185 results about "Satellite payload" patented technology

A satellite effective load management system and management method, the system comprises a load distributor, a storage module, a signal transformation processing module, a data multiple connection module and a center processor module, the modularization, integration, standardization effective load management system design conception provided by the invention integrates the data transmission, telemetering measurement remote control, power supply management to be centralized, meanwhile the standard, configurable effective load interface can be designed, so that the resource configuration and the interface configuration of the system are fixed and uniformed, the matching performance problem of effective load managing each task can be effectively solved, the good software and hardware platforms can be provided for time sequence, information and safety management of effective load management, thereby improving the effective load management design level and reducing the complicated degree, corresponding weight and power consumption and so on of the effective load management system, certain basis can be established for improving the whole star performance.

In Low Earth Orbit (LEO) satellite networks data routes for point to point and point to multipoint communication sessions in LEO satellite networks are determined by considering session bandwidth, priority, and the constantly changing terminal-satellite and satellite-satellite connectivity. Multiple routes are computed for a session to facilitate automatic adaptive rerouting by satellite payloads when they encounter network failures or congestion conditions.

An In Orbit Transportation & Recovery System (IOSTAR™) (10) One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

An In Orbit Transportation & Recovery System (IOSTAR(TM)) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR(TM) includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR(TM) (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

The invention relates to a piggyback satellite, a device for and a method of attaching piggyback satellites with multiple payloads, including a primary satellite and at least one secondary payload, such as a deployable micro satellite, affixed instrument package or the like, wherein the secondary payload is mounted on the nadir end of the primary satellite, preferably employing a universal adapter between the primary satellite and the at least one secondary payload that releases at least one secondary payload or the like in a certain orbit in space for allowing the secondary payload to continue its journey independently in space for transmitting and receiving data.

An In Orbit Transportation & Recovery System (IOSTAR™) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A satellite payload structure includes one side that includes at least two equipment mounting panels. Each of the equipment mounting panels are oriented substantially parallel to each other such that when the satellite is deployed, each of the equipment mounting panels are substantially parallel to a vector pointing from the deployed location of the satellite to a point on the earth's surface. The point on the earth's surface is located vertically beneath the satellite such that the vector is substantially normal to the earth's surface.

The invention belongs to the field of mechanical vibration isolation and particularly relates to a multiple-degree-of-freedom vibration isolator and a multiple-degree-of-freedom vibration isolating system for an effective load and a satellite. The multiple-degree-of-freedom vibration isolator provided by the invention consists of a casing, a motion guide rod, a metal wire network material member,flexible hinges, a fine tuning spring, an upper casing end cover and a lower casing end cover and has the advantages of compact structure, reliable performance, long service life and the like. The vibration isolating system consisting of the vibration isolator arranged according to a certain geometry configuration is arranged between the effective load and the satellite and can be used for isolating the vibration transmitted to the effective load from each direction of a satellite platform; and the vibration isolator is connected with the effective load and the satellite through the flexible hinges, and thus a gap, collision and the like are avoided.

The invention discloses a multi-satellite cooperative observation business scheduling method and relates to satellite observation business planning and scheduling technology in a satellite operation control field. The multi-satellite cooperative observation business scheduling method is based on a satellite operation control system, builds a star land resource management model, a satellite orbit calculation model and a satellite access information calculation model, adopts a satellite observation business dynamic planning algorithm to regulate satellite observation business rapidly and dynamically, adopts an automatic processing mechanism to make a satellite load control plan and a data receiving plan immediately, adopts a satellite actual load control instruction compilation business trigger mechanism to compile satellite actual load control instructions in real time, and schedules an electronic signal detection satellite and an imaging observation satellite to complete cooperative observation business. The multi-satellite cooperative observation business scheduling method has the advantages of being rapid in response speed, accurate in calculation method, high in automation and the like and is particularly suitable for a marine moving target multi-satellite cooperative observation field.

