11 results about "Circular orbit" patented technology

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 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.

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 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.

The present invention relates to a gonioradiometer and method for measuring at least one illumination or emission characteristic of a light source (1,S) in a direction-dependent manner. The gonioradiometer comprises at least one sensor (SR) and a device (4,5) for moving the light source (1,10,S) such that the sensor (SR) records measurements indicating the illumination or emission characteristic on a spherical surface centered on the radiative centroid (11,LS) of the light source (1,10,S). Each measurement of the illumination or emission characteristic is assigned a specific plane of a planar system (A) and a specific radiation angle (α) within this plane. The device (4,5) for moving the radiation source (1,10,S) includes a worktable (2) having a rotation axis (A4) to which the radiation source (1,10,S) can be attached. The device (4,5) is designed and configured to move the worktable (2) tangentially along a circular orbit (3) centered on the radiation centroid (11,LS) to set different radiation angles (α), and to rotate the worktable (2) around the rotation axis (A4) to set various planes of the planar system (A). The radiation centroid (11,LS) of the radiation source (1,10,S) attached to the worktable (2) remains stationary both while the worktable (2) is moving along the circular orbit (3) and while it is rotating around the rotation axis (A4).

To provide a reaping blade-driving device capable of readily ensuring the safety of operation by suppressing the generation of scattered materials during the operation and suppressing the generation of vibration and noise while suppressing the nipping of foreign materials by the reaping blade.SOLUTION: A base unit that supports the plurality of cutting blade units in a stacked state, and a driving mechanism that is supported by the base unit and performs a cutting operation by driving at least one of the cutting blade units, the cutting blade unit includes a plurality of cutting blades that protrude in different directions around the base unit, and the drive mechanism performs a cutting operation in which the cutting blade of one cutting blade unit and the cutting blade of another cutting blade unit overlap each other by causing at least one of the cutting blade units to perform a translational movement in an orbital path having no dead point.SELECTED DRAWING: Figure 1

The invention discloses a micro-nano satellite multi-pulse maneuvering method considering thrust constraint and J2 perturbation. The method aims at solving the problems that single speed increment implementation of the optimal maneuvering section of micro thrusters such as micro-nano satellite electric thrust / cold air and the like is limited, the maneuvering precision is reduced due to the fact that a long-term maneuvering orbit is influenced by J2 perturbation, and consequently the multi-satellite formation capture and reconstruction task of a near-circular orbit is difficult to complete. According to the method, the mapping relation between pulses and relative orbit parameters is established based on the relative orbit elements, a supplementary allocation method is provided for thrust constraint to achieve maneuvering pulse splitting, and a pulse sequence is corrected through iteration to compensate for J2 perturbation errors. According to the method, the multi-pulse maneuvering trajectory planning precision under the J2 perturbation influence can be effectively improved, the centimeter-level theoretical control precision can be achieved through 2-3 times of iterative computation, and the method has important practical value in the aerospace field.

The application provides a trajectory planning method for target window rendezvous, comprising the following steps: step one, calculating the latitude amplitude angle range that can be reached; step two, preliminarily calculating the time of strategy orbit transfer; step three, correcting the time of orbit transfer through the way of orbit recursion; step four, calculating the strategy of orbit transfer; and step five, long-term drift correction strategy. The application provides a trajectory planning scheme for reaching a fixed target orbit (circular orbit) at a specified phase within a given time, and proposes a scheme for correcting the strategy of orbit transfer based on orbit recursion.

The application discloses a kind of near-circular orbit space target initial orbit determination method based on single optical satellite, belong to space-based situation awareness technical field.The method includes: S1, using two groups of observation data successively, according to the right ascension and declination calculation space target true anomaly angle variation, based on circular orbit assumption through cosine theorem and search iteration estimation space target distance to geocenter within two groups of observation period;S2, repeat S1 for all observation data within period, using elliptical orbit equation and least square method calculates semi-major axis, eccentricity, and iteratively optimizes initial true anomaly;S3, according to semi-major axis, eccentricity and initial true anomaly calculation space target position and certain point velocity vector, utilize initial orbit normal vector and two-body orbit model restores initial orbit;If eccentricity is larger, then using Lambert specific orbit calculation velocity.The method can efficiently generate initial orbit using short arc angle data, avoid iteration not convergent, improve near-circular orbit space target orbit determination accuracy.

This invention discloses an optimization method for low-thrust transfer between circular orbits based on piecewise constant yaw angle guidance, belonging to the field of spacecraft orbit control technology. The implementation method of this invention is as follows: the low-thrust transfer process is divided into a transfer segment, a drift segment, and an adjustment segment; a piecewise constant yaw angle guidance strategy is proposed, using latitude argument angles α, β, γ as switching points, dividing each orbital period into four arc segments, maintaining a constant yaw angle and thrust switching state within each arc segment; based on the Gaussian perturbation equation, analytical solutions are derived for the long-term changes of the orbital semi-major axis, orbital inclination, and right ascension of the ascending node under this strategy and perturbation influence; the parameters of the optimal drift orbit are estimated based on the impulse assumption; a mixed-integer nonlinear programming model is established with the objective of minimizing the total velocity increment, and based on the dynamic relationship that "inclination changes are only generated by the thrust of arc segment 1," the model is transformed into several nonlinear programming sub-problems for solution, obtaining near-optimal guidance parameters. This invention avoids tedious numerical integration, significantly improves the efficiency of orbit design and optimization, and the resulting transfer scheme has fuel consumption close to the theoretical optimum. Moreover, the algorithm is robust and suitable for rapid orbit design tasks such as large-scale constellation deployment.