technology, in particular to a shared space-based system and method. The present application provides a near-equatorial orbit low-orbit open shared space-based system, which comprises a low-orbit communication satellite constellation, a space digital base and a resource fusion scheduling center arranged on a near-equatorial orbit. The low-orbit communication satellite constellation comprises a main constellation layer and a network supplement layer. The near-equatorial low-orbit circular orbit is arranged in the main constellation layer, and the near-equatorial medium-orbit circular orbit is arranged in the network supplement layer. The present application eliminates global communication and blind area, improves resource utilization, and has instant response and customization. With the advantages of near-equatorial orbit, space perception load is carried, and the system has the prospects of commercial value closed loop, ecological partner cooperation and sustainable development: the load carrying and resource sharing mechanism is opened, the redundant satellite launch is reduced, and the problem of near-earth orbit congestion is alleviated.