107 results about "Orbit plane" patented technology

A propulsion system for the orbit control of a satellite in Earth orbit driven at a rate of displacement along an axis V tangential to the orbit comprises two propulsion modules, fixed to the satellite, and facing one another relative to the plane of the orbit, each of the propulsion modules comprising, in succession: a motorized rotation link about an axis parallel to the axis V; an offset arm; and a plate supporting two thrusters, suitable for delivering a thrust on an axis, arranged on the plate on either side of a plane P at right angles to the axis V passing through a centre of mass of the satellite; each of the two thrusters being oriented in such a way that the thrust axes of the two thrusters are parallel to one another and at right angles to the axis V.

A method for sensing and monitoring ICBM threats using a constellation of earth orbiting satellites. The method includes arranging a first and a second plurality of threat sensing satellites in earth orbit. The first plurality is selected to have a first prograde orbit plane with a first inclination angle between about 30° to 70° with respect to the earth's equator and an Argument of Perigee equal to about 270°. The second plurality is selected to have a first retrograde orbit plane with a second inclination angle between about 110° to 150° with respect to the earth's equator and an Argument of Perigee equal to about 270°. The second inclination angle is selected to be a symmetric conjugate of the first inclination angle. The first and second plurality of threat sensing satellites are distributed respectively within the first prograde and first retrograde orbits with a Walker distribution of T/P/1. Synchronization of a plurality of epochs of the first plurality of satellites with a plurality of epochs of the second plurality of satellites is performed to provide uniform inter-plane satellite spacing for improved monitoring of an ICBM threat.

The invention discloses a fault diagnosis method for output current of on-orbit satellite solar cell array. A simplified semi-physical model of the output current of the solar cell array is provided by being combined with actual on-orbit satellite telemetering historical data, on-orbit attenuation conditions of the on-orbit satellite solar cell array, the solar incident light intensity, an included angle between a sunlight vector and an orbit plane and the like, the physical model of the output current of the solar cell array is optimized by using epsilon.(cosbeta)<n>, and a dynamic alarm threshold of the output current of the solar cell array is acquired on the basis, thereby carrying out fault diagnosis on the solar cell array more accurately and more reasonably, and greatly reducing the risk of missing alarm while improving the fault diagnosis accuracy. In addition, the fault diagnosis method is moderate in calculation amount of the model, thereby being convenient for project implementation.

The invention discloses a structure variable mixed-orbit satellite constellation, comprising a GEO (Geosynchronous Orbit) sub constellation, an IGSO (Inclined Geo Stationary Earth Orbit) sub constellation and an MEO (Medium Earth Orbit) sub constellation. The GEO sub constellation comprises n GEO satellites distributed on n orbit positions of a geosynchronous orbit, and coverage areas between the adjacent satellites are connected and cover medium and low latitude areas in a full longitude manner. The IGSO sub constellation comprises five IGSO satellites, wherein the IGSO1, IGSO2 and IGSO3 satellites are located in the same orbit plane, and the phase difference between the adjacent satellites is 120 degrees; and simultaneously, the IGSO3, IGSO4 and IGSO5 satellites have the same ground track. The MEO sub constellation comprises m*k satellites distributed on k orbit planes, m satellites are on each orbit plane, and the MEO sub constellation has global coverage capability. When the structure of the constellation is reconstructed, the IGSO4 and IGSO5 satellites flexibly enter an HEO (High-altitude Earth Orbit) and run. The constellation has excellent global coverage capability, self management capability and flexible and variable structure.

The invention discloses a method for determining intersection coordinates of a satellite wave beam and the earth. The method comprises the following steps of: (1) seeking a unit vector rfs of the satellite wave beam in a satellite body coordinate system; (2) calculating a conversion matrix of the satellite body coordinate system and a satellite track coordinate system by a yaw angle, a pitch angle and a roll angle of a satellite; (3) calculating the conversion matrix of a geographic coordinate system and an earth coordinate system by longitude lambda and latitude I of the satellite; (4) calculating the conversion matrix of an auxiliary coordinate system and the geographic coordinate system by an included angle beta of a meridian plane on which the satellite is and an instantaneous orbit plane of the satellite and an included angle alpha between a movement direction of the satellite and a geographic horizontal plane; (5) converting the unit vector rfs into the earth coordinate system to obtain rfe by coordinate system conversion; (6) associating a linear parameter equation of satellite beam incidence with an earth ellipsoid equation, and calculating two groups of coordinates [xje, yje and zje]T, wherein a group of coordinates near the satellite are coordinates of an intersection point of the satellite wave beam and the earth.

