Abstract: Computer systems (100) for satellite image acquisition and mission-plan calculation devices (120) for satellite and also Earth observation satellites (110). Based on the observation that the uncertainty connected with meteorological conditions can lead to the development of mission plans, for satellites, with low effectiveness, the uncertainty is characterized and used to improve the preparation of mission plans for satellite. A model of the uncertainties is built upon observations and/or past meteorological forecasts. Next, at the time of planning, the model is used by presenting the observations and/or the current forecast as input to the model. In that way, the acquisition zones for which it is unlikely to get acceptable images on the preferred acquisition date may be identified.

Abstract: A constellation of many satellites provide communication between devices such as user terminals (UTs) and ground stations that are connected to other networks, such as the Internet. A constellation management system (CMS) facilitates management and operation of the satellites in the constellation and facilitates information exchange with other authorized systems to provide for situationally aware operation. The CMS may ingest data such as satellite telemetry, space weather data, object ephemeris data about other orbital objects, and so forth. The CMS uses the ingested data to automatically operate satellites to perform routine activities such as station keeping maneuvers, maintenance activities, interference mitigation, and so forth. Confirmation from a human operator may be obtained before performing some activities. Activities may be planned and coordinated to minimize resource consumption for the individual satellite as well as the constellation.

Abstract: An LADCP and combined inertial navigation system combined observation system and method. The system includes an LADCP, combined inertial navigation system host, data processing unit, wireless transmission module, voltage-resistant wireless antenna, battery cabin, first and second GNSS antennae and an instrument support. The data processing unit is connected to the LADCP and the system host through serial ports to obtain ocean current profile measurement data, a GNSS position, GNSS orientation data, three-axis acceleration data, a three-axis gyroscope angular velocity, a roll angle, a pitch angle and a course angular rate. Real-time high-precision correction is performed on a three-dimensional flow velocity of an ocean current profile observed by the LADCP based on the orientation data, roll angle, pitch angle and the course angular rate. Accordingly, measurement precision of flow velocity and flow direction by the LADCP is improved, and working costs to carry out profile flow observation are reduced.

Abstract: In a satellite side, there are provided a satellite-side operation parameter storage and an operation planning unit; and on a ground side, there are provided a ground-side operation parameter setting unit, a ground-side operation parameter storage to store operation parameters the same as stored in the operation parameter storage of the satellite-side, and an operation plan estimation unit having the same function as that of the operation planning unit of the satellite side. The ground-side operation parameter setting unit determines operation parameters to be used for the operation plan; and when necessary, the ground-side operation parameter setting unit saves them to the ground-side operation parameter storage and uploads them to the satellite-side operation parameter storage. The ground-side operation plan estimation unit and the satellite-side operation planning unit each produce operation plans according to the same operation parameters.

Abstract: A signal receiver method to achieve satellite position fix by improving satellite orbit prediction includes: acquiring satellite signals and navigation data and calculating a position solution, which includes predicting the state or orbit of one or more satellites. The prediction includes using a model of the solar radiation pressure operating on a selected satellite.

Abstract: A constellation of Earth-orbiting spacecraft, the constellation having an orbital maneuver lifetime life (OML), includes a first spacecraft disposed in a first orbit and a second spacecraft disposed in a second orbit, each of orbit being substantially circular with a radius of approximately 42,164 km and having a respective inclination with respect to the equator specified within a range of 10° to 20°. The first orbit has, at beginning of life (BOL), a first right ascension of ascending node (BOL-RAAN1) and the second orbit has, at BOL, a second RAAN (BOL-RAAN2) the BOL-RAAN1 and the BOL-RAAN2 being separated by a first angular separation ?-RAAN1. A first stationkeeping delta-V (?V1) applied over the OML to the first spacecraft, in combination with a second delta-V (?V2) applied over the OML to the second spacecraft, maintains the ?-RAAN1 approximately constant and an actual inclination within specification, and ?V1 approximately equals ?V2.

Abstract: 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.

Abstract: A method of determining the appropriateness of satellite orbit modeling is provided. The method includes calculating values of parameters, that the predetermined model has, on the basis of predicted position data including a first predicted position at a first point of time and a second predicted position at a second point of time of a positioning satellite in time series, calculating a first and a second calculated positions of the positioning satellite derived from the predetermined model by using the values of the parameters; and determining the appropriateness of the predetermined model using the values of the parameters, on the basis of first difference between the first predicted position and the first calculated position, and seconded difference between the second predicted position and the second calculated position. The predetermined model is used when approximating a satellite orbit of the poisoning satellite.

