122 results about "Orbit/Orbital" patented technology

The invention discloses a method and device for cleaning space debris. According to the method, electromagnetic force is applied to the space debris running on an original orbit; the speed and/or direction of the space debris are/is changed under the action of the electromagnetic force; the space debris of which the speed is changed is changed to run on a new orbit under the action of gravity; at least the height of the perigee of the space debris on the new orbit is smaller than the height of the perigee of the space debris on the original orbit; the space debris running on the new orbit finally falls into the atmosphere of the earth under the action of atmosphere resistance. The device comprises a satellite bearing platform which is provided with an electric field device cleaning the space debris or a magnetic field device cleaning the space debris or a combination of the electric field device and the magnetic field device. The method and device for cleaning the space debris do not have high requirements for the accuracy of position detection of the space debris, micro space debris can be effectively cleaned, solar energy can be fully utilized for power generation, and self-carried energy is saved. The device for cleaning the space debris has the advantages of being simple in structure, low in manufacturing cost, and better in pulse action effect.

The invention discloses a method for compositely controlling attitudes and orbits of in-orbit dragging combination spacecrafts. The method includes steps of identifying mass property parameters of the combination spacecrafts in an in-orbit manner; determining composite control on the attitudes and the orbits of the combination spacecrafts in orbit maneuver periods. The method has the advantages that abandoned spatial targets can be deorbited and maneuvered under the conditions of high disturbance torque and uncertain parameters; the method can be used for cleaning abandoned satellite or space debris in the earth orbit, and accordingly clean running environments can be provided for satellites which run normally.

A system is dedicated to the control of the deployment of at least two spacecraft (ES1, ES2) which are provided with maneuvering means (MD1, MD2) and are intended to move according to a chosen formation. This system includes a control device (MM11, MM12, MT, MC1) comprising i) first measurement means (MM11, MM12, MT) responsible for determining substantially simultaneously and with high precision the orbital positions of the spacecraft (ES1, ES2), and ii) first calculation means (MC1) responsible for determining for each of the spacecraft, as a function of their orbital positions, maneuvers intended to position each of them at a chosen instant substantially in a chosen position with respect to a reference trajectory, having regard to the time law of a reference craft (ES1) on this reference trajectory (TR), so as to place the formation in a chosen configuration.

The invention relates to satellite autonomous target identification, and discloses a radar-based space debris autonomous target identification and early warning method, including steps 1, capturing the target and performing high-precision relative navigation measurement; 2, according to the spacecraft orbit and relative measurement data Calculate the orbit of space debris; 3. Recursively predict the orbit of space debris and calculate the minimum relative distance; 4. Release early warning information; 5. Measure and update. The invention realizes the autonomous identification of space debris by the satellite, and provides early warning to avoid the impact of the space debris, and can effectively avoid the damage caused by the space debris to the spacecraft.

The invention relates to an earth-moon libration point interorbital transfer design method. The method comprises the following steps that a, under the earth-moon mass center rendezvous coordinate system, an orbital kinetic equation of a detector is built; b, libration point invariant manifold data are stored; c, a transfer track splicing point range is searched; and d, a libration point interorbital optimal transfer track is determined. According to the libration point track invariant manifold state amount, splicing point positions of different track segments are rapidly determined, positionsof escape points and capture points are automatically adjusted, and the target track size are adjusted automatically, and the low-energy transfer track capable of meeting the task needs can be effectively designed.

The invention discloses a different-orbit single-satellite time-sharing frequency measurement positioning method based on star position optimization. The method comprises the following steps: measuring the frequency of a signal sent by a ground radiation source reaching an observation satellite; establishing a different-orbit single-satellite time-sharing frequency measurement positioning model, selecting a plurality of satellites with different orbits to respectively carry out frequency measurement at different moments, and establishing a positioning equation set according to a plurality of measured different signal frequencies; deriving geometric dilution precision factors, and analyzing the influence of various factors on the positioning precision; and establishing a star position optimization model, and optimizing the selection of orbit and satellite observation moments to improve the positioning precision. According to the invention, within a limited range of visibility, single satellites flying through different orbits of a positioning visual area at different moments are used for repeatedly and independently observing the same interference source; due to the difference in the aspects of flight direction, observation position distribution and the like, compared with a single-orbit single-satellite time-sharing frequency measurement positioning scheme, the positioning effect is stable, the precision is higher, the space between satellites is larger, and the selection of measurement positions can be more flexible.

