91 results about "Orbital maneuver" patented technology

The invention relates to a method for estimating orbit control comprehensive efficiency and propellant consumption. Based on the in-orbit calibrated disturbance torque on a satellite by an orbit control thruster during working, the method for estimating the orbit control comprehensive efficiency and the propellant consumption calculates out unexpected influences on the position of an orbit when other thrusters compensate for the disturbance torque. According to the unexpected influences on the orbit, working time of the orbit control thruster is corrected and the orbit control comprehensive efficiency of the orbit control thruster of the satellite is worked out. Propellant consumption and propellant surplus of the orbital maneuver are estimated. According to the method for estimating the orbit control comprehensive efficiency and the propellant consumption, when the satellite performs orbital maneuver in the X direction of an orbital coordinate system, determination of control parameters and estimation of propellant consumption can be more accurate, the precision of orbital maneuver of a medium-to-high orbit satellite is effectively improved and the service life of the satellite in the orbit is prolonged.

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.

The invention discloses a thruster layout method for a truss type GEO (geostationary orbit) satellite adopting subdivision optimization design. Basic demands of truss type GEO satellite thrusters for subdivision modular design in thrusting service module layout are met, and layout azimuth angles of four 10N thrusters mounted at four corners of a thrusting service module back plate are designed according to demands of highest possible satellite orbital maneuver efficiency, enough capability of overcoming disturbance torque of a high-thrust orbit maneuver engine, combination of sedimentation of propellants and three-axis attitude control, angular momentum unloading and the like; under the adverse condition that the centroid of the whole satellite is remarkably higher than the upper end of a thrusting service module, a thrust vector slant and parallel layout method for 10N thrusters in pairs is adopted for layout of thrusters for east-west station keeping and south-north station keeping, so that layout of the thrusters for east-west station keeping combines functions of pitch attitude control and angular momentum unloading, layout of the thrusters for south-north station keeping combines functions of rolling, pitching, off-course attitude control and angular momentum unloading, and propellants for attitude control and angular momentum unloading are saved.

The invention discloses a low-orbit constellation system phase keeping method, system and equipment and a storage medium, wherein the method comprises the steps: determining the deviation between an actual semi-major axis and a reference semi-major axis of a satellite and the variation of the deviation along with time according to a reference flat phase angle and an actual flat phase angle of thesatellite; according to the deviation between the actual semi-major axis and the reference semi-major axis of the satellite and the variation of the deviation along with time, determining the orbit maneuvering time and the variation of the semi-major axis of the satellite; and adjusting the phase of the low-orbit constellation system in a satellite phase keeping task period by utilizing the satellite orbit maneuvering time and the semi-major axis variation. The technical scheme provided by the invention can be used for realizing high-precision phase retention of the low-orbit constellation system; under the conditions that the satellite position and speed determination precision is not high and the positioning data sampling rate is low, high-precision phase keeping (the precision requirement such as +/- 0.1 degree) can be achieved, and the requirements of high-precision phase keeping for satellite orbit determination precision and orbit determination data sampling storage are lowered.

A multiple-constraint multi-spacecraft flight interval indicating and collision avoidance method includes the following steps: (1) confirming magnitude of attitude control average force according to telemeasuring data value of propellant consumed by attitude control of launched spacecrafts on an orbit under the cruising attitude, (2) conducting high-accuracy orbit prediction according to initial ephemeris information of each spacecraft and the attitude control average force confirmed in the step one, calculating ephemeris of each spacecraft in an inertial frame at any time and relative distance between each spacecraft at any time, and confirming the minimum relative distance between the spacecrafts, (3) changing the direction of the attitude control average force, repeating the step one and the step two, calculating the minimum value of the minimum relative distance between each spacecraft, namely, the direction of the attitude control average force corresponding to the minimum value is the worst situation, (4) comparing the minimum value confirmed in the step three with the minimum safe distance, when the minimum value is greater than the minimum safe distance, the spacecrafts has no risk of collision, otherwise, within a first circle observation and controlling track segmental arc, one spacecraft which is selected conducts orbital maneuver, and thereby the distance between the spacecrafts is widened and the collision risk of the spacecrafts is avoided.