Abstract: A navigation apparatus is capable of image-capturing and is used to implement a navigation method including the steps of: a) obtaining current location information associated with the navigation apparatus, and location information of a reference target along a planned route that is being traversed; b) capturing real-time navigation images of the planned route according to an image-capture configuration parameter; c) obtaining a display screen position associated with the reference target with reference to the current location information, the target location information, and the image-capture configuration parameter; and d) showing on the display device the navigation images together with a mark corresponding in position to the display screen position.

Abstract: An apparatus and method for determining an orbit of a geostationary satellite is provided. The orbit of the geostationary satellite may be determined using at least one pseudo-range of the geostationary satellite calculated based on an orbit and a position of at least one global positioning system (GPS) satellite, position information of the geostationary satellite, and an angle between the geostationary satellite and each GPS satellite.

Abstract: A method of enhancing on-board state estimation for a spacecraft utilizes a network of assets to include planetary-based assets and space-based assets. Communication signals transmitted from each of the assets into space are defined by a common protocol. Data is embedded in each communication signal transmitted by the assets. The data includes a time-of-transmission for a corresponding one of the communication signals and a position of a corresponding one of the assets at the time-of-transmission. A spacecraft is equipped to receive the communication signals, has a clock synchronized to the space-wide time reference frame, and has a processor programmed to generate state estimates of the spacecraft. Using its processor, the spacecraft determines a one-dimensional range from itself to at least one of the assets and then updates its state estimates using each one-dimensional range.

Abstract: The present invention provides methods for positioning at least one component, in particular a seat, in or on an aircraft or spacecraft comprising the following steps: controlling at least one light source of the aircraft or spacecraft to display at least one desired position of at least one component; and positioning the at least one component in the at least one desired position displayed.

Abstract: A method of determining the reliability of long-term predicted orbit data, includes: determining the reliability of long-term predicted orbit data, which is acquired by predicting a satellite orbit in a target period of at least one day, using predicted position data including predicted positions of a positioning satellite in time series and actual position data including actual positions of the positioning satellite corresponding to the predicted positions.

Abstract: An amount of propellant remaining in an orbiting satellite can be estimated in a more accurate manner than is possible with conventional approached. Pressure and temperature telemetry data received from the satellite can be analyzed using a maximum likelihood estimation approach that reconstructs a predicted tank pressure signal using the temperature data and determines a pressurant volume necessary to make the reconstructed pressure signal match the received pressure signal. Drift of the pressure data received from pressure transducers in the satellite can also be addressed using the current subject matter, as can issues including but not limited to inaccessible propellant due to satellite spin, tank expansion under pressure, and the like. Independent determinations of the amounts of propellant remaining can be made using moment of inertia calculations in situations in which an axis of spin of the satellite is known a priori.

Abstract: A geospatial object property assessment apparatus comprises a processing resource (102) arranged to support a geospatial object property processor (206) comprising a subject data source input (212) capable of receiving a first metadata associated with 10 quality of current source information content used to define a property (408) of a geospatial object; a resource data source input (214) capable of receiving a second metadata associated with quality of candidate resource information content for updating the definition of the property (408) of the geospatial object.

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 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 computing apparatus and method of manipulating a displayed orbit of an object by detecting a gesture having a correspondence between body movements as the gesture and a resulting change to one or more orbital parameters within a displayed orbit system, changing an orbital parameter of the displayed orbit system according to detection of the gesture, and changing visualization on the computer display screen of the orbit system according to the change to the orbital parameter.

Abstract: A first position of a satellite is calculated at a first time in dependence on received orbit data corresponding to an orbit path of the satellite. An orbit path of the satellite is modeled from the first position at the first time to a second time to determine a second position of the satellite at the second time. A third position of the satellite is then calculated at the second time in dependence on the received orbit data. The second position and third position are compared to determine a validity of the orbit data.

Abstract: An improved approach to satellite-based navigation (e.g., GPS) is provided. In one embodiment, a method includes receiving a first set of tracking information. A nominal orbital path for the navigation satellite is determined using the first set of tracking information. Ephemeris data corresponding to the nominal orbital path is computed and uploaded to the navigation satellite. Long-term navigation information corresponding to the nominal orbital path is transmitted to a communication system for broadcast to a plurality of navigation devices. A second set of tracking information is received, an orbital path of the navigation satellite using the second set of tracking information is predicted, and a difference between the predicted orbital path and the nominal orbital path is determined. Commands configured to instruct the navigation satellite to adjust an actual orbital path of the navigation satellite to substantially conform to the nominal orbital path are uploaded to the navigation satellite.

Abstract: A motion of an object is controlled from a geostationary transit orbit (GTO) of an earth to an orbit of a moon. A first trajectory of the motion of the object is determined from an intermediate orbit of an earth to a neighborhood of a stable manifold of a first Lagrange point (L1). A second trajectory of the motion of the object is determined from the GTO to the intermediate orbit. A third trajectory of the motion of the object is determined from the neighborhood to the stable manifold to an L1 orbit, and a fourth trajectory of the motion of the object is determined from the L1 orbit to the orbit of the moon. A trajectory from the GTO to the orbit of the moon is determined based on a combination of the first, the second, the third, and the fourth trajectories.

