18 results about "Sun-synchronous orbit" patented technology

The invention provides a solar synchronous orbit quick response SAR satellite load platform integrated configuration, which comprises a satellite body, an SAR antenna, a solar array, a phased array data transmission antenna, a star sensor and a body-mounted solar cell, the satellite body is in a quadrangular prism shape, the cross section of the satellite body is an isosceles trapezoid, the satellite body comprises a group of opposite parallel surfaces and a group of opposite inclined surfaces, the phased array data transmission antenna and the star sensor are respectively arranged on two parallel surfaces, and the SAR antenna and the solar array are respectively arranged on two inclined surfaces; the body-mounted solar cell piece and the solar array are installed on the same inclined plane and located on the inner side of the solar array, and the solar array comprises two solar array plates and has an unfolded state and a folded state. According to the method, load platform integration is taken as a design thought, so that the AIT process of the SAR satellite and the time required by AIT are greatly shortened; meanwhile, the attitude requirements of energy and load work are met, and the reliability of the initial injection stage is improved through sufficient energy.

The invention relates to a generalized carrier rocket trajectory calculation method which comprises the steps that a target orbit and constraint conditions corresponding to the target orbit are set, the target orbit comprises a sun-synchronous orbit or a low-dip-angle orbit, and the constraint conditions at least comprise the geocentric vector diameter, the absolute speed and the flight path angle; input parameters are simplified, wherein the input parameters comprise configuration parameters, mass parameters, pneumatic parameters and engine parameters; and based on the constraint condition of the target orbit, the corresponding control variable and the input parameter, calculating the value of the control variable through a Newton iteration method so as to obtain the trajectory of the carrier rocket. The invention provides a universal carrier rocket trajectory calculation method. The carrying capacity of a carrier rocket is ensured by optimizing a typical target orbit; according to the method, input parameters of trajectory calculation are simplified and carded, the trajectory is obtained through calculation of the Newton iteration method, the adaptability to different configuration schemes is improved on the premise that the calculation accuracy is guaranteed, rapid trajectory iteration calculation can be conducted, and generalization is achieved.

A method of sun-synchronous orbit design for remote sensing satellite based on revisit characteristics is proposed. The relationship between revisit and ground synchronization is established by Q value. The mathematical model of Q is established by node revisit sequence in a certain interval. Then the coverage rate is obtained by the mixed node access sequence and coverage area of orbit altitude and observation angle. Finally, the coverage rate of remote sensing satellite in sun-synchronous orbit is demonstrated from three aspects of time-varying coverage rate, minimum coverage time and minimum observation angle, which provides a reference for the design of sun-synchronous orbit of remote sensing satellite.

The invention provides a non-sun-synchronous orbit satellite energy optimal attitude sailboard control method and system. The method comprises the steps that S1, the illumination angle of each orbit in the satellite life period is calculated according to collected orbit parameters; s2, calculating the illumination area half angle and the illumination duration under the illumination angle of each track; s3, calculating the illumination energy of each track of the sailboard driving shaft under the combination of different swinging positions and different flight attitudes of the satellite; s4, comparing illumination energy of each track, and preferably selecting a combination of a posture with the maximum energy in each period and a sailboard driving position; and S5, obtaining a control strategy for the attitude of all orbital circles of satellites and the position of a sailboard driving shaft in the whole life period of the satellites. By comparing various feasible in-orbit attitude and sailboard position combinations, a sailboard position and attitude control method with optimal energy can be obtained through optimization, the in-orbit receiving energy of a two-dimensional driving solar sailboard satellite is improved, the required sailboard area is reduced under the same power consumption requirement, and the weight of the whole satellite is reduced.

