Abstract: Spacecraft control modules can steer a spacecraft during a slew maneuver. The spacecraft control modules can receive a target attitude for the spacecraft slew maneuver. An angle between a current attitude and the target attitude may be determined. A smooth attitude offset profile with respect to a steering frame may be established from the current attitude to the target attitude. A smooth attitude command may be generated from the smooth attitude offset profile. The smooth attitude command may be applied to an attitude control module of the spacecraft.

Abstract: Spacecraft payload orientation steering is provided for an orbiting spacecraft in motion along an orbit track around a celestial body, the orbit track having a nominal inclination with respect to an equatorial orbit, a substantial eccentricity, and a drift angle with respect to the nominal inclination. Coordinates of an optimal payload target location as a function of a spacecraft position along the orbit track are determined, the target location being on the surface of the celestial body and having a substantial motion with respect to the surface and with respect to a spacecraft nadir. A payload of the spacecraft is substantially aligned with the determined coordinates by steering the satellite body to correct for at least one of the inclination drift angle, and the eccentricity, thereby adjusting the spacecraft orientation as a function of the spacecraft position along the orbit track.

Abstract: Provided are an attitude control system and method of a spacecraft of an artificial satellite that may enhance a maneuverability and a controllability by simultaneously applying a reaction wheel and a thruster among drive units used to maneuver an attitude of the spacecraft of the artificial satellite. The attitude control system may include: a thruster-based attitude controller which control firing time of thrusters mounted on the spacecraft; and a reaction wheel-based attitude controller controlling driving of a reaction wheel mounted on the spacecraft. The spacecraft may include a plurality of reaction wheels. When a defect occurs in the spacecraft due to a partial malfunction of the reaction wheels, an attitude maneuverability of the spacecraft may be corrected by simultaneously applying the thruster-based attitude controller and the reaction wheel-based attitude controller.

Abstract: In a method for injecting a plurality of spacecraft into different circum-earth or interplanetary orbits individually in a single launch, a plurality of spacecraft coupled to an assist module are injected into an interplanetary orbit having a periodicity synchronous with the earth's revolution period. Then, in a maneuver allowing the assist module to re-counter with and pass near to the earth (Earth swing-by), the assist module appropriately performs an orbital change maneuver and a separation maneuver for each of the spacecraft in a sequential order. Through these maneuvers, a closest-approach point to the earth is changed for each of the spacecraft so as to drastically change a subsequent orbital element for each of the spacecraft. The assist module takes a sufficient time to determine a target orbit for each of the spacecraft with a high degree of accuracy until a half month to several days before a closest-approach time in the Earth swing-by.

Abstract: The present invention relates to a method for maneuvering an imaging satellite, and more particularly a method for commanding control moment gyroscopes on an imaging satellite to change the attitude of the satellite. In one embodiment, a method for generating a gimbal rate trajectory for maneuvering a satellite to point to a target includes providing a final attitude for pointing to the target and a final satellite rate for imaging the target, providing an initial attitude and an initial set of gimbal angles, and determining a gimbal rate trajectory from the initial set of gimbal angles to a final set of gimbal angles.

Abstract: A system and method are provided for operating a control moment gyro (CMG) in a spacecraft. The CMG includes a spin motor, and a CMG rotor that is coupled to the spin motor and is configured to rotate about a spin axis. Electrical power from a power source is supplied to the spin motor to rotate the CMG rotor about the spin axis. A determination is made as to whether the power source can supply sufficient electrical power to the spin motor. If the power source cannot supply sufficient electrical power to the spin motor, the spin motor is rotated by the CMG rotor to generate and supply electrical power to the power source.

Abstract: A method for calculating a trajectory of a satellite following a maneuver includes determining a first estimated trajectory of a satellite using a first sequential mode of operation; performing a thruster burn as a correction maneuver for the satellite; and providing a real-time assessment of the correction maneuver by utilizing a second sequential mode of operation. The second sequential mode of operation includes receiving a data point from an uplink/downlink facility, determining a second estimated trajectory and an estimated thruster performance of the satellite based upon the received data point, and determining a trajectory error of the satellite based upon the received data point. The method further includes repeating the second sequential mode of operation if the trajectory error is above a specified threshold level, and repeating the first sequential mode of operation if the trajectory error is below the specified threshold level.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.