The invention discloses a preferable short arc orbit determining method based on Gauss solution cluster, and belongs to the astrodynamics field, including: grouping the observation data, using Gauss method to obtain the target state vector at the corresponding time point for each group of data, forming a solution set of preliminary estimation; dividing the solution set of preliminary estimation into a position component vector solution set and a velocity component vector solution set for clustering to obtain a position component vector solution cluster and a velocity component vector solution cluster; based on the position component vector solution cluster and the velocity component vector solution cluster, generating a two-dimensional trajectory solution set; evaluating each of the two-dimensional trajectories by using a trajectory optimal method, calculating the number of root of orbits corresponding to the optimal two-dimensional trajectory, thereby completing determination of initial orbit.

The invention discloses an orbit planning method of a sun-synchronous circular orbit. A satellite launching orbit model, orbit regression characteristic constraints, target area illumination condition constraints and an accurate earth motion model (earth complex motion models such as age difference, nutation, rotation and polar movement are considered in the design) can be combined to design a sun-synchronous circular orbit scheme so as to realize rapid and effective detection of a target area. The problem to be solved by the satellite orbit planning design is to complete the planning design of a sun-synchronous circular orbit for different types of target area detection according to selected target area information, constraint conditions of orbit regression characteristics, constraint conditions of reconnaissance loads, launch deployment parameters, an accurate earth motion model and other data.

The invention discloses an earth-moon/moon-earth direct transfer orbit design method based on a reachable set concept. According to the method, the relative geometrical relationship between the earthlow-orbit berthing orbit and the moon is analyzed; concepts of an earth-moon direct transfer reachable set and a moon-earth direct transfer reachable set are defined, then numerical solution is performed on the earth-moon direct transfer reachable set by introducing a minimum near-moon-distance double-pulse earth-moon direct transfer orbit, and the moon-earth direct transfer reachable set is obtained based on symmetry of the reachable set. The earth-month direct transfer reachable set and the month-earth direct transfer reachable set effectively disclose geometrical conditions for realizing earth-month/month-earth direct transfer, and can be used for quickly analyzing orbit transfer windows and transfer orbit characteristics of the earth low-orbit space station to and fro the moon low orbit, so that the method can be directly applied to moon detection task analysis based on the earth low-orbit space station. According to the method, the high-precision earth-moon/moon-earth direct transfer orbit can be obtained, and the method is rapid and effective.

The invention discloses a novel intelligent orbital transfer structure for a tunnel inspection robot, and belongs to the technical field of machinery manufacturing. The invention aims to solve the problems that the existing inspection robot in a tunnel cannot carry out multi-orbit orbital transfer inspection tasks and cannot select inspection routes in the inspection process. The novel intelligentorbital transfer structure comprises a front orbital-transfer orbit translation slide, a rear orbital-transfer orbit translation slide, orbital-transfer orbit sliding connection blocks, sliding connection block rolling wheels, a left orbital-transfer orbit, a right orbital-transfer orbit, a limit travel switch and a travel switch paddle. The novel intelligent orbital transfer structure is characterized in that the left orbital-transfer orbit and the right orbital-transfer orbit are respectively provided with two orbital-transfer orbit sliding connection blocks, the connecting blocks are provided with the sliding connection block rolling wheels, and the sliding connection block rolling wheels can translate along the orbital-transfer orbits; and after the robot identifies an orbital transfer position identification code, the left orbital-transfer orbit and the right orbital-transfer orbit move to the position overlapping with an original orbit, the limit travel switch is triggered, andorbital transfer is completed, so that the robot can continue to move. The novel intelligent orbital transfer structure has the advantages of simple structure, clever design and low cost and can effectively complete the orbital transfer movement of the tunnel inspection robot, and the orbital transfer precision is high.