The present invention provides a satellite system and method for collaborative detection and imaging of low-orbit targets from a low-orbit orbit. The satellite system includes three sun-synchronous orbit reconnaissance satellites with orbital altitudes of 600km, 800km, and 1000km. The satellite carries a camera payload with optical imaging and detection capabilities. The satellite has three working modes: flyby imaging, fly-by imaging, and joint detection and positioning mode. The advantage is that it can perform real-time multi-angle detection of satellites within the orbital altitude range of 500 to 1100km, provide high-precision positioning and orbit determination of targets; and perform high-frequency flyby imaging, and the minimum imaging distance is short (less than or equal to 100km), the imaging effect is good, and it has multiple observation angles. It can quickly form a fly-by imaging configuration for sun-synchronous orbit satellites within the orbital altitude range of 500 to 1100km, and perform fly-by imaging of targets.

The invention discloses an autonomous control method for a sun synchronous regression orbit, and belongs to the technical field of spacecraft orbit design and autonomous control, and the method comprises the steps: executing an in-plane and out-plane instruction calculation step based on a current instantaneous orbit element vector when determining that a satellite passes through a descending node based on a real-time position vector and a real-time adjacent moment position vector; obtaining in-plane maneuvering final execution time and an out-plane maneuvering instruction; and executing in-plane maneuver based on the in-plane maneuver final execution time, and executing out-plane maneuver when the satellite passes through the ascending node based on the out-plane maneuver instruction. The invention provides an autonomous control method for a sun-synchronous regression orbit, and solves the technical problem that an existing high-precision autonomous orbit control method for the sun-synchronous strict regression orbit is not suitable for the sun-synchronous strict regression orbit at a non-morning-night non-midnight intersection point. And the long-term stability and robustness of the control are enhanced while the high precision of the autonomous control of the sun-synchronous orbit is ensured.

The invention provides an illumination constraint-based sun-synchronous orbit satellite imaging probability analysis method, which comprises the following steps of: determining an orbit height and a latitude argument under a set time interval in an orbit period of a satellite; determining the geographic latitude, the latitude argument and the orbit height of the sub-satellite point of the satellite at each position of the orbit descending section; determining a geocentric field angle of the satellite at each position pointing edge projection point and a sub-satellite point in the orbit descending section; determining whether the to-be-analyzed latitude with the visible light imaging condition points to the position of the edge projection point or not in the orbit descending section of the satellite; determining the local time of the boundary line position after the solar altitude constraint is considered; comparing whether the latitude of the to-be-analyzed satellite with the visible light imaging condition points to the position of the edge projection point in the orbit descending section, and considering the local position of the boundary line position after solar altitude constraint, judging whether the visible light imaging condition is met, traversing the time of one year and the related latitude, and carrying out statistics to obtain the visible light imaging probability of the satellite.

The application provides a satellite configuration and switching method of a multi-band planar phased array antenna, which comprises a satellite platform 1, a flat panel antenna group and a solar cell array 5, the flat panel antenna group comprises multiple flat panel antennas of different frequency bands and is connected to the top plate of the satellite platform 1, the solar cell array 5 is connected to the bottom plate of the satellite platform 1, when in a launching state, the multiple flat panel antennas of different frequency bands are all folded and close to the first side surface, and the solar cell array 5 is folded and close to the second side surface; when in a flying state, the multiple flat panel antennas of different frequency bands are all unfolded until the multiple flat panel antennas of different frequency bands and the top plate of the satellite platform 1 are located in the same plane, and the solar cell array 5 is unfolded until the solar cell array 5 and the bottom plate of the satellite platform 1 are located in the same plane. The satellite configuration has stable structure and can bear large-size and large-mass load, and has high adaptability to satellites with flat panel type effective load in the current sun-synchronous orbit and low-inclination orbit.

An autonomous mission planning method for ultra-low sun-synchronous orbit optical remote sensing satellites includes the following steps: extrapolating the satellite orbit considering atmospheric drag and thrust effects; planning orbit maintenance arcs according to an autonomous orbit maintenance strategy; planning imaging and data transmission arcs respectively according to imaging and data transmission mission requirements; eliminating overlapping orbit maintenance, imaging, and data transmission arcs based on priority; estimating whether the planned arcs meet the energy requirements; if not, eliminating the planned arcs; updating the surface-to-mass ratio based on the planned arcs, and replanning using the updated surface-to-mass ratio until the planned arcs remain unchanged compared to the previous one; completing the autonomous mission planning for the next few orbits. This invention uses electric propulsion for autonomous orbit maintenance as a key constraint in mission planning, prioritizing satellite energy for orbit maintenance, and on this basis, maximizing the performance of imaging and data transmission missions.