Abstract: The invention relates to a navigation systems and elements. A network element (M) includes a receiver (M.2.2) for forming assistance data relating to at least one navigation system. The network element (M) inserts indication of the navigation system and a selected mode into the assistance data and constructs the assistance data according to the selected mode. The network element (M) has a transmitting element (M.3.1) for transmitting the assistance data via a communications network (P) to a device (R). The device (R) includes a positioning receiver (R.3) for performing positioning on the basis of one or more signals of the at least one satellite navigation system; a receiver (R.2.2) for receiving the assistance data from the network element (M); and an examining element (R.1.1) adapted to examine the received assistance data. The assistance data is adapted to be used by the positioning receiver for performing positioning of the device (R).

Abstract: A flight dynamics subsystem (FDS), a velocity increment calculation module, and operational methods of the same are provided. A used fuel quantity actually used in a satellite is calculated, and a velocity increment is calculated using the calculated fuel quantity. Therefore, an orbit of the satellite may be estimated more accurately.

Abstract: A method and a device are provided for the optimization of the mass of a satellite. The method includes: a step of calculation of an elliptical second orbit obtained by rotation of a first orbit about an axis connecting the periapsis and the apoapsis, the elliptical second orbit being associated with a second maximum eclipse duration less than a first maximum eclipse duration; a step of determination of a maneuver enabling the satellite to move to the second orbit; and a step of calculation of a second battery mass making it possible to maintain the satellite in operation during the second maximum eclipse duration and of calculation of a mass of fuel necessary to effect the maneuver.

Abstract: Certain embodiments of the invention may include systems, methods, and apparatus for sensing flight direction of a spacecraft. According to an example embodiment of the invention, a method is provided for determining flight direction of a spacecraft. The method includes providing at least one imaging detector associated with a spacecraft; imaging at least a portion of a celestial body onto the at least one imaging detector; acquiring, by the at least one imaging detector, sequential images of at least a portion of the celestial body; and determining the spacecraft flight direction relative to the celestial body based at least in part on processing the sequential images, wherein the processing is performed by one or more computer processors.

Abstract: Systems and methods of controlling solid propellant gas pressure and vehicle thrust are provided. Propellant gas pressure and a vehicle inertial characteristic are sensed. Propellant gas pressure commands and vehicle thrust commands are generated. A propellant gas pressure error is determined based on the propellant gas pressure commands and the sensed propellant gas pressure, and vehicle thrust error is determined based on the vehicle thrust commands and the sensed vehicle inertial characteristic. Reaction control valves are moved between closed and full-open positions based on the determined propellant gas pressure error and on the determined vehicle thrust error. The system and method allow the reaction control valves to operate at variable frequencies or at fixed frequencies. The system and method also allows propellant pressure to be commanded to follow a predetermined pressure profile or commanded to vary “on-the-fly.

Abstract: A protective device for an electronic unit on a space exploration vehicle. A Laplace transform calculation unit generates a Laplace transform of an electronic unit state. A system parameter identification unit identifies system parameters based on the Laplace transform of electronic unit states. A Fourier transform calculation unit generates a ratio of the ground contacting mechanism state to a highest dominant frequency of the measured electronic unit state. A critical travel speed calculation unit generates a critical travel speed based on the identified system parameters. A calculation unit for threshold of travel speed generates a threshold of a travel speed based on the critical travel speed. An autonomous motion controller generates a control signal that drives the space exploration vehicle based on the threshold of the travel speed.

Abstract: According to one embodiment of the disclosure, a method for use in telemetry processing includes receiving telemetry data originating from a satellite, such that the telemetry data comprises a plurality of data segments. The method includes processing the plurality of data segments simultaneously and transmitting a signal to the satellite, in response to the processing, for effecting a change in the direction of the satellite.

Abstract: The invention set forth herein describes propulsive guidance methods and apparatus for controlling and shaping an atmospheric skip reentry trajectory for a space vehicle. Embodiments of the invention may utilize a powered explicit guidance algorithm to provide a closed-loop control method for controlling a space vehicle during a skip reentry maneuver.

Abstract: A method for transmitting additional information in a satellite navigation system includes providing a navigation message having a plurality of parameters, selecting at least one parameter from the plurality of parameters for the transmitting of the additional information, replacing the at least one parameter, at least partially, by the additional information so as to form a changed navigation message, and sending the changed navigation message.