An In Orbit Transportation & Recovery System (IOSTAR™) (10). One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A method and apparatus for switching signals through a switch matrix are disclosed. The apparatus comprises an input module and an output module. The input module has a plurality of inputs typically equal to a number of cells in a reuse pattern, and the inputs receiving at least one uplink beam. The input module also has a plurality of outputs, the plurality of outputs is typically equal to the number of subbands in the uplink beam. The output module is coupled to the input module and selectively couples the outputs from the input module to an output of the output module. The method comprises the steps of grouping the uplink signals into a plurality of groups, the number of groups typically equal to a number of cells in the cell-based transmission matrix, where each group comprises a signal from each frequency used in the frequency reuse pattern. Each group of uplink signals is then separated into subband signals, which are then grouped together into groups of similar subband signals. The groups of similar subband signals are then forwarded to demodulators for processing within the satellite system.

A satellite communication system (10) includes a satellite (20) which receives first and second data from sources. Each satellite receives control data from a satellite control center (100). An earth processing center (PC) is arranged to process the data received from the satellites, and a wide band network (30) is arranged to transmit the data to the processing center. A receptor terminal (A) is arranged to receive the first data and to place the first data on the network (30) for transmission to the processing center (PC). A second receptor terminal (B) is arranged to receive the second data and to place the second data on the network (30) for transmission to the processing center (PC).

The invention discloses a management system for payload information of a satellite based on an intelligent high-speed router. The management system comprises an intelligent high-speed router, a payload device, a data handling device, a data storage device and a data transmission device, wherein the intelligent high-speed router consists of a SpaceWire routing module, an intelligent control module and a SpaceWire standard bus interface; and the SpaceWire routing module is used for directly exchanging data from the payload device, the data handling device, the data storage device and the data transmission device, or used for outputting the data to a satellite payload low speed bus network by a CAN (Controller Area Network) interface of the intelligent control modular, and used for transmitting a remote control command from the satellite payload low speed bus network received by the intelligent control module to the payload device, the data handling device, the data storage device and the data transmission device. The system solves the problems that currently the interface of a satellite payload system and a data transmission system is complex, rapid assembly and testing are difficult, and the technology upgrade of a device is difficult.

A large-dip-angle orbiting satellite solar panel orientation method under simple posture control includes the steps that firstly, a satellite centroid orbit coordinate system, a spacecraft body coordinate system and a geocentric equatorial inertia coordinate system are built, the included angle between sunlight and the normal of a satellite orbital plane is obtained, then the included angle is judged, a satellite solar panel is controlled to rotate according to different included angles, the direction of a rotary shaft of the satellite solar panel is adjusted, and large-dip-angle orbiting satellite solar panel orientation control is finished. The method overcomes the defects that in the prior art, in order to meet the requirement that sunlight is perpendicular to a satellite solar panel, the satellite yaw angle of a satellite with a single-degree-of-freedom solar panel driving mechanism is controlled all the time, rotation speed of the solar panel is large, and stability of satellite effective loads is not benefited. Three solar panel orientation adjustment modes are designed according to the angle theta between the vector direction of sunlight and the normal direction of a satellite orbit to meet orientation requirements of a large-dip-angle orbiting satellite solar panel, and the method has the advantages of being high in load stability and easy to operate.

The invention relates to a communication satellite effective load test system simulation platform, and a test rule module stores test rules. A simulation core module reads the data interpretation results of a data analysis module, and a discrete event driving mechanism is adopted to determine the execution time of an event and execute the corresponding event. A simulation design module establishes the node composition and connection of a test system; in establishment, the attributes of a satellite effective load are classified into different working states, and each working state corresponds to one process; and all the working states of the same satellite effective load form a node. An event queue list storages the simulation events and the corresponding execution time. A data acquisition module receives data which is generated in some nodes or processes in simulation, classifies the data and forwards the same to the data analysis module. The data analysis module receives the data which reflects the simulation state, simultaneously acquires a test rule basis which corresponds to a current test project and interprets the received state data according to the test rule basis, and the interpretation results are sent to the simulation core module.

