Abstract: Provided are a method for maintaining Walker constellation formation and a terminal device. The method comprises: determining a first offset amount of each satellite within a simulation time period according to parameters of a Walker constellation; performing first offset on each satellite according to the first offset amount to obtain a Walker constellation after the first offset; determining a second offset amount of each satellite within the simulation time period according to parameters of the Walker constellation after the first offset; and superimposing the first offset amount and the second offset amount, and performing second offset on each satellite so as to maintain the formation of the Walker constellation.

Abstract: A scalable, extensible, multi-tenancy multi-mission configurable spacecraft system is provided that allows applications to be deployed and managed across many spacecraft. One embodiment includes a plurality of satellites in orbit, where each satellite includes an antenna, a memory configured to store a non-virtualized operating system and one or more software applications, and a processor connected to the antenna and the memory. The processor is configured to run the non-virtualized operating system and to run the one or more software applications. The processors and the applications can be managed by ground terminals or other satellites.

Abstract: The disclosure includes determining a regression period and a semi-major axis of an orbit, inclination, eccentricity and argument of perigee of the orbit; determining both the number of satellites and the number of orbital planes as n; determining right ascension of an ascending node and a mean anomaly of a first satellite, and sequentially determining right ascension of ascending nodes and mean anomalies of subsequent satellites according to satellite service requirements; determining a set of geostationary orbit satellite networks that need to be coordinated, and width of guard band for interference of non-geostationary satellite constellation on geostationary satellite; at any location on the ground, deployed satellites pass overhead successively along fixed trajectory, a user at ground can simultaneously see satellites when multi-coverage is formed; if satellite trajectory crosses the guard band for interference on the geostationary satellite, then when a currently-accessed satellite enters the guard band

Abstract: The positioning receiver according to the invention comprises means for storing configuration information and information on its reception conditions, which is processed in order to calculate a current and/or predicted precision of the positioning calculation. Advantageously, at least some of this information is integrated into a variation model of a Kalman filter integrated into the receiver. The invention notably allows a more rapid convergence on a reference precision in a mono-frequency operating mode and transitions between a mono-frequency and bi-frequency mode to be smoothed. Advantageously, the precision information is delivered graphically or in numerical form to the user.

Abstract: A system and method of providing an affordable navigation, guidance and control system for arbitrary nano/micro launch vehicles by integrating commercial grade sensors with advanced estimation algorithms in a manner that provides sufficient accuracy of the resulting vehicle state estimates to inject nano/micro satellites into low earth orbits. The system and method uses commercial grade sensors and an advanced sensor-fusion estimator software that estimates and removes the estimated measurement errors and filters noise produced by the commercial grade sensors, resulting in estimated states with suitable accuracy. The filtered data are sent to a guidance and control system where actuator commands are formulated based on the filtered data. A simulated launch and flight of the launch vehicle is performed using the filtered data to validate that the GNC system and launch vehicle are ready for launch.

Abstract: The disclosed technology relates to systems and methods for tasking satellite constellations. A method is disclosed herein for receiving, from a resource database of a satellite control system, knowledge data corresponding to a plurality of components associated with a satellite constellation communications system. The plurality of components can include one or more satellites associated with a constellation. The method includes processing the knowledge data according at least one received mission objective. Processing the knowledge data can include determining a status of at least one satellite in the constellation.

Abstract: A method for controlling the orbit of a satellite in earth orbit. The orbit of the satellite is controlled by commanding, according to a maneuver plan, a propulsion system having at least one thruster and a transporter to move the propulsion system. The maneuver plan includes at least two orbit-control maneuvers. The thrust powers of the propulsion system during the two orbit control maneuvers have respective thrust directions that are not parallel in an inertial frame of reference. Each thrust power is determined to simultaneously control the inclination and the position of the orbit of the satellite as well as to form a momentum that is suitable for unloading a device for storing angular momentum of the satellite in a plane orthogonal to the direction of thrust of the thrust power.

Abstract: Techniques are disclosed for providing a redundant telemetry transmission path on a spacecraft. The spacecraft includes a telemetry (TLM) transmitter having a modulation input coupled by a commandable switch with a signal source. The TLM transmitter is configured to receive an output from the signal source when the commandable switch is closed. The computer is configured to encode TLM data bits onto a TLM downlink signal by actuating the commandable switch, such that a signal from the signal source received by the TLM transmitter is temporally varied between an on condition and an off condition.