Abstract: Provided are an apparatus and method for executing a telecommand on a geostationary satellite, and an apparatus and method for verifying a telecommand execution status on a geostationary satellite ground control system. When a telecommand on a satellite is executed, the satellite generates command execution verification words for the executed telecommand and adds the generated command execution verification words to a telemetry transfer frame, and thus a satellite ground control system can easily verify the telecommand execution.

Abstract: The present invention relates to image satellite planning, and more particularly to a method for allowing a deeper search for high value targets in a time-limited planning environment. In an exemplary embodiment, a method of computing an ordered subset of targets includes using an approximation for the time needed for the satellite to re-orient to a new target, rather than calculating each maneuver time between targets. By approximating the maneuver time rather than calculating it, the calculation time is reduced. Each iteration through the traveling salesman problem takes less time, and more iterations can be accomplished between imaging windows. The iterative process can search deeper into the traveling salesman problem to find a better solution.

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: A method and system for a multi-spectrum celestial navigation system includes a first sensor responsive to at least a first and a second wavelength band of electromagnetic radiation. The sensor is configured to generate a first output related to the first wavelength band of electromagnetic radiation and to generate a second output related to the second wavelength band of electromagnetic radiation. The system also includes a processor programmed to receive the first and second outputs, determine a position of the sensor with respect to one or more stars using a stored star catalog and the received first and second outputs, and output the determined position.

Abstract: The particle swarm-based micro air launch vehicle trajectory optimization method is carried out by formulating a parameter optimization problem, which is solved using a particle swarm optimization procedure. The optimization problem is formulated using a single objective function having the explicit objective to maximize the payload mass. Constraints on terminal conditions are imposed.

Abstract: Disclosed is a method and device, which provide an enhanced ability to monitor the navigation of an aircraft during a phase of flight in which the aircraft is close to the ground in which the aircraft uses positional information supplied by a satellite positioning system for the navigation. A display screen is used to display a first characteristic sign representing a selected setpoint value for a height parameter of the aircraft. The display screen also displays a second characteristic sign representing a current auxiliary value expressed in the form of an achievable height parameter. An alert is emitted when the second characteristic sign is determined to be within a predetermined height value of the first characteristic sign, and the alert is shown in visual format on the display screen. An alarm is also emitted following a predetermined time after the alert is emitted, if the setpoint value is not replaced by a new setpoint value.

Abstract: A Parametric Systematic Error Correction (ParSEC) method is disclosed which provides improved accuracy for image navigation and registration (INR). This method may be employed in imaging systems that exhibit systematic distortion. The ParSEC method comprises employing a Parametric Systematic Error Correction (ParSEC) system wherein coefficients for the system are determined by employing an nth order Taylor Series.

Abstract: A method for controlling a vehicle may include sensing a position of each of a plurality of stars relative to the vehicle. The method may also include determining an attitude of the vehicle using the sensed positions of the plurality of stars, and the attitude may be determined either with or without using information from a gyro or sensor for measuring angular velocity. The method may additionally include implementing a set of strategies to optimize determination of the attitude of the vehicle when using only the sensed positions of the plurality of stars, without information from the sensor for measuring angular velocity. The method may further include controlling the vehicle based on the determined attitude of the vehicle.

Abstract: An improved network enabled extended ephemeris navigation system includes a network server able to collect ephemeris, clock, and almanac information from orbiting GPS satellites, and to use that information to build up extended ephemeris predictions that will be valid and useful for at least a week. A mobile client is able to request and use the extended ephemeris predictions to search for and track orbiting GPS satellites visible to it. The improvement is characterized by a satellite position and clock compact model construction and database unit that constructs a compact short-term satellite model to be sent first in response to a request for extended ephemeris predictions from the mobile client, and that constructs several consecutive long-term satellite models each representing a unique portion of a day in at least a seven day series.

Abstract: A flight dynamics subsystem (FDS), a velocity increment calculation module, and operational methods of the same are provided. A used fuel quantity actually used in a satellite is calculated, and a velocity increment is calculated using the calculated fuel quantity. Therefore, an orbit of the satellite may be estimated more accurately.

Abstract: A first position of a satellite is calculated at a first time in dependence on received orbit data corresponding to an orbit path of the satellite. Anan orbit path of the satellite is modeled from the first position at the first time to a second time to determine a second position of the satellite at the second time. A third position of the satellite is then calculated at the second time in dependence on the received orbit data. The second position and third position are compared to determine a validity of the orbit data.

Abstract: Improved orbit/covariance estimation and analysis (OCEAN) system and method are presented utilizing ground station observations collected from satellites passing overhead to estimate the positions, velocities, and other parameters of multiple satellites using weighted least squares (WLS) batch and/or Kalman filter smoothing (KFS) estimation algorithms to estimate each parameter, with or without a priori knowledge of the errors involved with each observed parameter.