A method is presented for utilizing excess satellite bus power comprising (1) sending a broadband signal in at least one feeder beam in a forward direction from a gateway terminal to a bent pipe repeater satellite for relay to at least one subscriber terminal, wherein the satellite is operable to provide a total amount of bus power, wherein an existing payload consumes an occupied portion of the total bus power, wherein an additional payload consumes a remaining portion of the total bus power and comprises a plurality of satellite-based transmission amplifiers, (2) receiving the broadband signal and amplifying the broadband signal using one of the plurality of satellite-based transmission amplifiers, (3) sending the amplified broadband signal, as one of a plurality of service spot beams, to the at least one subscriber terminal, and (4) receiving and retrieving data from the amplified broadband signal at the at least one subscriber terminal.

The invention discloses an accurate single-satellite one-way timing-based uplink quasi-synchronization time measurement method for low-earth-orbit satellite communication. The method comprises the selection of a spread-spectrum sequence of a terminal, error analysis, a testing method and implementation steps. The method is applied to a single-satellite or overlapped coverage area-free multi-satellite communication system, and is used for a satellite payload to accurately measure the time reference error of a satellite and the ground terminal and ensuring the quasi-synchronization window range of dozens of nanoseconds.

The invention discloses a satellite payload task management system. The satellite payload task management system comprises a data management computer, a data bus, a battery, a distributor, a payload, a data storage multiplexer and a satellite-ground telemetry and remote control processing unit. Through adoption of the satellite payload task management system, integral task management of functions such as power supply and distribution, telemetry and remote control, and health management of a satellite-borne payload by a satellite can be realized.

Provided is an Automatic Identification System-AIS-receiver comprising at least one processing section (PS1, PS2) for synchronizing, demodulating and detecting AIS messages contained in a received signal, said processing steps are carried out separately for a plurality of frequency sub-bands (SB1, SB2, SB3) spanning an AIS channel (CH1, CH2); the receiver is characterized in that: said sub-bands overlap with each others; and said or each processing section is adapted for synchronizing, demodulating and detecting said AIS messages within each sub-band on the basis of timing error and carrier frequency estimations obtained from filtered replicas of said received signal, propagating along respective auxiliary signal paths.

The invention provides a planning method for short-term and effective load work with the capability of increasing energy utilization efficiency of satellites. The planning method comprises following steps: model creation of an energy constraint condition, model creation of the energy constraint condition and an input condition affecting a satellite energy condition and creation of a charge-discharge condition model for a storage battery group. The invention discloses the planning method for short-term and effective load work and an application example. A mathematical model is capable of taking actual running parameters of each orbit ring in satellite running as an input condition for calculation. Compared with a conventional method which only takes a short-term and effective load work mode and work time constraints into consideration under the harshest condition, working modes of loads can be combined at will in one orbit ring. By setting work time and working time of loads at will, the satellites can be more practical.

The invention relates to an earth observation satellite mission effectiveness evaluation index system and an analysis method. The system comprises a mission completion degree index, a time response degree index and a resource utilization rate index. The task fulfillment index comprises a regional target coverage rate, a total task fulfillment rate, a moving target recognition degree, a moving target tracking track fit degree, a task fulfillment benefit and a credibility of a moving target search result. The aging responsiveness index comprises a task response time and a time resolution. The resource utilization index comprises a satellite payload rate, a total satellite side swing number, a total startup time, a total startup number, a large-capacity memory occupancy rate, a ground stationand a relay satellite downlink window utilization rate. The method of the invention can fully embody the physical connection between the operation logic of the system and the indexes, and the evaluation result is intuitionistic, readable and understandable, the method of the invention is an evaluation method with practical physical significance and effectively overcomes the subjectivity and fuzziness existing universally in the efficiency evaluation process.