The invention relates to a carrier rocket elliptical orbit online planning method, in particular to a rocket online trajectory planning method, and belongs to the field of aerospace guidance control.According to the method, the nonlinear constraint, namely the elliptical orbit injection constraint, can be converted into the convex constraint so that the elliptical orbit planning problem can be constructed into the convex planning problem, and online solving and implementation are facilitated. By updating the orbit injection point conjecture and establishing a new elliptical orbit convex programming model through the sequence, the orbit injection precision can be ensured, and the terminal constraint of elliptical orbit planning is accurately met.

The invention discloses an analytic low-thrust circular orbit different-plane intersection optimization method, which comprises the following steps of: firstly, obtaining an input spacecraft initial orbit, a target orbit needing to be intersected and a specified intersection moment, and calculating an orbit element difference between the spacecraft initial orbit and the target orbit needing to be intersected at the given intersection moment under the condition that the spacecraft is free of thrust and the target orbit needing to be intersected under the condition that the spacecraft is free of thrust; a simplified model for describing the intersection trajectory of the different planes of the small-thrust circular orbits is obtained; secondly, establishing a fixed thrust starting coefficient, and optimizing to obtain an optimal thrust starting coefficient, a semi-major axis, an inclination angle and right ascension variation of an ascending node; obtaining the phase difference of the rendezvous time relative to the target orbit needing rendezvous; obtaining a new dip angle and ascending node right ascension variation; and finally, outputting a corresponding small thrust control rate and a rendezvous transfer orbit. The method solves the problems that in the prior art, a parameter optimization method is time-consuming in integration and too many in solving variables, and the optimal rendezvous trajectory is not easy to obtain.

The invention provides a satellite orbit error refinement method, which comprises the steps of respectively obtaining satellite-ground clock errors based on one-way and two-way time synchronization, comparing the clock errors, and obtaining an original orbit radial error sequence of precise orbit determination; based on the obtained original orbit radial error sequence, adopting a second-order polynomial to carry out orbit error fitting to obtain an orbit radial error correction sequence; and performing orbit error prediction based on a satellite orbit period by utilizing the orbit radial error correction sequence to obtain an orbit error prediction sequence, and converting the orbit error prediction sequence from an orbit coordinate system to an earth-fixed coordinate system to finish orbit error correction of precise orbit determination. According to the satellite orbit error refinement method, satellite orbit errors are effectively monitored, fitted, predicted and corrected under the conditions that dynamic factors are not considered and the range of a monitoring station is not expanded by means of comparing the two clock errors, and a practical Beidou satellite orbit performance improvement method is provided.

The invention discloses a communication coverage constraint-oriented earth-moon balance point task orbit rapid determination method, and belongs to the technical field of aerospace. The implementationmethod comprises the steps of establishing an earth-moon rotating coordinate system, determining an approximate analytical solution of a Halo orbit, obtaining an accurate solution by utilizing a differential correction method, and generating a plurality of Halo orbits for communication constraint amplitude calculation; conducting meshing on a back area of the moon, and calculating the geometricalrelationship between feature sample points and the Halo orbits; selecting characteristic geometric parameters according to communication constraint indexes, constructing a regression equation for thecommunication coverage constraint, and determining a quantitative relation between the communication constraint and the orbit amplitude as well as the landing point through the regression equation; and according to the task constraints, rapidly evaluating communication conditions by utilizing the regression equation, and realizing relay task orbit determination. The method can achieve the quick determination of task orbits with different communication constraints, guarantees the communication coverage of a landing point at the far side of the moon, and has the advantage of high orbit determination efficiency.

The invention discloses a regression track design method in a high-precision gravitational field. The method comprises the steps of building a regression track design coordinate system; setting a regression orbit condition according to the precision requirement of task implementation; constructing high-order Poincare mapping of orbital state change of the satellite after a regression period from an initial state; and according to the constructed high-order Poincare mapping and set regression orbit conditions, solving an optimization problem to obtain a design initial value of a regression orbit. According to the invention, high-precision and rapid track design is realized.