The invention relates to the technical field of satellite orbits, in particular to a sun-synchronous orbit dip angle offset calculation method and device, equipment and a medium. According to the method, an orbit perturbation equation containing earth non-spherical gravity and atmospheric damping perturbation is constructed, and the influence of semi-major axis attenuation caused by atmospheric resistance on the inclination angle offset is considered. According to the method, the influence of atmospheric damping on the orbit height and different orbit height maintaining strategies are comprehensively considered, more accurate orbit injection parameters are calculated in combination with the prior art, and a more reliable guarantee is provided for subsequent on-orbit normal operation of a spacecraft. In the initial stage of orbit design, more accurate spacecraft orbit injection parameters can be calculated by introducing more perturbation models. The orbit height maintaining strategy is also considered, which means that the later-stage spacecraft in-orbit operation and maintenance cost is also considered, the engineering implementation is more reliable, the result is more accurate, and meanwhile, flexible adjustment can be realized by fully combining actual conditions.

The present invention provides a global carbon inventory satellite, comprising: a frozen orbit unit, configured to set an orbital inclination, so that the global carbon inventory satellite operates on a frozen orbit, with the apogee of the frozen orbit frozen above the latitude of an area with intensive human activities; and a sun-synchronous orbit unit, configured to enable the global carbon inventory satellite to continue operating on the sun-synchronous orbit, so that the global carbon inventory satellite is always in an illuminated area when it operates to the apogee; wherein the relationship between the perigee orbit altitude and the apogee orbit altitude is obtained based on the value of the ascending node of the sun-synchronous orbit, the value of the orbital inclination, a first function, and a second function; the first function represents the relationship between the semi-major axis of the orbit and the perigee orbit altitude and the apogee orbit altitude; the second function represents the relationship between the orbital eccentricity and the perigee orbit altitude and the apogee orbit altitude.

The application provides a satellite system for observing GEO targets from low orbit and an observation method thereof. The satellite system is composed of a plurality of sun-synchronous orbit satellites, and the orbit altitudes of the plurality of sun-synchronous orbit satellites are distributed in the 300-1500 km orbit altitude. The plurality of sun-synchronous orbit satellites all adopt a north polar staring observation method, that is, the satellites work 20-30 minutes before and after passing the north polar region, the satellite reference attitude is a geodetic attitude, based on the VVLH coordinate system, the camera is fixed along the satellite Y axis, the satellite adjusts the longitude direction pointing around the Z axis of the VVLH coordinate system to ensure the light avoidance requirement of the load, the satellite adjusts the latitude direction around the X axis of the VVLH coordinate system to point to the latitude region between-15° and 15° corresponding to the GEO band, and the pointing of the latitude band of each satellite at different times within one year is calculated and distributed according to a planning scheme. The miniaturized satellite system can realize high-time-efficiency observation of GEO targets.

The present invention discloses a design method for a solar panel device of a sun-synchronous orbit satellite and a solar panel device designed by using this method, belonging to the technical field of satellite solar panels, and solving the problems of difficult attitude adjustment of existing solar panels, high manufacturing cost and long development cycle. The method of the present invention includes: establishing a satellite body coordinate system, which coincides with the orbital coordinate system; making the rotation axis of the solar panel coincide with the rotation axis of the solar vector; calculating the angle β between the solar vector and the orbital plane; according to the angle β, obtaining the angle α between the normal line of the solar panel and the rotation axis of the solar panel, and the angle α and the angle β are complementary angles; designing the solar panel device according to the angle α. The present invention is applicable to the design of solar panel devices of remote sensing satellites with unchanged attitudes.