Abstract: A control system of a spacecraft for controlling two or more sets of collinear control moment gyroscopes (CMGs) comprises an attitude control system. The attitude control system is configured to receive a command to adjust an orientation of the spacecraft, determine an offset for a momentum disk for each of the two or more sets of CMGs that maximizes torque, determine a momentum needed from the two or more sets of CMGs to adjust the orientation of the spacecraft, and calculate a total torque needed by taking the derivative of the momentum. The control system further comprises a momentum actuator control processor coupled to the attitude control system, the momentum actuator control processor configured to calculate a required gimbal movement for each of the CMGs in each of the two or more sets of collinear CMGs from total torque.

Abstract: A vehicle control system and related sub-components together provide an operator with a plurality of specific modes of operation. The various modes of operation incorporate different levels of autonomous control. Through a control user interface, an operator can move between certain modes of control even after vehicle deployment. Specialized autopilot system components and methods are employed to ensure smooth transitions between control modes. Empowered by the multi-modal control system, an operator can manage multiple vehicles simultaneously.

Abstract: A method and apparatus for determining long term orbit (LTO) models using variable time-horizons to improve the orbit and clock model accuracy. The method and apparatus use either historic ephemeris or historic measurements for at least one satellite to produce an orbit parameter prediction model (an LTO model). The parameter predicted by the model is compared to an orbit parameter of a current broadcast ephemeris. The result of the comparison (an indicia of accuracy for the model) is used to establish a time-horizon for the orbit parameter prediction model for that particular satellite. Such a time-horizon may be established in this manner for each satellite within a satellite constellation.

Abstract: A visualization system for a computer system includes a modeling portion configured to specify the virtual geographic location in response to a three-dimensional model of the virtual geographic location and in response to the position of the viewer with respect to the virtual geographic location, a satellite data portion configured to specify a representation of satellite status data in response to the position of the viewer with respect to the virtual geographic location and in response to the satellite status data, and a three-dimensional output portion configured to provide at least two images of the virtual geographic location and the representation of the satellite status data to the viewer.

Abstract: A method of providing heading awareness on a flight display is provided. The method comprises displaying a heading awareness symbology to indicate heading relative to track, and varying the prominence of the heading awareness symbology on a display element based on at least one sensed condition.

Abstract: A routing tool is disclosed. In one embodiment, the method and system include receiving flight data and geographic data in the aircraft, and generating route data based on the flight data and the geographic data. The route data provides information about attainable landing areas for the aircraft.

Abstract: Techniques for providing singularity escape and avoidance are disclosed. In one embodiment, a method for providing control moment gyroscope (CMG) attitude control singularity escape includes calculating a Jacobian A of a set of control equations, calculating a measure of closeness to a singularity, and comparing the calculated closeness to a threshold value, when the calculated closeness is less than or equal to the threshold value, recalculating the Jacobian A. Recalculating may include determining a new direction of virtual misalignment of ? and ?, recalculating the Jacobian inputting the new direction of the virtual misalignment, recalculating the measure of closeness to a singularity, and comparing the measure of closeness to the threshold value. Further, the method may include calculating a gimbal rate command if the of closeness is greater than the threshold value and generating a torque from the gimbal rate command to control the attitude of a satellite.

Abstract: Methods and apparatus for assigning weighting coefficients to measurements of a succession of stars acquired by a star sensor connected to a client device in order to determine a spatial orientation, characterized in that higher or lower preference is given to refreshment of the positions of measurements with the highest weights and/or stars on which these measurements are made by the star sensor and/or its client device, so as to displace part of the power of the error associated with the set of star measurements within the frequency spectrum.

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: A mechanical flight control system for a rotary-wing aircraft is disclosed. The flight control system comprises an upstream portion, a downstream portion, and a booster means for connecting the upstream portion to the downstream portion. The booster means may comprise dual concentric valve actuators and/or a variety of system load limiting features.

