263results about "Cosmonautic vehicle tracking" patented technology

A wireless power transfer system includes: a plurality of power transmitters, each of which transmits a microwave; and a rectenna base station which receives the microwave to generate power. The rectenna base station includes: a rectenna; and control section which specifies an identification code for identifying each power transmitter and generates a command signal to change a phase of the power transmitter specified by identification code so as to increase a power value received at the rectenna. Each of the power transmitters comprises: a plurality of transmission antenna elements, each of which transmits the microwave to the rectenna base station; and a phase controller which makes phase change of the microwave based on the command signal from the phase monitor and control section of the rectenna base station if the identification code matches a stored identification code.

A rocket-powered vehicle used for entertainment, namely races, exhibitions, competitions, and revenue-generating events.

A method includes: a) producing initial scheduling plans based on requests related to tasks to be performed within a time period by a remote sensing satellite; wherein each initial scheduling plan schedules respective tasks, which do not conflict with each other in time and in using satellite resources; and each task is scheduled in at least one of the initial scheduling plans; b) applying a genetic-algorithm-based processing to the initial scheduling plans to produce a genetic-algorithm-based scheduling plan which is for mission objectives, and complies with constraints related to the satellite resources, to the tasks, and to the time period; and c) applying a simulated-annealing-based processing to the genetic-algorithm-based scheduling plan to produce a simulated-annealing-based scheduling plan that fits the mission objectives, complies with the constraints, and in which a larger number of tasks is scheduled than in the genetic-algorithm-based scheduling plan.

A universal micro-satellite integrated test platform based on PXI system is formed by interconnecting tested object, direct measurement-control layer, central control management layer and terminal monitoring layer to each other.

The invention provides a control method and a control device used for micro-satellite cluster flying based on atmosphere resistance, and relates to the technical field of micro-aircraft and astronautics measuring and controlling. The signal output terminal of a GPS receiver is connected with the signal input terminal of a track determining device; the signal output terminals of a solar battery array and a three-axis magnetometer are connected with the input terminal of an attitude fusion filter; the output terminal of the three-axis magnetometer is connected with the input terminal of a pure magnetic measuring attitude fusion filter; the signal output terminal of the attitude fusion filter is connected with the second input terminal of a satellite line; the first output terminal of the satellite chain is connected with the input terminal of a cluster formation control law; the second output terminal of the satellite chain is connected with the input terminal of the satellite attitude control law; the formation and attitude of the satellite cluster flying are controlled by controlling a pneumatic plate. The method and the device achieve the objects of long service life of the aircraft, low power dissipation, low cost and no pollution on the basis of atmosphere resistance.

The invention relates to a different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method. The different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method aims at solving the problems that the uncertainty of inertia of a spacecraft is not considered in the prior art, the convergence time can not be freely adjusted depending on the state initial value, and compensating moment generated in the singular direction of a flywheel needs to be designed artificially. The method comprises the particular steps that 1, a tracking satellite and a target satellite are supposed to be located on a different-surface crossover track, and the expected posture needs to be determined; 2, an expected posture tracking control rule is designed; 3, buffeting of the expected posture tracking control rule is eliminated; 4, the expected posture of the crossed points of the tracking satellite and the target satellite changes along with distance between the crossed points of the tracking satellite and the target satellite, a configuration scheme of an execution mechanism is determined according to the expected posture tracking control rule, and the expected posture control torque is solved. The different-surface crossover quick-change track fixed time stable posture pointing direction tracking control method is applied to the field of satellite control.

The invention relates to a method for automatic fault diagnosis of a satellite and a diagnostic system thereof. The method comprises the following steps: dividing the satellite fault into a component level fault, a subsystem level fault and a system level fault; adopting a fault detection method based on a fuzzy set to fuzzify a precise sensor sampling value; defining different alarm weight values for fuzzy subset based on the membership grade belonging to each fuzzy subset; and determining whether to alarm or not by calculating the relation between the summation of the alarm weight values in a time window and an alarm threshold value. The method can effectively overcome the defects of a common fault detection method and effectively reduce the false alarm rate and the rate of missing report; and when certain fault is caused by a plurality of sub-faults, the method performs heuristic search according to the reliability of a system corresponding to each sub-fault and searches a sub-fault node corresponding to a system with low reliability preferentially. Besides, in order to adapt to the synergistic diagnosis between distributed satellites in the future, a synergistic interface is designed to provide diagnosis services for other satellites, thus the aim of the invention is achieved.