The invention relates to a full-simulating sun-moon-earth instrument which comprises a sun rotation and lighting device, an earth rotation device, an earth revolution device, a moon rotation device, a moon revolution device, an earth-moon rising and dropping device, and speed regulating and control circuits, wherein the earth-moon rising and dropping device is arranged at the external end of a long arm of the earth revolution device; the earth rotation device, the moon revolution device, the moon rotation device and the moon rising and dropping device are arranged on the earth-moon rising and dropping device; all the devices above run around the sun at the centre; and the full-simulating sun-moon-earth instrument is characterized in that the earth synchronously rises and drops during performing the revolution around the sun, and the moon synchronously rises and drops during performing the revolution around the earth; with the adoption of the composite motion, the due intersect angles can be formed between an earth revolution orbit plane and the sun equator plane and between a moon revolution orbit plane and the earth equator plane; the speed ratio of the sun to the moon to the earth under the rotation and the revolution can be similar to the actual speed ratio by carrying out electronic speed regulating; and the star image drawn on a transparent celestial sphere is similar to the actual star image, therefore, the running of the three balls can be comprehensively and really displayed in a three-dimensional and vivid way, and the running rules and resulting astronomical phenomena and the natural phenomena can be popularly and quickly indicated; and the practical significance is brought to popularization of the astronomical knowledge.

The invention discloses a same orbit plane satellite observation orbit design method under a nature accompanying condition; the method comprises the following steps: designing an observation satellite initial orbit satisfying the nature accompanying condition with a target satellite; setting a visual direction, relative to the observation satellite, of the target satellite as the observation direction, and keeping space inertia orientation of the observation satellite in a flying process; selecting a visual angle of an observation camera; determining an observation direction and target satellite connecting line angle long-time cycle radiate model; adjusting the observation satellite gesture according to the long-time cycle radiate model. The satellite observation orbit design method can reduce the observation satellite space flying orbit gesture orbit coupling maneuvering requirement; a control system and an execution mechanism are simple, and cost is low.

A system which comprises a sun-synchronous orbit satellite (3) revolving around the sun (5) and apparently revolving around the earth at a period of approximately one year while keeping the direction and the geometry to the sun and the earth (4) constant at an ultrahigh altitude of several millions km keeping clear of the influence zone of the earth, and which provides any of astronomical or global observation service and communication service. An observation satellite, which meets such conditions that a vector normal to a virtual orbital plane centering around the earth is substantially fixed regardless of revolution, can sweep the substantially entire celestial sphere during a single revolution while avoiding the effect of light and heat from the earth, and can keep a local solar time at a sub-satellite point on the surface of the earth constant, can be attained. A communication satellite, where the substantially hemisphere of the earth can be covered at a time simultaneously with one satellite and the substantially entire region of the earth can be covered with a small number or three satellites, is attained. Geometrical relations with the sun and the earth can be frozen approximately by the adjustment of the i/e ratio or by the use of perturbation by the attraction of the earth gravity. Operation from launch to insertion into an intended orbit can be achieved with significantly low fuel consumption by utilizing the perturbation owing to the attraction of the earth gravity.

The invention provides a method for driving a solar wing to face the sun through a low-inclination-angle orbit uniaxial SADA. The method comprises the following steps: S1, determining coordinates of asolar vector in an orbit system by an on-board sensor; S2, calculating a satellite yaw angle according to the projection of the sun vector on the XoOYo plane of the orbit; S3, calculating expected yaw angle and angular velocity information in satellite attitude motion; S4, calculating the rotation angle of a solar wing driving device according to an included angle between the normal coordinate ofthe solar panel and the solar vector when the solar wing driving device is in a zero position; and S5, controlling the satellite to maneuver by the on-board actuating mechanism, and rotating the solar wing driving device in a matching mode to ensure that the solar wing points to the sun. The invention further provides a system for driving the solar wing to face the sun through the low-inclination-angle orbit single-shaft SADA. The method has the advantages that the single-axis SADA is used for replacing the double-axis SADA in the low-inclination-angle orbit to achieve sun orientation of thesolar wing, cost is reduced, reliability is improved, and control is easy.

The invention provides a low-orbit internet constellation satellite-ground link planning method based on constraint satisfaction, and relates to the technical field of satellites, and the method comprises the steps: constructing a satellite-ground geometric visible model; constructing an inter-satellite distance visible model; constructing a satellite-ground link visible model according to the satellite-ground geometric visible model; acquiring measurement data between the target satellite and ground gateway station equipment according to the satellite-ground link visible model; determining constraint conditions; determining the validity of the measurement data according to the constraint condition, and establishing an effective measurement link matrix according to the validity of the measurement data; determining a satellite-ground networking planning principle; establishing a target satellite-ground link planning model according to a satellite-ground networking planning principle, constraint conditions and the effective measurement link matrix; and calculating the target satellite-ground link planning model to obtain a satellite-ground link resource scheduling result. According to the method, the planning constraint problems of neglecting constellation configuration, constellation orbit plane balanced use and the like during multi-task scheduling of a low-orbit constellation without an inter-satellite link are solved.