Abstract: 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.

Abstract: A parking control surveillance system is provided for monitoring vehicles within parking spaces for providing parking enforcement personnel information concerning authorized and unauthorized vehicles within the parking space. The parking control surveillance system includes a parking stop forming a curb or having dimensions substantially the same as traditional cement parking stops located at the rear of a parking space. The parking stop includes a video recorder for recording the identity or for obtaining images of vehicles entering or exiting the designated parking space. The parking stop further includes a transmitter for transmitting images and information such as the time of parking transgression to vehicle enforcement personnel. Preferably, the vehicle identity and time of parking are transmitted to a central office for billing the vehicle owner.

Abstract: Provided is a method for controlling an attitude of a satellite using target track approximation. The method includes the steps of: a) receiving coordinate information of at least one of target areas; b) generating a target track by approximating a track of a satellite based on the received coordinate information; and c) detecting a location of the satellite on a current track, calculating an attitude angle of the satellite for the target track using the location of the satellite and the location of the target area, and applying the calculated attitude angle of the satellite to control the attitude of the satellite.

Abstract: A method of controlling the attitude of a satellite (1) including two gyrodynes (3,4) and a third main actuator (2) which delivers torques at least along the Z axis. The method includes: fixing the gimbal axes, A1 and A2, of the gyrodynes (3, 4) parallel to Z; setting a non-zero bias (?) between the angular momentum vectors (Formula I) of the gyrodynes; using the measurements provided by the sensors on board the satellite to estimate the kinematic and dynamic variables necessary in order to control the attitude of the satellite (1); calculating set variables in order to the objectives assigned to the satellite (1) attitude control system; and using the deviations between the estimated variables and the set variables to calculate control orders and to send same to the main actuators (2, 3, 4).

Abstract: A method for minimizing the amount of propellant required to be carried onboard the satellites of an Earth orbiting satellite constellation to maintain the minimum angle of elevation between a selected geographical area of the Earth and at least one satellite of the constellation during the life of the constellation includes establishing an optimal initial inclination and RAAN of each satellite of the constellation. The method enables the satellites of the constellation to carry additional hardware and other payload mass, or alternatively, the design life of the constellation to be increased.

Abstract: Techniques for providing singularity escape and avoidance are disclosed. In one embodiment, a method for providing control moment gyroscope (CMG) attitude control singularity escape includes calculating a Jacobian A of a set of control equations, calculating a measure of closeness to a singularity, and comparing the calculated closeness to a threshold value, when the calculated closeness is less than or equal to the threshold value, recalculating the Jacobian A. Recalculating may include determining a new direction of virtual misalignment of ? and ?, recalculating the Jacobian inputting the new direction of the virtual misalignment, recalculating the measure of closeness to a singularity, and comparing the measure of closeness to the threshold value. Further, the method may include calculating a gimbal rate command if the of closeness is greater than the threshold value and generating a torque from the gimbal rate command to control the attitude of a satellite.

Abstract: An airborne processor is linked to a ground-based, vehicle based, or satellite-based supplier of approaching aircraft data and/or terrain data, including airfield structures. The processor also has access to the aircraft's state, e.g. position, speed, intended route, aircraft dimensional data, or the like. Algorithms within the processor calculate a zone of protection about the aircraft in light of aircraft type, turning radius, and other identifying data. The processor combines the present and estimated future positions of the aircraft with the approaching aircraft and/or airfield structure data, and creates an alert to the crew if a threat of a ground incursion is detected. The display provides a visual of the zone of protection about the aircraft, and approaching aircraft and/or airfield structure to draw the attention of the crew to the threat.

Abstract: A satellite has a depletion detector arranged in a propellant line such that, when depletion is detected, the amount of propellant remaining in the propellant lines is sufficient to dispose of the satellite, and may include a margin sufficient for 6-12 months of stationkeeping. This provides a simple and reliable method of determining when decommissioning is required.

