45 results about "Inclined orbit" patented technology

A retro-directive antenna for communicating with a geostationary satellite autonomously detects the direction from which a signal is received, and transmits a beam that points back along the same direction. An array feed is used to illuminate a parabolic reflector. Each feed element of the retro-directive antenna is associated with a unique pointing direction of the beam in the far field. As the transmit energy is switched to different feed elements, the far-field beam is scanned, making it possible to track a geostationary satellite in a slightly inclined orbit. This eliminates the need for mechanical tracking and maintains high antenna gain in the direction of the geostationary satellite. The use of a toroidal reflector with multiple linear array feeds spaced in the azimuth direction enables multi-beam operation, allowing multiple geostationary satellites, spaced by up to fifteen beam widths in azimuth, to be tracked simultaneously and independently.

The present disclosure is directed to an inclined geosynchronous orbit satellite system that can efficiently provide continuous communication to multiple geographic regions across the world using satellites in inclined geosynchronous orbital paths having an equatorial crossing and enabling the reuse of frequencies assigned within GSO orbital locations. The inclined orbit satellite system can include multiple inclined orbit satellites that are capable of co-existing with geostationary satellites to provide continuous uninterrupted service.

The invention relates to a method for improving the orbit-transferring safety of an inclined orbit satellite, wherein the problem of energy-source shortage of the inclined orbit satellite in the period of orbit-transferring motorization under the condition of high solar angles is solved through the reasonable adjustment of flight events by considering the condition of solar altitude angles in the designing process of a flight program, and meanwhile, flight-control operation can be simplified. The discharging time of a storage battery in the period of orbit-transferring ignition can be shortened by designing the flight program according to the method, so that the load power in the power-supplying period of the storage battery is reduced, therefore the discharging depth of the storage battery is reduced, the energy-source safety of the satellite is ensured, and a launch window is not limited by solar illumination angles.

The invention discloses a double-direction solar panel control method applicable to an inclined orbit satellite. The double-direction solar panel control method comprises the steps that a mathematic model of the positional relation between the inclined orbit satellite and the sun is built and used for quantitative description of the motion law of the sun relative to the satellite; a solar panel driving error is defined according to the mathematic model and used as input of solar panel driving control; the law of solar panel capture control is determined, so that a solar panel is made to find the sun; the law of solar panel tracing control is determined, so that solar panel sun-targeted orientation is realized through sun tracing. According to the double-direction solar panel control method applicable to the inclined orbit satellite, the solar panel driving law and satellite yawing maneuvering are combined, and the problem that the inclined orbit satellite realizes sun-targeted orientation through a one-dimensional solar panel driving mechanism is solved; judgment and adjustment can be carried out autonomously according to the motion direction of the sun according to the forward flight and inverted flight of the satellite, and correct capture of the sun by the solar panel under different initial conditions is realized; forward driving control and inverted driving control of sun-targeted orientation of the solar panel are realized according to the forward flight and inverted flight of the satellite.

A method of flying a constellation of inclined geosynchronous satellites at the same station longitude with specific spacing but without the possibility of collision and provides the basic equations defining the initial positions of satellites such that the satellites will continue to remain in synchronized positions relative to each other for a number of years with little or no north-south positioning. In preferred embodiments the number of satellites in the constellation is five or ten. Communication with the satellites in the constellation is provided with existing prior art tracking radio systems.

The invention relates to an inclined orbit spacecraft energy obtaining method based on yaw steering. The inclined orbit spacecraft energy obtaining method based on yaw steering includes the steps of calculating the solar unit vector of a J2000 inertial coordinate system according to Julian century numbers, calculating a solar elevation angle beta according to the solar unit vector, the satellite position and the satellite running speed, calculating a transfer matrix Aoi from the J2000 inertial coordinate system to an orbital coordinate system according to the satellite position and the satellite running speed, calculating the solar angular altitude theta s of the orbital coordinate system according to the solar unit vector and the transfer matrix Aoi from the J2000 inertial coordinate system to the orbital coordinate system, calculating the yaw steering angle psi and a sailboard drive angle R according to the solar elevation angle beta and the solar angular altitude theta s, and enabling the solar vector to be perpendicular to the plane of the solar sailboard through yaw posture control and sailboard one-dimensional driving to ensure that a spacecraft can obtain energy. The normal of the solar sailboard of the spacecraft can point to the sun under the situation of the large change range of the solar elevation angle and the complex lighting condition, so that it is ensured that the spacecraft can obtain solar energy under the situation of an inclined orbit.