Abstract: An operation of a spacecraft is controlled using an inner-loop control determining first control inputs for momentum exchange devices to control an orientation of the spacecraft and an outer-loop control determining second control inputs for thrusters of the spacecraft to concurrently control a pose of the spacecraft and a momentum stored by the momentum exchange devices of the spacecraft. The outer-loop control determines the second control inputs using a model of dynamics of the spacecraft including dynamics of the inner-loop control, such that the outer-loop control accounts for effects of actuation of the momentum exchange devices according to the first control inputs determined by the inner-loop control. The thrusters and the momentum exchange devices are controlled according to at least a portion of the first and the second control inputs.

Abstract: The disclosed technology relates to systems and methods for tasking satellite constellations. A method is disclosed herein for receiving, from a resource database of a satellite control system, knowledge data corresponding to a plurality of components associated with a satellite constellation communications system. The plurality of components can include one or more satellites associated with a constellation. The method includes processing the knowledge data according to at least one received mission objective. Processing the knowledge data can include determining a status of at least one satellite in the constellation.

Abstract: A drive test information exchange method and apparatus are disclosed. The method includes: receiving, by a positioning controller, a positioning request message sent by user equipment UE; sending, by the positioning controller, positioning assistance information to the UE after the positioning controller receives the positioning request message sent by the UE; receiving, by the positioning controller, a measurement result sent by the UE; obtaining, by the positioning controller, positioning information according to the measurement result; and sending, by the positioning controller, drive test information to a drive test unit according to the indication information. The method and the apparatus of the present invention may be used to enable a process of obtaining drive test information by a drive test unit to be simpler, and a delay to be shorter.

Abstract: A spacecraft includes a payload subsystem including a digital channelizer. The digital channelizer provides at least a portion of spacecraft command or telemetry functionality. The spacecraft may optionally also include a telemetry and command (T&C) subsystem, the T&C subsystem including one or more of a command receiver, a command decoder, a telemetry encoder and a telemetry transmitter. The digital channelizer may be communicatively coupled with at least one of the command receiver, the command decoder, the telemetry transmitter and the telemetry encoder.

Abstract: A robotic system has arms, arm processors, arm supervisor, and system supervisor. Each arm includes nodes for controlling motors in the arm. Each node, including each arm processor, detects faults affecting the node, places the node into a safe state upon detecting a fault, propagates a fault notification, diagnoses the fault and classifies it, and sends an error message to the supervisor processor. The arm supervisor may detect faults affecting an arm and also perform fault reaction activities. The system supervisor handles the fault as either a system or local fault depending upon its class. For system faults, a fault notification is sent to the arm processors of non-failed arms so that the non-failed arms are placed in the safe state. For local faults, a degraded operation option is provided to a user and if the fault is classified as recoverable, a recovery option is provided to the user.

Abstract: An apparatus and method for controlling a geostationary orbit satellite is provided. The method including generating remote measurement data by measuring a state of a geostationary orbit satellite, transmitting the remote measurement data, receiving a remote command signal, and controlling an orbit and a pose of the geostationary orbit satellite relative to inclined geosynchronous space debris.

Abstract: A system and method for controlling de-orbit of a spacecraft is presented. Embedded command modules are commanded directly from a central command and telemetry module. Latch valves, thruster valves, and solar wing drive of the spacecraft are operated in response to inputs to the embedded command modules. The spacecraft is maneuvered to a safe disposal orbit in response to commands from the central command and telemetry unit.

Abstract: A hybrid network of kinematic sensors of an AOCS, made up of a star sensor including an optical camera head, and a processing unit provided as the central master processing unit, and additional kinematic sensors, each made up of a sensor element and a processing unit connected to the central processing unit via a first bus. An additional processing unit is equivalent to the processing unit and is a redundant central processing unit. The central processing units and—are connected via an additional bus of a spacecraft provided with the hybrid network with the aid of a central computer. The particular active central processing units-provide all kinematic sensors with a uniform time pulse via a synchronization line, and supply the central computer with hybridized kinematic measuring data formed according to a method for hybridization based on the synchronous kinematic measuring data of the star sensor and the measuring data of the other sensors.

Abstract: A system and method for rapid dissemination of image products. In the system, a data consumer display device sends a geospatial request for a map image of a specific area of interest to a rapid image distribution system (RIDS), which forwards the request to a sensor ground station. The sensor ground station processes data received from a sensor platform and sends the processed data to a georectification processor. The georectification processor creates georectified data and sends the georectified data to the RIDS, which further processes the data, exposing it to data consumers using network optimized data services (e.g., KML/KMZ, TMS, GeoRSS, image chipper) based on geographic coordinates provided in the query that is a smaller subset of the sensor data. The RIDS sends the image product to the data consumer display device for display.