The invention discloses a space bipropellant orbit-control engine based on a two-way single-nozzle centrifugal injector. The space bipropellant orbit-control engine is suitable for meeting the task requirements of orbital maneuver, orbital transfer and the like of aerospace crafts. The space bipropellant orbit-control engine comprises a mounting connection plate, a heating device, a control valve,screws, a body structure, an inner nozzle, an outer nozzle, a temperature measuring sensor, a combustion chamber and a jet pipe. After the control valve is started, the two propellants of an oxidizing agent and a combustion agent enter a swirling cavity of the outer nozzle and a swirling cavity of the inner nozzle of the centrifugal injector correspondingly to rotate at a high speed, rotational flow fog cones are formed on the nozzles and are further atomized and sprayed out according to a certain mass flow proportion, the two propellants are mixed for combusting in the combustion chamber coated with a high-temperature resistant and oxidation resistant coating, high-temperature fuel gas is formed ultimately and is further expanded and accelerated in the Laval jet pipe and accordingly thrust is formed.

The invention discloses a space-target positioning method based on multi-model filtering. The space-target positioning method comprises the following steps: firstly, selecting the position and the velocity vector of a space target as state variables, selecting different system noise variance matrixes, and establishing a plurality of models for describing the uncertainty of the orbital maneuver acceleration of the space target; secondly, establishing a plurality of parallel-extension Kalman filtering algorithms for the plurality of models, and respectively carrying out filtering to obtain a plurality of state estimated values; thirdly, calculating corresponding filtering weights according to the conforming degree of all the state estimated values and observed quantities; and finally, calculating the weighted sum of the state estimated values of all the parallel-extension Kalman filtering algorithms according to the filtering weights to obtain navigation filtering results, namely the estimated values of the position and the speed vector of the space target. The space-target positioning method disclosed by the invention has the advantages that the system noise variance matrixes playing a leading role can be selected in a self-adaptive manner, so that the tracking and positioning capabilities of a filter to the maneuver target is enhanced, and the space-target positioning method is favorable for solving the problem that the uncertainty of the maneuver acceleration influences the positioning accuracy of the space target.

The invention discloses a spacecraft cluster only-ranging relative navigation method based on consistency filtering. According to the method, measurement sensors are not added to members of the spacecraft cluster, and related orbital maneuvers are not performed. Under the condition that the absolute orbit is unknown, only only-ranging relative navigation of the spacecraft cluster can be realized by constructing the consistency expansion Kalman filter only by relying on the lever arm effect of the communication receiver antenna which is installed to deviate from the spacecraft mass center and the geometric topological constraint between cluster members. The method comprises: performing evolution of a relative orbit by taking a relative orbit motion equation among members of a spacecraft cluster as a navigation state equation; taking relative distance information measured in a receiver TOA mode installed in a manner of deviating from the mass center of the spacecraft as a measurement quantity; introducing geometric topology information among the members of the cluster as a consistency constraint; and designing a consistency-based extended Kalman filter to finish only-ranging relativenavigation of the spacecraft cluster.

An active-passive-combined rope net dragging safety orbit leaving control method includes: tethered system mass center acceleration is designed according to tethered satellite orbit leaving task, and a propelling system applies corresponding push force on a primary satellite so as to track the designed orbital maneuver path in real time; tension control quality is designed, a tethered winch mechanism generates corresponding tension to be applied to the primary satellite so as to actively control the in-plane swing angle of a tether, and orbital maneuver of the tethered system is achieved by dragging secondary satellites through the orbital maneuver of the primary satellite and tether tension control. The method has the advantages that the large vibration of the in-plane swing angle of the tether is avoided during dragging, the in-plane swing angle of the tether is stabilized in a constraint range, and collision is prevented; meanwhile, the in-plane swing angle is restrained by the appropriate orbital maneuver path design and the active tension control applied to a combined body, and safety orbit leaving control of a target is achieved.

The invention discloses a space-based laser flight device. The flight device comprises two parts of a space-based platform and a laser load. The space-based platform mainly comprises a propulsion subsystem, a power subsystem, a communication subsystem, a control subsystem and a temperature control subsystem. The laser load mainly comprises a high-energy laser device, a transmitter-telescope and atarget acquisition tracking and aiming subsystem. The flight device can remove space debris with high efficiency, meanwhile, a large-scale high-frequency orbital maneuver can be avoided, and the service life of the flight device is prolonged.