A satellite communication ground station configured for communicating over an inclined orbit geostationary satellite may include a tracking antenna having three fixed axis and one moving axis, a motor for swinging the antenna along the moving axis, a controller for controlling the motor, a receiver configured to receive a signal arriving from the satellite via the tracking antenna and an estimator configured to estimate reception quality of a signal received by the receiver. The invention described herein presents a method for tracking an inclined orbit geostationary satellite, the method comprising a learning step and a tracking step, wherein the learning step includes use of a filter for reducing alignment errors and reducing the amount of peaking required for aligning the antenna with the satellite and tracking its movement.

An inexpensive method and system for maintaining communication between fixed receivers on the earth surface having fixed antennas and inclined orbit geostationary satellites. In preferred embodiments the fixed antennas of each receiver are provided with more than one feed horn (such as three or more feed horns) each positioned to form a beam with its antenna in a slightly different direction from beams of the other feed horns in the north-south direction. An algorithm is developed based on the known daily swings of the inclined orbit geostationary satellite which determines which of the feed horns provides a beam providing the maximum signal from the satellite.

The present disclosure is directed to an inclined geosynchronous orbit satellite system that can efficiently provide continuous communication to multiple geographic regions across the world using satellites in inclined geosynchronous orbital paths having an equatorial crossing and enabling the reuse of frequencies assigned within GSO orbital locations. The inclined orbit satellite system can include multiple inclined orbit satellites that are capable of co-existing with geostationary satellites to provide continuous uninterrupted service.

The invention discloses an autonomous control method for inclined orbit satellite yaw maneuvering. The method comprises the steps of measuring and calculating the solar altitude and the satellite yaw axis attitude at the current moment; obtaining the moving direction of the sun relative to a satellite according to the solar altitudes at two different moments; determining the expected attitude of the satellite yaw axis according to the solar altitude at the current moment; determining a yaw maneuvering target value and a yaw maneuvering mode according to the solar altitude at the current moment, the satellite yaw axis attitude at the current moment and the moving direction of the sun relative to the satellite; starting control through a satellite actuator, and controlling the satellite to change to the expected attitude of the satellite yaw axis from the satellite yaw axis attitude at the current moment to carry out yaw axis attitude maneuvering. According to the autonomous control method for the inclined orbit satellite yaw maneuvering, the satellite forms an approximately fixed nightside through the yaw axis attitude maneuvering of the satellite, layout and design of star sensors can allow the star sensors to avoid sun shine, the star sensors are protected, and continuous output of measured data of the star sensors is guaranteed.

An advanced multiple-beam fixed ground terminal is achieved that is capable of simultaneously tracking multiple inclined orbit satellites, increasing and suppressing gain in multiple directions. The fixed user terminal equipped with digital beam-forming and null-forming technique can track and identify signals from multiple inclined orbit satellites at the same time. This technique enables a geostationary satellite drift to an inclined orbit without losing communication with ground terminals which not only increase the life span of an inclined orbit satellite, but also relieve the scarcity of geo-stationary orbit. In extreme cases, satellite can be placed in the same slot which further enhanced the usage of geosynchronous orbits. Another present invention is to from double nulls whose null width is much wider than a single null. A wider null increases the system robustness to frequency drift and change of signal direction, thus in turn reduce the system's complexity by lowering update beam wave vectors. To use the same beam wave vector on wider frequency spans, an FIR filter need to be designed according to system requirements.