The invention discloses a method for distributing and transferring multiple satellites launched by one rocket to different orbit planes. Firstly, orbit injection dip angles i, target heights Hf, time T for completing deployment of tasks and orbit plane drifting distances delta omega needed by all the satellites need to be determined, and critical height differences delta a for transferring the satellites to the target obit planes in the assigned time T are calculated according to the input conditions. Satellite drifting starting heights are provided through a piggybacking separation mode or an upper-level orbital transferring mode, or when the quantity of fuel carried by the satellites enables corresponding orbit transferring capability to be achieved, drifting is begun by directly transferring to the drifting starting heights from the orbit injection height. The different drifting starting heights corresponding to all the satellite are set under the premise that the satellites can arrive in the assigned time T, and the satellites are restored to designed heights respectively after arriving at the target orbit planes. Selection of obits and piggybacking is carried out on piggybacking satellites or multiple constellation satellites according to given drifting time in tasks and existing orbit injection conditions, different fuel demands of all the satellites are deployed when piggybacking conditions exist or constellation tasks meet requirements, and different orbit distributing and transferring are completed after obit injection according to designed drifting modes.

A method for determining orbit of a spacecraft by using relative position increment belongs to the technical field of determination of spacecraft orbit, and aims at solving the problem that a conventional orbit determination method based on measured epoch difference needs to employ other orbit determination method for obtaining initial position information. The method for determining orbit of the spacecraft by using relative position increment comprises: firstly, determining the motional orbit plane of the spacecraft in a three-dimensional space; then, converting the three-dimensional relative position vector into a spacecraft relative position coordinate in a two-dimensional plane, fitting the spacecraft motion orbit shape in the two-dimensional plane by employing a least square method, and determining eigenvalue of the fit matrix by employing lagrangian multiplier; and finally, determining the spacecraft orbit type according to the eigenvalue of the fit matrix, and obtaining the absolute position information of the spacecraft in an inertia reference system according to the position of a central celestial body, so as to determine the initial orbit of the spacecraft. The method is applicable to orbit determination of spacecraft.

Under the existing technical conditions, the high resolutions and large breadths of images obtained by remote sensing satellites are mutually contradictory. The invention discloses a method for obtaining high-resolution and large-breadth remote sensing images by utilizing a small satellite group. According to the method, multiple satellites are grouped to be operated, so that high-resolution and large-breadth earth observation images can be obtained. By adoption of the method, each small satellite has a strict requirement for resolutions of observation instruments and the breadth requirement can be properly broadened; and after orbit injection, all the small satellites can be uniformly distributed on a same orbit plane, and also can be operated in a relatively concentrated way. When an observation task is carried out for each time, a large area is segmented, on and off moments of reasonable observation instruments are distributed for each satellite, lateral tilt angle and pitch angles of satellite attitudes are set, finally, images obtained by all the small satellites are spliced to realize one-time transition of the satellites and obtain high-resolution and large-breadth earth observation images, and if the quantity of the small satellites in the satellite group is more, the breadths of the images are larger. Moreover, by adoption of a satellite group design, the risk of causing the whole satellite task to lose efficacy due to the failure of individual satellites is reduced.

The invention discloses a construction method of a space-based space target monitoring photoelectric fence system. The method comprises the steps of: determining observation initial parameters according to parameter information of an observation object and parameter information of observation satellites; determining an observation distance and an observation angle according to the observation initial parameters; determining a load requirement according to the observation distance and the observation angle; and constructing the space-based space target monitoring photoelectric fence system according to the load requirement. Through the constructed space-based space target monitoring photoelectric fence system, multiple observation satellites are arranged on a same orbit plane, and the observation range of each observation satellite is intersected with that of the adjacent observation satellites to form an observation loop, so that the space targets running on low orbits can pass throughthe observation loop in each circle and then high-aging observation can be realized.

The invention discloses a high-orbit target approaching observation method based on finite time constraint. The method comprises: determining the relative position relation between a high-orbit service spacecraft and a target spacecraft; determining an approaching mode, an observation moment and an observation position; calculating the relative longitude drift rate, the orbital transfer amount andthe orbital transfer moment; the high-orbit service spacecraft realizing the rounding and floating of the orbit through twice orbital transfer; calculating a traversing moment set of the high-orbit service spacecraft traversing the target orbit plane; calculating an included angle between the orbit plane of the high-orbit service spacecraft and the target orbit plane; selecting an intermediate moment of the crossing moment set to carry out orbital transfer on the high-orbit service spacecraft, so that an orbital plane of the high-orbit service spacecraft coincides with a target orbital plane;and the high-orbit service spacecraft running on the new orbit after orbital transfer for 12 hours before a preset observation moment; and the high-orbit service spacecraft reaching a predetermined observation position through one-time orbital transfer. According to the method, the problems of non-standardization and complex illumination condition constraints of the current approach process method are solved, and the development of subsequent high-orbit observation tasks is supported.

The invention relates to a distributed remote sensing satellite system. The system comprises a plurality of first satellites and a plurality of second satellites, the plurality of first satellites arelow-orbit remote sensing satellites and are distributed on at least two orbit planes; wherein each of the at least two orbit planes is provided with at least three first satellites, the second satellite is a geosynchronous orbit satellite, and remote sensing data collected by the first satellite can be directly transmitted to a ground station or indirectly transmitted to the ground station through the corresponding second satellite. High-definition remote sensing images are collected through the low-orbit remote sensing satellites, remote sensing data can be transmitted to the ground stationby means of the synchronous orbit satellites, and the transmission efficiency of the remote sensing data is greatly improved and guaranteed.