The invention discloses a regional hovering orbit control method based on constant continuous thrust, and the method comprises the steps: giving a theoretical track of a regional hovering orbit according to a raindrop shape hovering configuration design method; then selecting two points from the two sides of the raindrop sharp point to serve as a starting point and an ending point for constant continuous thrust control over the task spacecraft orbit, and calculating the total control duration and the theoretical value of the absolute orbit element variation of the task spacecraft after control; and finally, by constructing a constant continuous thrust orbit control equation, calculating a regional hovering orbit control strategy based on constant continuous thrust, and achieving long-termhovering of the mission spacecraft relative to the target spacecraft in a specified region. Aiming at the requirement of an on-orbit service task for a hovering technology, the invention provides a constant continuous thrust control strategy for realizing spacecraft regional hovering, and solves the problem that the existing regional hovering orbit pulse control method is not applicable under thecondition that a spacecraft is equipped with a continuous thrust engine.

The invention discloses a thrust on-orbit calibration test method, belongs to the technical field of spacecraft electric propulsion, and can solve the problem that an existing on-orbit calibration test method is relatively large in deviation, so that the control precision of a micro thruster on an orbit is relatively low. The method comprises the following steps: when the number of running turns of a satellite is an even number of turns, applying thrust to the satellite to obtain a non-thrust acting ring and a thrust acting ring of the satellite; obtaining a thrust estimated value range of thethruster according to the average orbit semi-major axis of the thrust-free action ring and the thrust action ring adjacent to the thrust-free action ring; obtaining prediction positions of all sampling points according with the thrust estimation value range according to the orbit dynamics model; according to the prediction positions and the actual measurement positions of all the sampling pointsin each thrust application time period, obtaining a thrust optimal estimation value in each thrust application time period; and obtaining thrust calibration information of the thruster according to the thrust optimal estimation value. The on-orbit calibration test method is used for on-orbit calibration of the thrust of the satellite thruster.

The invention discloses an orbit planning method for a sun-synchronous circular regression orbit. The design of the sun-synchronous circular regression orbit needs to meet two index requirements of ascending node right ascension derivative constraint and regression coefficient constraint of the sun-synchronous orbit. According to the two index requirements, the two important parameters of the orbit semi-major axis and the orbit inclination angle of the sun-synchronous circular regression orbit can be obtained through iterative calculation; the position of a launching point near the central position of a launching area is traversed and selected according to the calculated semi-major axis of the orbit, the calculated inclination angle of the orbit, the central position of the designated launching area and other parameters; and deviation distribution of subsatellite point geocentric longitude and target point geocentric longitude are calculated when the sun-synchronous circular regression orbit subsatellite point passes through the target latitude circle for multiple times corresponding to different launching point positions, the orbit design with the minimum deviation value is selected to carry out threshold judgment, and the effective orbit design is stored. According to the invention, rapid and effective detection of the target area can be realized.

The invention relates to a method for measuring the angular velocity ({right arrow over (ω)}) of a body (2) orbiting, or in motion anyway, in space depending on the detection of the trajectory of a plurality (n) of feature points (Pi) to be observed of said body (2), said trajectory being detected on the basis of data acquired by at least one remote sensor (1); the invention relate also to the related system for measuring the angular velocity ({right arrow over (ω)}) of a body (2) orbiting, or in motion anyway, in space using such method and that comprises at least one remote sensor (1), it being possible for said at least one sensor (1) to be installed on board a spacecraft (3) or to be housed in a earth station (5).

The present invention has a preferred application for measuring the angular velocity in fields such as for recovering and de-orbiting space debris.

Method for placing a spacecraft into a lunar orbit, either by standard (i.e., impulsive) or ballistic (i.e., non-impulsive) capture, from an Earth orbit that is significantly inclined relative to the lunar orbit plane, with no constraint on the local time of perigee for the starting orbit.

The invention belongs to the field of space detection, particularly relates to an impact detection track design method and system based on a carrying form, and aims to solve the problem that in the prior art, an observer track and an impactor track cannot be cooperatively designed according to impact detection task requirements of a detector in a form that an observer carries an impactor. The present invention comprises: a main body; establishing a reference frame, constructing a two-body model, an accurate dynamic model and an engineering constraint model based on the reference system; and calculating an initial orbit set error constraint condition of the impactor through the constructed model, screening out a nominal orbit of the impactor, calculating an orbit transfer moment and an orbit transfer speed of the observer based on the nominal orbit, and calculating an accurate nominal orbit of the observer based on an accurate power model through a targeting method. According to the invention, the cooperative design of the observer and the impactor track is realized, and a reference is provided for a small celestial body impact detection task.