The invention discloses an image navigation / registration demonstration and verification system based on a real-time simulation platform and a control method. The system comprises a data management computer, an image navigation and registration coprocessor, a payload simulator based on the real-time simulation platform, a dynamics and attitude and orbit control computer simulator based on the real-time simulation platform, a payload work instruction generation tool, a performance demonstration, verification and evaluation system and a ground telemetering and telecontrolling system, wherein the payload simulator is used for simulating the in-orbit work state of a satellite payload, and the dynamics and attitude and orbit control computer simulator is used for simulating the in-orbit work state of a satellite attitude and orbit control subsystem. The system and the method can be used for carrying out evaluation, demonstration and verification on a design result during the designing of an image navigation and registration subsystem and can also be applied to failure determination and inversion during the in-orbit operation of a satellite. The system and the method have the advantages that the practical engineering problem that ground verification is carried out when dynamically-compensated image navigation and registration is carried out on the satellite is solved, and the validity of an image navigation and registration design is guaranteed.

The invention provides a satellite effective load sealing protective shade, which comprises a butt joint and installation mechanism, a protective shade protection cover, a locking mechanism and a flicking mechanism, wherein the locking mechanism consists of components such as an electromagnetic switch, a lever mechanism, a stretched spring and the like, the sealing protective shade adopts an unlocking method of using the electromagnetic switch for providing the unlocking motive power, the protective shade protection cover is tightly locked to the preset position through the pulling force of the stretched spring arranged in the locking mechanism during locking, and the protective shade is ensured to be tightly locked by the mechanical method. Through a method of electrifying the electromagnetic switch arranged in the locking mechanism during unlocking, the locking mechanism is unlocked through overcoming the stretching force of the stretched spring, and the sealing protective shade is opened under the effect of the flicking mechanism. The structure of the invention can protect the effective load internal components from being polluted by external environment, simultaneously protectthe internal low-intensity components from being impacted by external airflows in the sending process, and enhance the survival capability in the stress environment, and in addition, the invention has the advantages of simple structure, low cost, reusability, wide environmental temperature range and high reliability.

The invention provides a management and control system for medium and high-orbit satellite payloads. The system comprises a hardware part and a software part, wherein the hardware part includes a CPU as well as an FPGA module, a communication module, a bus isolation driving module and a data memory which are connected with the CPU, and the software part includes guide software, system program software and FPGA software, wherein the guide software and the system program software operate on the CPU and are mainly responsible for system hardware initialization, electrified self checking and data processing, and the FPGA software operates on the FPGA module and is mainly responsible for system decoding control, command control, reset control and data memory collocation control. The management and control system of the invention are applied to commonly-used application of the satellite payloads, such as program reconstruction, communication management and data storage, and can load a variety of programs for loading satellites and make loading systems complete different functions in different conditions. The management and control system has the advantages of high reliability, long service life and anti-radiation performance.

A reusable module is affixed atop a reusable orbital vehicle (OV). Various configurations of the reusable module have identical external dimensions in the region of attachment to the OV, aerodynamic characteristics, and mounting configurations to permit interchangeability. Different configurations can accommodate a variety of missions of different type and duration. The module may be a cargo module, a satellite payload module or a passenger module. The passenger module is provided in a variety of configurations to accommodate a different number of passengers and cargo based on mission parameters.

Methods, systems, and apparatus are disclosed for high power thermal vacuum testing of satellite payloads using pick-up horns. Such pick-up horns can include at least one outer metal wall forming a metal body and at least one interior surface disposed in the metal body, forming at least one chamber in the metal body. The pick-up horn further includes a front metal surface disposed at a front end of the metal body, having at least one opening corresponding to the at least one chamber, and at least one high-power absorbing load disposed within the at least one chamber and in contact with the at least one interior surface. A pick-up horn may further include a serpentine coolant path disposed within the metal body between an outer surface of the at least one outer metal wall and the at least one high-power absorbing load. Related systems and methods are described.