The invention discloses an inclined orbit satellite yaw maneuvering opportunity judgment method. The method comprises the steps that inclined orbit satellite yaw maneuvering constraint conditions are determined and judgment threshold values of the constraint conditions are determined; according to the constraint conditions and the judgment threshold values of the constraint conditions, the yaw maneuvering opportunity is judged. According to the inclined orbit satellite yaw maneuvering opportunity judgment method, the inclined orbit satellite in-orbit automatic yaw maneuvering constraint conditions are comprehensively analyzed, the property of the yaw maneuvering constraint conditions is analyzed, conditions and the requirement for the occurrence frequency are met, and the effective and reliable basis is provided for accurately judging the yaw maneuvering opportunity and time; the yaw maneuvering opportunity judgment criteria and the yaw maneuvering starting specific time judgment method are provided, and satellite-borne software can accurately and automatically carry out yaw maneuvering starting and control.

The invention discloses a satellite orbit determination system inserted in a mobile communication network. The satellite orbit determination system comprises a plurality of mobile communication base stations, a satellite data server and a plurality of mobile users. The satellite data server transmits calculated high-precision orbit parameters to the mobile users; the mobile communication base stations send collected original satellite data to the satellite data server; the satellite data server provides the calculated high-precision orbit parameters for a satellite ground control center; the satellite ground control center transmits satellite orbit correction to navigation geosynchronous satellites and inclined orbit geosynchronous satellites through satellite earth stations. Large-scale distributed mobile communication base stations are utilized to measure the original satellite data, and the satellite data server collects measurement data of the mobile communication base stations through networks, analyzes the measurement data, and calculates orbit parameters of the navigation geosynchronous satellites and the inclined orbit geosynchronous satellites to precisely determine satellite positions, so that the objective of improving accuracy of satellite navigation systems is achieved.

Techniques for redirecting a satellite antenna beam from a satellite antenna of a satellite in an inclined orbit to a preferred satellite antenna beam pointing position on the earth are provided. Satellite antenna maneuver data is computed based upon information associated with the satellite, such as telemetry, orbital, and antenna nominal pointing information. Command code sequences are then generated that correspond to the maneuver data. The command codes are executable at the inclined orbit satellite to modify the antenna position. The generated command codes are then transmitted to the inclined orbit satellite to redirect the satellite antenna beam center to the preferred position.

A concealed satellite navigation and positioning system comprises a concealed satellite constellation, a concealed navigation and positioning signal, a concealed navigation and positioning ground master control station, an intelligent receiver and a passive orbit determination station. The concealed satellite constellation is a generic GEO communication satellite constellation and is composed of aC / Ku frequency band GEO commercial communication satellite, an SIGSO small-dip-angle synchronous orbit satellite and an IGSO inclined orbit satellite. The concealed navigation positioning signal is concealed under an original service signal of a generic GEO communication satellite, the signal power is less than one percent of the full-load power of a satellite transponder, the bandwidth of the navigation signal is more than 40MHz, and the signal-to-noise ratio reduction amplitude of the original signal does not exceed 0.3 dB; wherein the concealed navigation positioning ground master controlstation is a concealed navigation positioning fixed station which is fixedly arranged or a concealed navigation positioning vehicle-mounted communication-in-static station which can move along with avehicle; the intelligent receiver comprises an antenna module, a radio frequency module, a baseband processor module and an autonomous navigation module; the passive orbit determination station comprises at least four orbit determination reference stations located on the surface of the earth. The orbit determination base station receiver collects satellite signals, calculates the position correction amount of a measured satellite, counts the orbit determination precision, arranges the satellite position correction amount information into a navigation message, and injects the satellite positioncorrection amount information into the satellite through the concealed navigation positioning ground master control station for positioning by a user.