Abstract: A hybrid network of kinematic sensors of an AOCS, made up of a star sensor including an optical camera head, and a processing unit provided as the central master processing unit, and additional kinematic sensors, each made up of a sensor element and a processing unit connected to the central processing unit via a first bus. An additional processing unit is equivalent to the processing unit and is a redundant central processing unit. The central processing units and—are connected via an additional bus of a spacecraft provided with the hybrid network with the aid of a central computer. The particular active central processing units-provide all kinematic sensors with a uniform time pulse via a synchronization line, and supply the central computer with hybridized kinematic measuring data formed according to a method for hybridization based on the synchronous kinematic measuring data of the star sensor and the measuring data of the other sensors.

Abstract: An autonomous system for a satellite which calculates collision paths of debris from anywhere within the spheroid around the satellite by using its radar/ladar data and from data on its own orbit derived by onboard sensors such as star, earth and sun sensors or from stored data sent from its ground control station through the satellite's command subsystem. If a collision would be likely, the system calculates the minimum change in the satellite's orbit to avoid such collision and generates and executes commands for firing on-board orbital control thrusters to put the satellite in a suitable avoidance orbit.

Abstract: A technique to assist guidance techniques for a free-flying inspection vehicle for inspecting a host satellite. The method solves analytically in closed form for relative motion about a circular primary for solutions that are non-drifting, i.e., the orbital periods of the two vehicles are equal, computes the impulsive maneuvers in the primary radial and cross-track directions, and parameterizes these maneuvers and obtain solutions that satisfy constraints, for example collision avoidance or direction of coverage, or optimize quantities, such as time or fuel usage. Apocentral coordinates and a set of four relative orbital parameters are used. The method separates the change in relative velocity (maneuvers) into radial and crosstrack components and uses a waypoint technique to plan the maneuvers.

Abstract: A method for stationing a satellite, comprises: determining a predefined trajectory for stationing of the satellite based on a model of movement of the satellite; determining parameters of the predefined control law of approximation of the trajectory and by minimizing impact on the control law of deviation from a trajectory followed by the satellite using parameters of the control law; determining a state vector of the satellite; determining a deviation between the state vector and the predefined trajectory; determining Lagrange multipliers based on a current state vector of the satellite, on a deviation between current state vector and predefined trajectory and on parameters of the predefined control law; determining parameters of the current control law of the engines based on the Lagrange multipliers and by derivation of a parameter representative of an effect of engines on the real trajectory; and controlling engines based on parameters of the current control law.

Abstract: A technique to assist guidance techniques for a free-flying inspection vehicle for inspecting a host satellite. The method solves analytically in closed form for relative motion about a circular primary for solutions that are non-drifting, i.e., the orbital periods of the two vehicles are equal, computes the impulsive maneuvers in the primary radial and cross-track directions, and parameterizes these maneuvers and obtain solutions that satisfy constraints, for example collision avoidance or direction of coverage, or optimize quantities, such as time or fuel usage. Apocentral coordinates and a set of four relative orbital parameters are used. The method separates the change in relative velocity (maneuvers) into radial and crosstrack components and uses a waypoint technique to plan the maneuvers.

Abstract: An autonomous system for a satellite which calculates collision paths of debris from anywhere within the spheroid around the satellite by using its radar/ladar data and from data on its own orbit derived by onboard sensors such as star, earth and sun sensors or from stored data sent from its ground control station through the satellite's command subsystem. If a collision would be likely, the system calculates the minimum change in the satellite's orbit to avoid such collision and generates and executes commands for firing on-board orbital control thrusters to put the satellite in a suitable avoidance orbit.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station-keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and angular momentum control using.

Abstract: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station-keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and control using Attitude Control Systems (ACS).

Abstract: A method of generating orbital transfers for spacecraft. The method provides an innovative technique for transferring spacecraft from one Earth orbit to another Earth orbit using significant solar gravitational influences. In one particular implementation, the multi-bodies in the transfer determination are the Earth (about which the spacecraft is to orbit) and the Sun (e.g., the Earth and the Sun are the first and second celestial bodies providing multi-body dynamics). The transfer orbit or trajectory is determined to make use of efficient tangential maneuvers by leveraging solar gravitational influences to improve transfer performance. Based on the generated transfer orbit, the spacecraft is controlled to perform one or more maneuvers to achieve a transfer orbit that traverses into a regime where the spacecraft's trajectory is significantly affected by gravity from both the Sun and the Earth. The spacecraft performs a near-tangential orbit insertion maneuver to enter the final orbit.

