5 results about "Elliptic orbit" patented technology

This invention discloses a tether-based approach docking control method for an on-orbit service robot. Focusing on the approach docking process, it studies the relative dynamics and control of the tether-based on-orbit service robot. A relative dynamic equation is established for the service robot to approach a customer satellite using tether tension. The influence of the customer satellite's orbital eccentricity, orbital altitude, and the robot's tether-retrieval rate on the system's relative motion is analyzed, along with the pitch motion characteristics under constant tether length. A stable tether-retrieval control law with a constant pitch angle constraint is proposed. By designing a tether length variation law, the robot can smoothly retrieve the tether to approach the customer satellite. Numerical simulation is used to verify the tether-retrieval control law. This control method can be used on elliptical orbits with non-zero eccentricity. By maintaining a constant pitch angle through tether-retrieval control, it is beneficial for changing the system's relative motion state and stabilizing the pitch motion.

This invention provides a method and apparatus for energy planning during the descent phase of an aircraft. After the first-stage engine separates, the minimum RCS attitude of the aircraft is calculated based on its current position and the position of the target detection point. When the aircraft's flight altitude exceeds a preset altitude, its attitude is adjusted to the minimum RCS attitude, and the aircraft is controlled to fly without power. After the aircraft reaches the highest point of its orbit without power, the optimal firing time and optimal firing attitude of the second-stage engine are searched based on the aircraft's tilt angle constraints and reentry point coordinate constraints. The second-stage engine is controlled to ignite at the optimal firing time, and the aircraft's attitude is adjusted to the optimal firing attitude. This achieves the control of the second-stage engine not igniting during the ascent phase of the elliptical orbit. After the aircraft flies without power over the highest point of the elliptical orbit, it flies according to the searched optimal firing time and optimal firing attitude. After the engine is exhausted, it flies according to the attitude constrained by the reentry angle of attack, ensuring that the aircraft's orbit meets the constraints. This increases the difficulty of detection, tracking, and orbit prediction for the aircraft during the second-stage flight phase.

The application provides a large-elliptical-orbit satellite electric propulsion orbit transfer strategy planning method and system, and the method comprises the following steps: S1: according to the satellite-rocket separation point parameters, calculating the initial starting time orbit six elements of the satellite in the inertial coordinate system, and setting the target end time orbit six elements; S2: using the rectangular coordinates to represent the position and velocity vector, establishing an orbit dynamics model containing the electric propulsion force under the J2 perturbation, taking the initial starting time orbit six elements as the initial input of the orbit dynamics model, judging whether the output of the orbit dynamics model meets the preset requirement based on the target end time orbit six elements, and if the preset requirement is not met, performing an orbit change strategy until the preset requirement is met.

The application provides a ground-visible space display system and relates to the technical field of satellite display. The system comprises annular display arrays, each annular display array comprises a plurality of satellite stars, the plurality of satellite stars perform relative annular flying motion around a virtual main star, a plurality of annular flying orbits are centered on the virtual main star, and the plurality of annular flying orbits are concentrically nested. The virtual main star is located on an elliptical orbit centered on the earth, the configuration parameters of the satellite stars satisfy a preset condition, and the orbit root number difference between the configuration parameters of the satellite stars and the orbit parameters of the virtual main star satisfies a correlation configuration model. By setting the configuration parameters of the satellite stars and the orbit parameters of the virtual main star, the orbit parameters of the satellite stars are obtained, the annular display arrays are formed, the brightness and color of the light emitting devices on the satellite stars are controlled, the expected display pattern is presented, the observer can see the delivered information without communication media, and global information delivery and transmission can be realized.

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.