The invention discloses a sun-facing orientation control method for a single-degree-of-freedom solar panel of an inclined orbit satellite. The method comprises the steps: a solar panel sun-facing control mode is selected according to a satellite sun altitude angle beta; when the absolute value of beta is smaller than or equal to 15 degrees, the satellite does not yaw, and a solar panel driving mechanism drives the solar panel to turn over. When the absolute value of beta is larger than or equal 15 degrees and smaller than or equal to 45 degrees, the satellite yaws at a fixed angle, and the solar panel driving mechanism drives the solar panel to turn over. When the absolute value of beta is larger than 45 degrees, the satellite yaws in real time, and the solar panel driving mechanism drivesthe solar panel to turn over. The invention discloses the sun-facing orientation control method for the single-degree-of-freedom solar panel of the inclined orbit satellite, the relation between thesolar energy obtaining efficiency and the attitude control precision is comprehensively considered, on the premise that the good sun-facing performance of the solar panel is guaranteed, energy consumption and the design difficulty of an attitude and orbit control system can be reduced, the satellite attitude control precision is improved, and a satellite can achieve the better comprehensive performance.

The invention provides a solar cell array control strategy method suitable for a yaw maneuvering satellite. The method comprises the steps that the light condition of an orbit where the satellite is located is obtained; a satellite yaw maneuvering scheme is determined according to the light condition; and a solar cell array control strategy is obtained according to the satellite yaw maneuvering scheme. The problem of energy of on-orbit flying of the inclined orbit satellite is solved.

The invention discloses a method for maneuvering large-angle yaw attitudes by the aid of momentum wheels. The method has the advantages that existing yaw attitude offset ability of an onboard control subsystem is utilized and is combined with a ground remote control process, offset target quantities are repeatedly injected by the method at multiple steps, and a control procedure with a high offset quantity is divided into a plurality of control procedures with small offset quantities, so that large-angle yaw is finally maneuvered via the multiple control procedures with the small offset quantities; the method is smart in design, simple in implementation and reliable in running, and actual requirements of capturing positions of inclined orbit satellites during flight control and keeping the inclined orbit satellites at in-orbit positions for a long time can be effectively met on the premise that onboard software is not modified.

The invention provides a zero-bias lasting day number determination method suitable for a navigation inclined orbit satellite. The zero-bias lasting day number determination method comprises the steps that an included angle between the ecliptic plane and the orbit plane is worked out according to the orbit ascending node right ascension and the orbit inclined angle of the navigation inclined orbit satellite; the arc section of the sun on the ecliptic during the satellite zero-bias period is obtained according to the included angle between the ecliptic plane and the orbit plane as well as an included angle between the sun vector in the dynamic deflection zero-bias condition and the orbit plane; and the satellite zero-bias lasting day number is determined according to the arc section of the sun on the ecliptic during the satellite zero-bias period. According to the zero-bias lasting day number determination method suitable for the navigation inclined orbit satellite, the satellite zero-bias lasting day number can be worked out through geometrical analysis according to the two orbit elements of the satellite during the orbit design period and after the design is completed; and the method is simple in calculation and little in required orbit information and meanwhile also can be reversely applied to orbit design under constraint of the zero-bias lasting day number.

The invention provides an overall optimization design method based on inclined orbit satellite illumination conditions, which comprises the following steps: a calculation step: obtaining a solar vector by on-board orbit parameter recursion or sun sensor measurement, calculating to obtain a solar altitude angle, and representing the energy supply change and heat dissipation surface change of the whole satellite by taking the solar altitude angle as a parameter; determining a flight scheme, wherein a satellite U-turn flight scheme considering whole satellite energy and fixed heat dissipation surface design is adopted; determining satellite flight polarity: determining the satellite flight polarity according to the sun vector change direction; an applicability determination step: adaptabilityof a shadow area judgment method and a shadow area turning-around flight scheme; and a basis selection step: selecting a basis for the turn-around flight time. Starting from the perspective of overall optimization design of the whole satellite, thermal control design of the whole satellite is facilitated; enough illumination time is ensured, so that the energy of the whole satellite is balanced;and the constraint of sensor and load layout design is also solved by the fixed illumination surface.

The embodiment of the invention discloses a control method for a double-shaft solar wing driving mechanism of a near-earth inclined orbit satellite. The mechanism comprises a first shaft mechanism, a satellite body, a second shaft mechanism and a solar wing, the first shaft mechanism is installed on a star body, the second shaft mechanism is connected with the first shaft mechanism through a connecting rod, and the solar wing is connected with the second shaft mechanism. According to the installation and control method for the double-shaft solar wing driving mechanism of the near-earth inclined orbit satellite, the configuration of a mechanism system is optimized, and the interference of the movement of a solar panel on the attitude of the whole satellite is reduced. Meanwhile, a control parameter calculation method and detailed steps of the solar wing driving mechanism under different illumination conditions are given, and the satellite can obtain solar energy to the maximum extent on the premise that the requirement of the whole satellite attitude control system is met.