Abstract: Techniques for providing singularity escape and avoidance are disclosed. In one embodiment, a method for providing control moment gyroscope (CMG) attitude control singularity escape includes calculating a Jacobian A of a set of control equations, calculating a measure of closeness to a singularity, and comparing the calculated closeness to a threshold value, when the calculated closeness is less than or equal to the threshold value, recalculating the Jacobian A. Recalculating may include determining a new direction of virtual misalignment of ? and ?, recalculating the Jacobian inputting the new direction of the virtual misalignment, recalculating the measure of closeness to a singularity, and comparing the measure of closeness to the threshold value. Further, the method may include calculating a gimbal rate command if the of closeness is greater than the threshold value and generating a torque from the gimbal rate command to control the attitude of a satellite.

Abstract: Method for the computer-assisted determination of an optimum-fuel control of nozzles according to a control instruction b=Ax. A defined matrix transformation of starting constraints for the mass flow of the nozzles and of the minimization criterion thereby takes place in a computer-assisted manner, a data processing representation of a geometric description of the matrix-transformed starting constraints, a computer-assisted determination of limiting point sets of the geometric description of the starting constraints through a computer-assisted geometric search procedure in the vector space and the application of the matrix-transformed minimization criterion to the points of the limiting point sets.

Abstract: A method and apparatus are described for controlling the attitude of a vehicle in a space having at least two opposed viewable regions, each region being viewed by a respective first sensor for sensing a first frequency band of electromagnetic radiation and a respective second sensor for sensing a second frequency band of electromagnetic radiation, wherein respective first and second data sets from the first and second sensors for each regions are produced, these second data sets are subsequently modified and combined with the first data sets to form respective third data sets for each region. The attitude of the vehicle is then adjusted until the third data sets are substantially equal.

Abstract: A control device (D) for a spacecraft of a group of spacecraft intended to travel in a chosen formation comprises i) a set of at least three send/receive antennas (A1-A3) installed on at least three differently oriented faces of its spacecraft and adapted to send/receive radio-frequency signals, ii) first measuring means (M1) responsible for determining the power of the signals received by each of the antennas (A1-A3) and for delivering sets of powers each associated with one of the other spacecraft of the group, iii) storage means (BD) responsible for storing sets of cartographic data each representative of the normalized powers of the signals received by each of the antennas (A1-A3) as a function of chosen send directions, and iv) processor means (MT) responsible for comparing each set of powers delivered by the first measuring means (M1) to the stored sets of cartographic data in order to estimate each send direction of the signals sent by the other spacecraft of the group with respect to a system of axes

Abstract: A method of and apparatus for determining and controlling the inertial attitude of a spinning artificial satellite without using a suite of inertial gyroscopes. The method and apparatus operate by tracking three astronomical objects near the Earth's ecliptic pole and the satellite's and/or star tracker's spin axis and processing the track information.

Abstract: An in-vehicle device includes: a hard disk drive for storing data; a communication element for transmitting a message to an user; an input element for receiving an instruction from the user; a communication control element for obtaining altitude information; and an activation control element for controlling activation of the hard disk drive. The communication control element obtains the altitude information before the hard disk drive is activated when the in-vehicle device is activated for the first time since standby power supply stopped while the in-vehicle device was not activated. The communication control element transmits the message when the current position is equal to or higher than a predetermined altitude. The activation control element starts to read the data from the hard disk drive when the input element receives the instruction that the user permits the activation of the hard disk drive.

Abstract: Methods and systems for controlling multi-body vehicles with fuel slosh are provided. In a method for stabilizing a vehicle with fuel slosh, a controller for a multi-body assembly is formed by selecting vehicle configuration and velocity as states of a rigid body. Equations of motion are written for the multi-body assembly, and velocity states are determined that are unactuated. A decoupling transformation is defined from the unactuated velocity states. The equations of motion are transformed using the decoupling transformation such that unactuated states are decoupled from actuated states. A slosh state estimator is coupled to the controller such that it is in operative communication with an input to the controller and an output from the controller.

