8 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.

The invention discloses a periodic time-lag sliding-mode control method for a nonlinear spacecraft elliptical orbit intersection system. A nonlinear spacecraft elliptic orbit rendezvous system model subjected to external interference is simplified into a nonlinear periodic system according to reversible linear transformation. On the basis, the nonlinear periodic system is written into a strict nonlinear system matched with interference. A periodic time-lag sliding mode controller with a strict nonlinear system matched with interference is designed by using a periodic time-lag sliding mode control method, and a closed-loop system formed by the strict nonlinear system matched with interference and the controller is stable within a preset true anomaly under the action of the controller. The control design method overcomes the defects that a traditional sliding mode control method of an existing spacecraft elliptic orbit rendezvous system is prone to neglecting control constraints, singularity occurs in preset time control implementation, convergence time depends on an initial value condition, the convergence speed is low, and the traditional sliding mode control method is only suitable for a low-order or single-input system. The problem of spacecraft rendezvous in the preset time is solved.

The invention discloses a method for evaluating the access capability of an artificial satellite to ground targets at different latitudes. Belongs to the field of aerospace dynamics and artificial satellite orbits. The method is based on mathematical modeling, main features of an orbit dynamics physical model are extracted, and the access capability of a satellite constellation to a ground specified latitude target is rapidly evaluated with low calculation cost, including the number of access windows and the total access duration in a given time period. According to the method, satellite motion is simplified into an elliptical orbit dominated by two-body gravitational force, and a static annulus model formed by projection of a satellite side-sway range on the earth surface is introduced; by analyzing the geometrical relationship and the time characteristic that a ground target passes through the annulus along with earth rotation, and combining a satellite orbit period and an earth rotation period, an analytical expression of the access probability, the window number and the time length is deduced. The method can effectively estimate the distribution of the access capability along with the latitude, does not need a large amount of numerical simulation, and remarkably improves the calculation efficiency.

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 invention relates to the technical field of mechanical design, and discloses a smooth guiding and positioning structure of an elliptical orbital .The smooth guiding and positioning structure comprises a box body, an elliptical orbital structure, a planet body and a driving structure, the elliptical orbital structure is installed on the front face of the box body and comprises an outer orbital and an inner orbital arranged on the inner side of the outer orbital, and the planet body is arranged between the outer orbital and the inner orbital; the driving structure is installed at the lower end of the box body and used for driving the planet body to move along the elliptical orbit structure and comprises an outer gear and an inner gear, the outer gear is arranged on the outer side of the outer orbit, and the inner gear is arranged on the inner side of the inner orbit. And the movement of the planet body can be effectively guided. The smooth guiding and positioning structure of the elliptical orbit solves the problem that the smooth guiding and positioning structure of the elliptical orbit is not designed for a planet in a traditional solar system model, dynamic display of the planet cannot be achieved, and the display effect is poor.

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.