The invention discloses an occultation atmospheric detection system based on a tail sub-level cluster, and the system is characterized in that a GNSS occultation detection load is loaded on an orbit-remaining tail sub-level cluster provided with an orbit-remaining platform, and the orbit-remaining tail sub-level cluster comprises a plurality of orbit-remaining tail sub-levels on a sun-synchronous orbit task and an inclined orbit task; the orbit remaining tail sub-levels adjust the attitude through an orbit remaining platform, receive a direct signal of a GNSS navigational star and an ascending occultation signal and a descending occultation signal of GNSS occultation through a GNSS occultation detection load, and analyze parameter data including ionosphere electron density, atmospheric refractive index and temperature and humidity profile to realize atmospheric detection. The orbit remaining tail sub-level under high-density launch of a carrier rocket is fully utilized to carry out adaptive modification, the orbit remaining tail sub-level and an existing GNSS occultation detection satellite constellation form a giant hybrid detection network, and the temporal-spatial resolution and the detection precision of GNSS occultation atmosphere detection can be effectively improved.

The invention discloses an inclined orbit satellite platform control method, electronic equipment and a storage medium. The method comprises the steps of determining an optical axis of a star sensor according to a preset star sensor installation rule; and performing attitude control over the star sensor according to a continuous rolling attitude path planning strategy so that the star sensor is enabled not to be interfered by sunlight. According to the inclined orbit satellite platform control method, the star sensor is not influenced by strong light under the condition that the change range of the solar altitude angle is large, and the star sensor is ensured to effectively measure attitude information.

The invention discloses a solar sailboard variable-speed driving method under yaw guidance of an inclined orbit satellite. The solar sailboard variable-speed driving method comprises the following steps: calculating a solar altitude angle and a satellite orbit angle in real time; calculating a theoretical rotating angle and a theoretical driving angular speed of the solar sailboard at the current moment in real time according to the solar altitude angle and the satellite orbit angle; calculating an angle needing to be driven by the solar sailboard at the current moment according to the theoretical rotating angle of the solar sailboard at the current moment; and selecting the gear closest to the theoretical driving angular speed of the solar sailboard at the current moment in the gears of the solar sailboard driving mechanism as a center value, and determining the micro-amplitude adjustment amount according to the angle needing to be driven by the solar sailboard at the current moment, so that the driving angular speed of the solar sailboard at the current moment is obtained. The solar sailboard adopts a variable angular velocity driving method, so that the solar sailboard is always perpendicular to the sun, and satellite energy is guaranteed.

The invention discloses a method for determining a sun-facing optimal fixed yaw angle of an inclined orbit satellite with a solar panel. The method comprises the following steps: as for a given solarelevation angle lambda, taking the step length of the yaw angle theta and the step length of the solar azimuth angle as theta step according to the solving precision, marking the yaw angle and the solar azimuth angle corresponding to each step length as theta k, and as for different theta k, calculating: when lambda is greater than 0, taking a sequence point from a positive zero crossing point theta k to a negative zero crossing point theta k in a negative value interval of theta k as the solved optimal fixed yaw angle; and when lambda is less than 0, the zero crossing point theta k of the sequence points from positive to negative in the positive value interval of theta k is the solved optimal fixed yaw angle. According to the method for determining the sun-facing optimal fixed yaw angle of the inclined orbit satellite with the solar panel, the mutual influence relation between the fixed yaw angle and the operation rule of a panel driving mechanism is comprehensively considered, and finally the theoretically optimal fixed yaw angle is obtained. The satellite is matched with the solar panel driving mechanism to operate at the optimal yaw angle so that the theoretically maximum solarenergy acquisition efficiency in the mode can be realized.