The invention relates to a method for controlling a rigid spacecraft for target attitude tracking, and belongs to the technical field of the high-precision and high-stability attitude tracking control of spacecrafts. The method solves the problem that when the attitude tracking spacecraft runs in a low orbit in outer space, the conventional control method cannot eliminate the inherent flutter of a sliding mode variable structure. The method comprises the following steps: 1, establishing a kinetic model and a kinematic model of the rigid spacecraft; 2, setting an attitude tracking error and anexpected attitude parameter of the rigid spacecraft, and combining the attitude tracking error and the expectation attitude parameter with the kinetic model and the kinematic model to establish a mathematical model for the attitude tracking; 3, adopting a control algorithm of a sliding mode variable structure controller to adjust a control law of the mathematical model which is established in thestep 2 and is used for the attitude tracking, and simultaneously combining an observation result of a disturbance observer to modify the control law; and 4, controlling the rigid spacecraft by using the modified control law obtained in the step 3 to realize the attitude tracking. The method is suitable for the attitude tracking of targets running in the outer space.

A collection and distribution system wherein information is collected from one or more aerial vehicles and from one or more ground-based sources before being processed and provided to an end-user.

An iterative method enabling a request plan to be established for an observation satellite. A plan consists in a succession of requests which are associated with pluralities of opportunities for satisfying said requests. The plan must also comply with a plurality of constraints. Each iteration k of the iterative method is made up of the following steps:new opportunity is selected;a provisional plan is derived from the preceding plan k-1 as calculated in-the preceding iteration, and from the new opportunity;provisional plan is verified for compliance with said plurality of constraints;the quality of said provisional plan is evaluated; andit is determined whether the provisional plan should be confirmed as plan k as a function of the quality of said provisional plan and of the quality of said preceding plan k-1.In the method, it is decided whether or not to confirm the provisional plan by a probabilistic metaheuristic of the simulated annealing type, and by the rule for constructing the provisional plan as a function of the newly selected opportunity.

Provided are an automatic operation system for automating satellite control operation, a method thereof, and an automatic satellite ground control system using the same. The object of the present research is to provide an automatic operation system that can reduce operation manpower and cost by automatically operating an entire satellite ground control system, which requires a plurality of operators to be properly operated, all the times, a method thereof, and an automatic satellite ground control system using the same. The satellite ground control system of the present research includes: a mission timeline receiving unit, an operation procedure editing unit, an operation procedure analyzing and code transforming unit, an operation procedure executing unit, a subsystem process state monitoring unit, and an access managing unit for providing interface with each of the subsystems.

A method of determining an orbit of an orbital object includes computing predicted tracking measurement values based on the orbit computed from the initial conditions factoring in any modeled environmental forces and realistic maneuvers; computing the differences between the actual and predicted tracking measurements; determining an improved estimate of the initial conditions that reduces the measurement errors using a minimization or root finding algorithm; after the algorithm has converged, reviewing the hypothetical maneuvers in the force model, taking each value and determining which values came up as near-zero in the minimized solutions and which values came up as those of measurable thrust; determining overall burn duration using the first and last burn times; determining the thrust profile of the orbital object over the observation period using the integrated thrust values; and determining the actual maneuver based on the observation data.

A method and apparatus determines the shape of an orbiting or airborne object. A radar determines the general location and a telescope is directed toward the object to form an image of background stars, which will be occluded by the object. The image is compared with a memorized star map, to identify the region of the image in the map. Stars visible in the map which are not visible in the image are listed. The invisible stars are paired with next adjacent visible stars to form star pairs. The midpoints are identified of lines extending between star pairs. Segment lines are drawn between a midpoint and the next closest midpoint. The segment lines define an outline of the object.

The invention relates to the field of radio navigation and, more specifically, it relates to methods of determining the current values of the orbital position of the artificial satellite of the Earth which is used in a satellite radio navigation system (SRNS), by the ephemeredes and can be used when performing the radio navigation measurements in the equipment of the consumers of the SRNS signals. The variants of a fast method for determining the coordinates of the satellite are proposed, which consist in that the satellite coordinates are determined at preset points of the orbit spaced from each other by a selected time interval, and further determining the velocities of the satellite at the nodal points, while at the points between the nodal points the satellite coordinate are determined in the geocentric inertial system of coordinates are in a system the Greenwich geocentric coordinates using special formulas or the Taylor polynomial of the second or third order, and determining these at the nodal points using the six-order Taylor polynomial.