Abstract: A spacecraft (P) has a propulsion system making it possible to exert a force of variable magnitude and orientation on the spacecraft, a control system designed to control the propulsion system in terms of magnitude and orientation so as to make the spacecraft approach a target around a planet, with the aid of a force having at least one component (fx, fy, fz), in the rotating reference frame tied to the target, which depends substantially linearly on the corresponding coordinate (x, y, z) of the craft in this reference frame.

Abstract: A method of managing an eccentricity vector of a geosynchronous satellite orbit is provided. The method includes determining a desired target locus of an acquisition control of a satellite in a geosynchronous orbit, where the acquisition control ensures that an osculating trajectory of the satellite converges in mean to the target locus.

Abstract: A system for determining and controlling an orbit of a spacecraft may include a GPS receiver mounted to the spacecraft to receive GPS information from a plurality of GPS satellites visible to the GPS receiver for determining a position of the spacecraft relative to a reference coordinate system. The system may also include a filter configuration control module to sense a thruster maneuver operation to adjust the orbit of the spacecraft. The system may also include an orbit determination filter to determine at least one of a spacecraft position, velocity, and residue acceleration based on the GPS information and information from the filter configuration control module related to the thruster maneuver operation.

Abstract: A system is dedicated to the control of the deployment of at least two spacecraft (ES1, ES2) which are provided with maneuvering means (MD1, MD2) and are intended to move according to a chosen formation. This system includes a control device (MM11, MM12, MT, MC1) comprising i) first measurement means (MM11, MM12, MT) responsible for determining substantially simultaneously and with high precision the orbital positions of the spacecraft (ES1, ES2), and ii) first calculation means (MC1) responsible for determining for each of the spacecraft, as a function of their orbital positions, maneuvers intended to position each of them at a chosen instant substantially in a chosen position with respect to a reference trajectory, having regard to the time law of a reference craft (ES1) on this reference trajectory (TR), so as to place the formation in a chosen configuration.

Abstract: An improved approach to satellite-based navigation (e.g., GPS) is provided. In one embodiment, a method includes determining a nominal orbital path of a navigation satellite. The method also includes transmitting ephemeris data corresponding to the nominal orbital path from the navigation satellite to a plurality of navigation devices. The method further includes determining an actual orbital path of the navigation satellite locally at the navigation satellite. In addition, the method includes determining a deviation between the actual orbital path and the nominal orbital path locally at the navigation satellite. The method also includes autonomously adjusting the actual orbital path locally at the navigation satellite to reduce the deviation between the actual orbital path and the nominal orbital path.

Abstract: 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.

Abstract: An autonomous unmanned space flight system and planetary lander executes a discrete landing sequence including performing an initial velocity braking maneuver to remove velocity at altitude, coasting during which the planet surface is imaged and correlated to reference maps to estimate cross-track and along-track navigation errors and one or more lateral braking maneuvers are performed to reduce cross-track navigation error, and performing a terminal velocity braking maneuver(s) to reduce the along-track braking maneuver and remove the remainder of the velocity just prior to landing. A bi-propellant propulsion system provides a very high T/M ratio, at least 15:1 per nozzle. Short, high T/M divert maneuvers provide the capability to remove cross-track navigation error efficiently up to the maximum resolution of the reference maps.

Abstract: Systems and methods are disclosed employing electric propulsion stationkeeping in a cyclical manner to better match the cyclical pattern of power generated by the solar array system. For a typical orbit design, e.g. a geostationary orbit, North-South stationkeeping can be intermittently suspended, tolerating some additional drift but yielding in a very significant reduction in the required solar power system. If necessary, stationkeeping can be supplemented with a chemical thrusters during off periods for the electric propulsion. Because of this, the overall electrical power margin for the solar array system design can be reduced without compromising the mission performance.

Abstract: The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station -keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and control using Attitude Control Systems (ACS).

Abstract: Systems and methods for are adapted for automatic implementation of exclusion zone avoidance for target-tracking vehicles, such as spacecraft. The systems and methods are configured to monitor pointing commands (commanded attitude and angular rates) generated for target tracking, and modify these commands as necessary to avoid pointing a boresight into an exclusion zone.

Abstract: According to one embodiment of the invention, a stationkeeping method for a geostationary satellite includes determining a gravitational force of the sun on the satellite at a beginning of a stationkeeping operation and a gravitational force of the moon on the satellite at the beginning of the stationkeeping operation. An initial inclination vector of the satellite is determined at the beginning of the stationkeeping operation that accounts for a first set of one or more perturbations affecting the orbit of the satellite. A maneuver strategy is determined to correct for a second set of one or more perturbations affecting the orbit of the satellite without accounting for the first set of one or more perturbations affecting the orbit of the satellite. Finally, a maneuver is performed on the satellite according to the maneuver strategy.