The invention discloses a universal satellite parametric modeling method, and relates to the technical field of satellite modeling. The modeling method comprises the steps of satellite platform parametric modeling and satellite load parametric modeling. In the satellite platform parametric modeling stage, typical characteristic parameter spaces are constructed for a sun-synchronous orbit satellite platform and a geosynchronous orbit satellite platform respectively; configuring a platform parameter selection principle, a platform model range, platform model composition, a platform model interface and an algorithm related to the platform model; in a satellite load parametric modeling stage, a typical characteristic parameter space is constructed for a satellite remote sensing load; and configuring functions, composition and related algorithms of the load model. The invention provides a universal satellite parametric modeling method which is wide in application.

The spacecraft task time sequence arrangement test system comprises an instruction layer state machine, a task layer state machine and an orbit layer state machine, the starting time and the task duration of a task are determined in an instruction layer, and constraint on a single task is completed; then, in a task layer, according to whether the task has a precondition or a postcondition, arranging the time sequence of the task so as to form a task group formed by sorting and combining a plurality of tasks; and finally, arranging the task group on an orbit layer according to whether the orbit is required or not, so as to complete the task time sequence arrangement of the spacecraft in the on-orbit application scene. The multi-layer state machine is simple in structure and high in interpretability, multi-task arrangement of the satellite in the geosynchronous orbit and the sun-synchronous orbit can be met at the same time, and the efficiency of satellite mode test design is effectively improved.

The present application relates to a kind of flexible solar wing for sun synchronous orbit satellite constellation and on-orbit control method, solar wing is flexible stacked structure, stack on satellite body in retracted state, solar wing includes oppositely arranged single wing, the single wing includes: substrate assembly, it is configured to carry out photoelectric conversion, the substrate assembly includes: rigid substrate and several flexible substrates;Unfolding mechanism, it is configured to provide unfolding stiffness for solar wing, to ensure that the flexible substrate keeps planar form;And unfolding motor, it is configured to drive the unfolding mechanism unfolding or folding, to realize the retraction or unfolding of solar wing;First solar wing driving mechanism and lifting mechanism are provided between oppositely arranged single wing.The present application is overall used flexible stacked structure form retraction envelope size small, light in weight, does not need additional use satellite attitude control system to carry out attitude adjustment and realizes sun tracking.

The method for calculating the collision avoidance area of space targets in a sun-synchronous orbit disclosed by the present invention calculates the areas of various space targets in a sun-synchronous orbit, takes into account the semi-major axis orbit determination deviation and the semi-major axis decay rate caused by atmospheric drag, uses the latest space target catalog data, and iteratively calculates the area of the collision warning of space targets in a sun-synchronous orbit, and finally gives the area for orbit control. The present invention takes into account the orbit perturbation caused by the particularity of the sun-synchronous orbit, and derives the collision avoidance area of space targets in a sun-synchronous orbit. Compared with the traditional calculation of the collision avoidance area of space targets in a sun-synchronous orbit, more accurate results can be obtained.

The present invention discloses an occultation atmospheric sounding system based on a final-stage cluster. By loading a GNSS occultation sounding payload on a final-stage cluster with an orbital retention platform, the final-stage cluster includes final-stage satellites in a number of sun-synchronous orbit missions and inclined orbit missions. These final-stage satellites adjust their attitudes through the orbital retention platform, and receive direct signals of GNSS navigation stars, rising occultation signals and falling occultation signals of GNSS occultations through the GNSS occultation sounding payload, and analyze parameter data including ionospheric electron density, atmospheric refractive index and temperature and humidity profiles to achieve atmospheric sounding. By making full use of the final-stage satellites in the high-density launch of the launch vehicle for adaptive modification, a giant hybrid detection network is formed with the existing GNSS occultation sounding satellite constellation, which can effectively improve the spatio-temporal resolution and detection accuracy of GNSS occultation atmospheric sounding.