A propellant budge-based low orbit elliptic track satellite successful injection determining method is provided. Through acquisition of relationships between propellant waste and apogee, perigee, eccentricity ratio and inclination angle adjustment amount, an injection success determination formula containing the apogee, perigee, eccentricity ratio and inclination angle adjustment amount can be achieved via total propellant waste provided by satellite orbital transferring; primary orbital transferring propellant amount allowed by a satellite can be determined according to limitations of each party; the apogee, perigee, eccentricity ratio and inclination angle deviation value of the satellite can be determined after separation of the satellite and a rocket; and whether the satellite successfully enters an orbit is determined via the determination formula. An orbital maneuver theory and formula are employed to achieve an analytic calculation formula, so accuracy, rationality, simplicity and high efficiency can be achieved; easy operation is provided and the method is in particular suitable for quick determination of success carry launch; and basis and instruction are provided for implements of emergency measures of the satellite upon problems during the carrying launch.

The invention discloses a track control method for a space rendezvous ground guide section with limited orbit maneuvering, aiming to realize high-precision track control of the space rendezvous ground guide section under the condition of limited orbit maneuver. The technical solution is to first determine the orbital maneuvering scheme of the tracking spacecraft as five orbital maneuvering schemes, determine the orbital control parameters M1~M5 that need to be calculated; calculate the orbital control parameter M1 after the tracking spacecraft enters orbit, and upload it to the tracking spacecraft. The maneuvering system performs the first orbital maneuver; after the first, second, and third orbital maneuvers are completed, the orbital control parameters M2, M3, and M4 are calculated respectively, and uploaded to the tracking spacecraft, and the orbital maneuvering system performs the second, third, and fourth maneuvers Orbital maneuvering; finally implement the fifth orbital control according to the orbital control parameter M5 (obtained when calculating M4). The invention has less number of track maneuvers, lower requirements on the attitude control system, improves the track control accuracy of the ground guidance section under the condition of limited track maneuvers, and has fast calculation speed and good real-time performance.

The invention discloses a free base space cooperation task motion reappearance experiment system, and belongs to the field of space mission ground validation. In order to overcome the shortcoming that an existing space cooperation task ground validation system cannot simulate orbital motion of a spacecraft, orbital motion of the spacecraft is simulated by a free base, and a task execution mechanism is mounted on the free base and used for finishing cooperation tasks. The complete process of the space cooperation tasks from long distance navigation, orbital maneuver, target approaching to final task execution can reappear on ground motion, and feasibility of an orbital maneuver control scheme and a task control scheme is verified.

The invention discloses a method for detecting orbital maneuver of a Beidou satellite in real time based on a phase observation value. The method comprises eight steps of data acquisition, data preprocessing, error correction, equation establishment, parameter calculation, empirical threshold calculation, orbital maneuver judgment and maneuver satellite number judgment. According to the invention,epoch difference is directly carried out on an original phase observation value; an inter-epoch differential observation model is formed; the influence of partial common errors can be eliminated; theerror correction level is improved, meanwhile, an integer ambiguity parameter is eliminated; real-time estimation is carried out on the three-dimensional speed and the clock difference epoch change parameter; the three-dimensional speed and the corresponding parameter estimation information of an observation station can be directly obtained; the accuracy of orbital maneuver detection is improved,meanwhile, real-time operation is facilitated; orbital maneuver can be efficiently detected; the influence of satellite maneuver on precise orbit determination is reduced; the available observation time of a user side satellite is prolonged; and the method has an important value for improving precise navigation positioning time service.

The invention discloses a high-integration three-unit cubic satellite capable of maneuvering orbital transfer. The cubic satellite comprises a satellite basic platform, and the satellite basic platform comprises a main force bearing structure, a solar cell array, a satellite computer, a lithium battery pack, a power supply control panel, a UV antenna, a UV communication machine, an attitude measurement and control system, other related accessories and a propulsion and control system; and the control system comprises a double-pulse solid rocket thruster, an ignition control panel of the double-pulse solid rocket thruster, a cold air micro thruster, a pulse modulation module of the cold air micro thruster, a mass moment executing mechanism and a mass center adjusting mechanism. The three-unit cubic satellite has the thrust vector control capability and is high in integration degree, rapid orbital transfer and attitude orbit control of the cubic satellite can be conducted, large-range orbital maneuver of the satellite is achieved, and the application field of the cubic satellite is greatly expanded.

The invention provides a method for detecting GNSS satellite orbital maneuver. The method comprises the steps of: acquiring global observation station GNSS observation values and broadcast ephemeris through files or networks; dividing a global GNSS observation network into a plurality of sub-networks and selecting an observation station with an atomic clock for each sub-network as a base station,the other observation stations being reference stations; building a linearized ionosphere-free combination smooth pseudo-range observation equation for each observation station in the sub-networks, and then performing satellite maneuver judgment through two threads. The method can perform reliable all-weather single-epoch maneuver detection on all satellites and can provide a feasible thought forfuture real-time satellite maneuver.