Abstract: Methods and apparatus are provided for reorienting control moment gyros (CMGs) to compensate for CMG failure or change in spacecraft (S/C) mass properties or mission. An improved CMG comprises a drive means for rotating the CMG around an axis not parallel to the CMG gimbal axis. Releasable clamps lock the CMG to the spacecraft except during CMG array reorientation. CMGs arrays are combined with attitude sensors, a command module, memory for storing data and programs, CMG drivers and sensors (preferably for each CMG axis), and a controller coupling these elements. The method comprises determining whether a CMG has failed or the S/C properties or mission changed, identifying the working CMGs of the array, determining a new array reorientation for improved spacecraft control, unlocking, reorienting and relocking the CMGs in the array and updating the S/C control parameters for the new array orientation.

Abstract: An apparatus and method provide flight operations integrated planning and analysis tools (FOIPAT) to perform integrated planning and analysis tasks for autonomous (unmanned) systems operations. The FOIPAT may be used for space launch and satellite early orbit analysis. The FOIPAT improves the efficiency and accuracy of space launch and satellite operations pre- and post-launch tasks performed by, for example, the US Air Force, the National Reconnaissance Office (NRO) and their support contractors. The FOIPAT provides greater functionality than existing software tools, reduces risk of errors, enhances quality control and data reliability, and reduces launch and flight operations planning cycle time. The FOIPAT may provide methodology for unmanned aerial vehicles (UAV) as well as terrestrial and seagoing systems.

Abstract: 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.

Abstract: Provided is a method and system for maneuvering a first spacecraft relative to a nearby second spacecraft that occupies a finite volume of space including obstacles in the vicinity of the second from a measured present relative position to a desired ending relative position. A trajectory is computed for the first spacecraft from the present relative position to the ending relative position using the A* algorithm. One or more thrusters of the first spacecraft are then fired. The first spacecraft's relative position and velocity are then measured and compared to the computed trajectory.

Abstract: Methods and apparatus are provided for reaction wheel (RW) assemblies for spacecraft. The apparatus comprises a M/G coupled to an inertia wheel and a controller coupled to the M/G that, in response to commands it receives, couples power to or from an M/G of another RW assembly over a shared transfer connection (VXFR), so that one M/G acts as a generator powering another M/G acting as a motor. The controller compares generator voltage VGEN to the BEMF of the motor on VXFR and reconfigures a multi-winding M/G or steps-up the generated voltage to maintain VGEN>VXFR to maximize direct energy transfer from one RW to the other as long as possible. When VGEN declines sufficiently, the controller can couple the motor to the spacecraft power bus and/or the generator to a power dump so as to continue to provide the commanded torque if needed. Operation is automatic.

Abstract: Disclosed is an attitude change control system that is designed to efficiently output a torque for attitude change of a space craft using CMGs and realizes a real time CMG driving rule. A CMG gimbal steering law 15 generates target profiles for setting angle and angular velocity for each gimbal by applying an anisotropic weighted gradient method based upon the necessary torque calculated by a feed back controller 13 and a feed forward controller 14 from the angle and angular velocity in the target direction from the attitude navigator 12 and the current angle and angular velocity of the space craft estimated by the attitude estimator 11 as well as the current condition of each gimbal from the CMG 40, thereby controlling the CMG 40 for changing the attitude of the space craft dynamics 50 to the target direction.

Abstract: An emulator, which is a computer program, for simulating 32 bit ARINC-429 data output by an Inertial Navigation System on board a P-3 aircraft. The value for each word output by the emulator is adjustable by user. The user adjusts each value by using the computer's mouse to engage slider controls located on a computer applications window. ARINC-429 data which the user adjusts includes aircraft latitude, longitude and altitude data as well as pitch and roll data for the aircraft, and time and distance to a destination by the aircraft.

Abstract: A method for controlling an actuator of a vehicle comprises providing a dynamic condition sensor generating a vehicle movement signal and a position sensor for generating a reported position. A processor is coupled to the inertial sensor and the position sensor and comprises an estimator, a position measurement predictor having a filter, a comparator and a control shaping block, said estimator generating a vehicle position based upon the dynamic condition sensor, said position measurement predictor generating an estimated position measurement in response to the reported vehicle position and a matched frequency response to the movement signal, said control shaping block generating an actuator control signal in response to a comparison of the estimated position measurement and the reported vehicle position.