A satellite communication system has a first deployment of a plurality of satellites deployed in a medium earth orbit and two later deployments of a plurality of satellites deployed in the medium earth orbit. The first deployment is spaced so that the second deployment may be easily deployed and interleaved into the first deployment. A ground terminal is used for communicating with the satellites in the first and second deployments.

The invention is a space measurement technology, involving the improvement to the star sensor calibration method. The invention uses the calibration system which comprising the air cushion platform, single star starlight simulator, star sensor, two-dimensional axial turntable and data processing computer, and its steps are: 1. establish the star sensor imaging model; 2. parameters calibrating. The invention uses no deviation strip to calibration focus and radial distortion coefficient, eliminating the internal parameter estimation effect by the external installation error; the method has simple model and small computing volume, particularly applied to the star sensor calibration based on radial distortion.

Provided are a space debris removing device and a method which enable easy installation of a deceleration device to space debris undergoing a tumbling motion. The space debris removing device includes: a propulsion device (3) for performing approach and attitude control on target debris (1); a capture device (4) having a harpoon (41) which can be ejected toward the target debris (1); an observation device (5) for calculating a capture position (E) and a capture attitude at which the harpoon (41) can be driven into a tank (11) (hollow portion) of the target debris (1) by observing a motion of the target debris (1); a deceleration device (6) directly or indirectly connected to the harpoon (41), for decelerating the target debris (1); and a body part (21) on which the propulsion device (3), the capture device (4), the observation device (5), and the deceleration device (6) are mounted.

The present invention relates to satellite systems and more particularly, to the provision of a satellite system for weather and climate monitoring, communications applications, and scientific research at higher latitudes, referred to as the circumpolar region and defined here as the area with latitudes greater than 60°, in either the northern hemisphere or the southern hemisphere. Contrary to the teachings in the art it has been discovered that a satellite system and method may be provided using satellites in 24 sidereal hour orbits (geosynchronous) with inclinations (70° to 90°), orbital planes, right ascensions and eccentricities (0.275-0.45) chosen to optimize coverage of a particular service area located at high latitudes. A constellation of two satellites can provide continuous coverage of the circumpolar region. The satellites in this orbit avoid most of the Van Allen Belts.

A distance measurement system for locating a geostationary satellite equipped with a transponder includes: a central station emitting a distance measurement signal; and at least one measurement station receiving the emitted signal, including means for measuring the time of arrival of the emitted signal, and transmitted via the transponder of the satellite. The central and measurement stations, of known positions, include synchronization means with a common time base. The transponder is that of a satellite telecommunications system, and has a bandwidth B. The central station includes means for emitting a specific distance measurement signal, distinct from the telecommunications signals transmitted via the transponder, with periodic events and with spread spectrum, the bandwidth of which is included in B.

A starlight simulator with high accuracy consists of light source system formed by concave lens and surface light source, star atlas analog computer being set with star atlas analog algorithm and basic star table for calculating out mapping positions of each navigation star on liquid crystal display system, frequency divider for dividing said mapping positions to different units, display unit for carrying out directly viewing display, liquid crystal display system for displaying distribution of fixed star, parallel light tube being used to set liquid crystal display system on its focal plane and being used to simulate parallel star light of infinite truly under specific visual field and shaft.

Disclosed is a command planning apparatus of a low-earth orbit satellite, and a low-earth orbit satellite control system including the same. The present invention automates the process of executing the command plan for converting the satellite task schedule planned on the ground into telecommands available by the satellite in the low-earth orbit satellite control system, and automatically selects a data set established by the parameters related to the execution task of the satellite according to the mapping rule.

The invention discloses a method for performing real-time measurement on the track and the position of a satellite or an aircraft, relating to space technology. The method comprises: A) establishing a database in a ground central station; B) transmitting ranging codes and telegraph texts by ground branch microminiature measurement terminals, forwarding by the satellite, receiving by heavy aperture antennae of the ground central station to obtain the telegraph text message and the pseudo ranges between the microminiature measurement terminals and the phase center of the heavy aperture antennae of the ground central station; C) subtracting two sections of pseudo ranges which are received by the antennae on the same side of the central station by the same satellite via two adjacent microminiature measurement terminals, using the pseudo range difference to solve the observation equation to obtain the high precision position or track solution of the satellite; D) using the measured values of the pseudo ranges from the same microminiature measurement terminal to different satellites to calculate to obtain the precision positions or pseudo ranges of the microminiature measurement terminals; and E) subtracting the obtain precision positions or pseudo ranges with the predetermined precision positions or pseudo ranges of the microminiature measurement terminals to obtain deviation values, using the deviation values to perturbate the observation equation to correct the track positions of the satellites to obtain the millimeter-level position coordinate values of the satellite track.