The invention discloses a medium and long distance space non-cooperative target orbit maneuver detection method, which comprises the following steps: taking relative angular momentum of a spacecraft as a center test parameter, establishing a relative kinetic model by a linearized relative motion equation, and carrying out relative measurement in a manner that an optical camera installed deviating from the mass center of a tracking satellite only measures an angle; setting a maneuver detection threshold value through a correlation coefficient of data in a sudden change simulation array and a sliding window, so that orbit maneuver detection of any in-orbit target satellite is realized, and measurement and tracking of any target satellite which may maneuver by a simple algorithm are realized through a mode of restarting a filter. According to the method, the calculation steps are simplified, the maneuvering detection accuracy of the short-range kilometer-level target satellite is improved, and improvement is carried out for the long-range hundred-kilometer-level target satellite.

The invention discloses a low-orbit spacecraft quasi-real-time orbital maneuver detection method based on measured data. Aiming at the characteristics of on-orbit operation rules and actual measurement data of a low-orbit spacecraft, the invention provides the low-orbit spacecraft quasi-real-time orbit maneuver detection method based on the measured data, and the method is characterized in that according to a preset detection threshold value, and after each arc section of the low-orbit spacecraft is tracked, whether an orbital maneuver phenomenon exists in the on-orbit flight process of the low-orbit spacecraft or not can be detected in a quasi-real-time mode, so that the problems that the iterative computation is difficult to converge during the orbital maneuver in the orbital determination process of the low-orbit spacecraft, and the spacecraft is difficult to track after orbital transfer, are effectively solved.

The present invention discloses a flying-around orbit design method for continuous visible light detection of a GEO space target. The method comprises an approaching stage of a detection spacecraft toa target spacecraft and a flying-around stage of the detection spacecraft to the target spacecraft. Firstly, a flying-around orbit is designed, and for the designed flying-around orbit, the orbital maneuvering scheme in the approaching stage is designed by taking the most fuel-saving effect as an optimization goal; the key to the design of the flying-around orbit is to determine an initial phasewindow of the detection spacecraft flying around relative to the target spacecraft, an initial phase that meets the requirements is selected, and the flying-around orbit is designed according to the CW equation; and then an approaching-stage orbit is designed, and the key to the design of the approaching-stage orbit is to determine the two moments of the Lambert orbital maneuver. According to themethod provided by the present invention, the periodic factors of the GEO space target flying around are combined, the solar lighting conditions are fully considered, and the designed flying-around orbit can keep the detection spacecraft always facing the sun to perform visible light detection on the target spacecraft and can effectively meet situational awareness and on-orbit service missions ofthe GEO space target.

The invention discloses a space target track maneuvering fusion detection method for only space-based angle measurement tracking. The space target track maneuvering fusion detection method comprises a detection quantity selection method and a hypothesis test method. The detection quantity selection method takes measurement information in state estimation as detection quantity. The hypothesis test method performs hypothesis test by use of five test methods, namely, an X<2> test method, a skewness-kurtosis test method, a Jarque-Bera method, a Kolmogorov-Smirnov method, and a Lilliefors method. The space target track maneuvering fusion detection method further comprises a weighted fusion detection method. Linear weighted fusion is performed on the detection results of the five test methods to get a linear weighted detection fusion result. When the detection fusion result is greater than or equal to a maneuvering detection judgment threshold V-horizontal line, it can be judged that orbital maneuver occurs to a space target. According to the invention, the advantages of a variety of hypothesis test methods are fully utilized, the success rate of space target track maneuvering detection is improved, the detection result of a short detection window is improved, the demand for the length of a detection window is reduced, and the redundancy and robustness of the system are improved.

The invention discloses a method for realizing fast moving of non-coplanar orbits by utilizing a suspension orbit. The method is realized through moving process planning, moving price analysis and contrast, stability analysis and a multi-pulse manner; and through related simulated analysis, the feasibility and advantages of the method for serving as a non-coplanar orbit moving manner are verified. According to the method, the suspension orbit is used as a middle transfer orbit; and through reasonable design aiming at tasks, the synchronous adjustment of the orbit surface and phase angle can be realized, so that the engine thrust demand is greatly reduced, the task time consumption is saved, benefit is brought to realize the large-range moving tasks in future space, and technical support is provided for the large-range rapid orbit moving tasks.