Abstract: Propellant management systems and methods are provided for controlling the delivery of liquid propellants in a space launch vehicle utilizing multiple rockets. The propellant management systems and methods may be configured to enable substantial simultaneous depletion of liquid propellants in each of a plurality of active rockets during operation of various booster stages of the launch vehicle.

Abstract: A method and apparatus for determining the attitude of a satellite using crosslink information is disclosed. The method and apparatus integrates data from the crosslinks with data available from other sources including, for example, star sensors, inertial sensors, and earth limb sensors to derive an accurate estimate of the satellite attitude, even in harsh nuclear environments.

Abstract: A control moment gyroscope system for delivering a target torque to a spacecraft including a rotor assembly having a rotor and a motor to spin the rotor about a rotor axis. A gimbal assembly has a gimbal for supporting the rotor assembly and a gimbal torque motor to rotate the gimbal about a gimbal axis, which is normal to the rotor axis, to generate an output torque. A control system has a sensor for determining the output torque and a processor in communication with the rotor assembly, the gimbal assembly and the sensor. The processor requests the target torque and establishes a feedback control loop to generate a torque error signal based on the output torque for bringing the output torque within a predetermined range of the target torque.

Abstract: A hand-held electronic celestial object-locating device assists in identifying a celestial object or directing a user to a desired celestial object. The device is useful for locating or identifying any celestial object including stars, constellations, planets, comets, asteroids, artificial satellites, and deep sky objects to name a few. The device utilizes sensors for 3-axis magnetic field and 3-axis gravitational field detection. The device utilizes a processor and an electronic database to perform the required calculations. The device's database may be updated through access to the Internet through which the updates may be purchased.

Abstract: A method for avoiding singularities in the movement of CMGs in an array of CMGs in a spacecraft includes a first step where a maneuver command to rotate a spacecraft orientation is received. Then, the torque needed to rotate the spacecraft's orientation is determined. Then, the torque is integrated to determine a momentum path. The momentum path is decomposed into a sequence of straight line segments. For each line segment, a unit vector along the straight line segments is determined. Then, it is determined if there is a continuous path connecting a start point and an end point of the line segment in a plane perpendicular to the unit vector. For each point along the path in the plane perpendicular to the unit vector, a set of gimbal angles is determined.

Abstract: A simulation system for simulating a GPS (global positioning system) satellite is configured for acquiring a navigation data in real time from the GPS satellite, generating a simulation signal from the navigation data while carrying out a clock time synchronization with an output signal from the GPS satellite, and outputting the signal to a GPS receiver as a target of evaluation.

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.

Abstract: A ship includes a star tracker mounted on a platform stabilized in ENU by the inertial navigation system (INS). The line-of-sight (LOS) of the star tracker is directed toward two separated stars, and the LOS difference angles are noted. The angles are processed to generate vector triads representing geodetic (ephemeris) and navigation system attitude. The triads are processed to generate a coordinate transformation matrix. The transformation matrix is separated into systematic error and reference attitude error. The reference attitude error is summed with the inertial navigation system attitude to generate corrected ENU attitude. The corrected attitude is used as a reference for shipboard sensors, to reduce errors when the sensor data is linked to other platforms.

Abstract: In one embodiment of the disclosure, a computing system includes a processor and a computer readable medium. The computer readable medium is operable to, when executed on the processor, determine an estimated trajectory of a satellite using a sequential mode of operation, perform a thruster burn, receive a data point from an uplink/downlink facility, determine an estimated trajectory and an estimated thruster performance of the satellite based upon the received data point. If the trajectory error is above a specified threshold level, repeat receiving an updated data point from the uplink/downlink facility, determining an updated estimated trajectory and an updated estimated thruster performance of the satellite based upon the updated data point, and determining an updated trajectory error based upon the updated data point.