The invention relates to the modification to the tracking method of the star tracker. The process includes: it gets the reference star map of the K+1 frame by the star mapping; the every star in the star map collected by the (K+1) frame is matched to the star in the reference star map of the K+1 frame; the character is to map the star by the threshold value mapping; while it searches the star by the zone catalogue; in the matching process, it arrays the sequence firstly then to match. The computing of the invention cost the short time, so it has the high speed.

A GNSS ultra-tight coupling (UTC) receiver architecture applicable to space borne orbit platforms is described. A receiver in accordance with this architecture retains the rotational motion sensors typically found in an inertial measurement unit (IMU) of a conventional UTC receiver, but replaces the IMU accelerometer sensors with precise orbital dynamics models to predict the translational motion of the platform center of gravity (CG). Drag and radiation pressure may be modeled as well. The various models can be implemented in software. The IMU rotational sensors are retained for compensation of the GNSS antenna lever arm effect due to platform rotation.

An attitude estimation method for an on-orbit three-dimensional space object comprises an offline feature library construction step and an online attitude estimation step. The offline feature library construction step comprises: according to a space object three-dimensional model, acquiring multi-viewpoint characteristic views of the object, and extracting geometrical features therefrom to form a geometrical feature library, where the geometrical features comprise an object main body height-width ratio, an object longitudinal symmetry, an object horizontal symmetry, and an object main-axis inclination angle. The online attitude estimation step comprises: preprocessing an on-orbit object image to be tested and extracting features, and matching the extracted features in the geometrical feature library, where an object attitude characterized by a characteristic view corresponding to a matching result is an attitude estimation result. A dimension scale and position relationship between various components of an object are accurately acquired in a three-dimensional modeling stage, thereby ensuring subsequent relatively high matching precision. An attitude estimation system for an on-orbit three-dimensional space object is also provided.

A rocket-powered race for entertaining spectators wherein computer-generated images are optionally provided to at-least partially define a race-course.

The double-view field star sensing detector and its star chart distinguishing method belongs to the field of posture sensor technology. The double-view field star sensing detector has one semi-transparent mirror in angle of 45 deg to the main optical axle installed before the lens system of common star sensing detector, and two lens hoods in the two incident light directions of the semi-transparent mirror to avoid spatial stray light interference. The light rays of the two view fields are transmitted and reflected separately in the semi-transparent mirror, passed through the same lens system and imaged in the same photosensitive detector. Compared with common single view field star sensing detector, the double-view field star sensing detector has higher precision, high chip utilization, small volume, low power consumption, easy installation and use, and other features. In addition, the corresponding star chart distinguishing method is also disclosed.

The invention is a photoelectric detection method for a satellite fault in a geostationary transfer orbit. The method comprises: acquiring a light curve of the satellite in the geostationary transfer orbit by an optical telescope detection system; obtaining an OCS value of a target through analyzing a phase angle sequence light characteristic curve, to evaluate the unfolding state of a solar panel of the satellite; obtaining an angular spin rate of the target through analyzing a time sequence light characteristic curve, to evaluate satellite platform attitude control capacity; and evaluating satellite platform attitude regulation and control capacity through accumulating multi-circle satellite phase angle sequence light characteristic curve, so as to evaluate satellite fault state and provide possibility of rescuing the satellite in time. The method does not depend on a sensor and downlink communication data of the satellite system, and has the property of passive receiving, and characteristics of high sensitivity, high precision, long detection distance, capability of observing in real time, quick data processing and so on.

The invention relates to the space technology, and to the improvement of the star sensor calibration method. The invention uses the calibration system which including air cushion platform, single-star starlight simulator, star tracker, two-dimensional axial turntable platform and data-processing computer, its steps being: establishing point spread function model, establishing a centroid algorithm formula, establishing a centroid algorithm pixel frequency error model, and data acquisition. The invention provides a more practical point spread model; improving satellite sensor calibration precision; the calibration model is simple and easy to calculate.