Abstract: A plurality of sensors provide spacecraft attitude signals to reaction wheel assemblies which provide converted signals to a flight controller that controls the reaction wheel assemblies. The reaction wheel assemblies include a secondary power supply to power the sensors, and power can be regenerated from the reaction wheels. A microcomputer in the reaction wheel assemblies can control the reaction wheel assemblies in place of the flight controller.

Abstract: Disclosed are systems and methods of determining or estimating an orientation of a spacecraft. An exemplary system generates telemetry data, including star observations, in a satellite. A ground station processes the telemetry data with data from a star catalog, to generate display data which, in this example, includes observed stars overlaid with catalog stars. An operator views the display and generates an operator input signal using a mouse device, to pair up observed and catalog stars. Circuitry in the ground station then processes two pairs of observed and catalog stars, to determine an orientation of the spacecraft.

Abstract: A method, apparatus, article of manufacture, and a memory structure for compensating for optical sensor data corrupted by angular acceleration is disclosed. The method comprises the steps of determining an angular acceleration of the optical sensor and modifying the optical sensor data according to the determined angular acceleration of the optical sensor. The apparatus comprises a sensor for determining an angular acceleration of the optical sensor and a navigation system for modifying the optical sensor data according to the determined angular acceleration of the optical sensor.

Abstract: A reconfigurable digital processing system for space includes the utilization of field programmable gate arrays utilizing a hardware centric approach to reconfigure software processors in a space vehicle through the reprogramming of multiple FPGAs such that one obtains a power/performance characteristic for signal processing tasks that cannot be achieved simply through the use of off-the-shelf processors. In one embodiment, for damaged or otherwise inoperable signal processors located on a spacecraft, the remaining processors which are undamaged can be reconfigured through changing the machine language and binary to the field programmable gate arrays to change the core processor while at the same time maintaining undamaged components so that the signal processing functions can be restored utilizing a RAM-based FPGA as a signal processor.

Abstract: Attitude acquisition methods and systems are provided which reduce the time generally required to acquire spacecraft attitude estimates and enhance the probability of realizing such estimates. The methods and systems receive, over a time span &Dgr;t, successive frames of star-sensor signals that correspond to successive stellar fields-of-view, estimate spacecraft rotation &Dgr;r throughout at least a portion of the time span &Dgr;t, and, in response to the spacecraft rotation &Dgr;r, process the star-sensor signals into a processed set of star-sensor signals that denote star positions across an expanded field-of-view that exceeds any of the successive fields-of-view. The expanded field-of-view facilitates identification of the stars that generated the processed set of star-sensor signals to thereby acquire an initial attitude estimate.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method comprises the steps of predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage, computing a magnitude and a direction of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque, and storing the computed magnitude and direction of momentum in at least one of the momentum wheels before deploying the appendage.

Abstract: Method and apparatus for using long term satellite tracking data in a remote receiver is described. In one example, long term satellite tracking data is received at a remote receiver from a server. The long term satellite tracking data is used to compute acquisition assistance data in the remote receiver. The remote receiver then uses the acquisition assistance data to acquire satellite signals. The acquired satellite signals may be used to locate position of the remote receiver.

Abstract: The wind turbulence prediction system according to the present invention provides a measurement system, which is capable of measuring three-dimensional wind turbulence, exhibits limited aerodynamic and structural effects when mounted on an aircraft, is capable of measurement even in cases in which the velocity is equal to or less than 20 to 30 m/s, at which velocity a Pitot tube is incapable of measurement, and even in cases where the air flow direction differs greatly from the axis of aircraft body, and does not produce an error resulting from disruption of the air flow by the aircraft body itself. The wind turbulence prediction system of the present invention adopts a system that measures the speed of remote three-dimensional air flow by mounting a laser wind speed indicator utilizing the Doppler effect on an aircraft, irradiating laser light while scanning same in a cone shape, and then receiving scattered light from wind turbulence regions forward of an aircraft body in flight.

Abstract: A fuel efficient technique for changing the inclination, with respect to the Earth's equator, for a satellite includes first maneuvering the satellite to the moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so-called fuzzy boundary or weak stability boundary. A negligibly small maneuver can then bring it back to the Earth on a reverse BCT to the desired Earth inclination. Another maneuver puts it into the new ellipse at the earth. In the case of satellites launched from Vandenberg AFB into LEO in a circular orbit of an altitude of 700 km with an inclination of 34°, approximately 6 km/s is required to change the inclination to 90°. The previous flight time associated with this method was approximately 170 days. A modification of this method also achieves a significant savings and unexpected benefits in energy as measured by Delta-V, where the flight time is also substantially reduced to 88 or even 6 days.

Abstract: A new steering logic based on a mixed weighted two-norm and least squares optimization solution overcomes a deficiency of being trapped in the momentum saturation singularities. The new steering logic utilizes an additional weighting matrix, W, in conjunction with deterministic dither signals in a weighting matrix V. The new steering logic also provides a simple means of avoiding troublesome, internal elliptic singularities which are commonly encountered by most other pseudoinverse-based steering logic. Various CMG systems, such as a typical pyramid array of four single-gimbal CMGs, two and three parallel single-gimbal CMG configurations, and four parallel double-gimbal CMGs of the International Space Station are described as illustrative embodiments demonstrating the simplicity and effectiveness of the new steering logic.

Abstract: An integrated inertial stellar attitude sensor for an aerospace vehicle includes a star camera system, a gyroscope system, a controller system for synchronously integrating an output of said star camera system and an output of said gyroscope system into a stream of data, and a flight computer responsive to said stream of data for determining from the star camera system output and the gyroscope system output the attitude of the aerospace vehicle.

Abstract: Spacecraft network and communication method thereof. A spacecraft network includes a first server spacecraft disposed in a first server orbit, a first client spacecraft disposed in a first client orbit, and a wireless local area network formed between at least the first server spacecraft and the first client spacecraft. The wireless local area network includes at least one communication channel to transmit and receive spatial information, at least one receiver to receive a first communication signal including at least routing information, at least one routing system to determine a desired route, and one transmitter to transmit the first communication signal. The first client spacecraft is free from the at least one routing system, and the first server spacecraft includes one of the at least one routing system.

Abstract: A satellite command system and method that provides satellite commanding using remotely-controlled modulation of telemetry parameters on-board a satellite. Software is provided on each of the redundant processor units that creates separate DMA command input pathways to the redundant processor units. A command translator is used at a ground station to translate a command that is to be implemented on the satellite into a form ready for modulation. The command information is uplinked to the satellite by remotely manipulating (modulating) a quantity that can be measured by the on-board computer using its telemetry collection functions. The software implemented in processor units recognizes the modulation and processes it to reconstruct the command. The reconstructed command is passed to the processor unit, which then executes the command.

Abstract: A system, a method, and a computer program product for determining tactile cueing of a flight control input apparatus includes an interface for receiving observed parameters relating to the flight envelope of an aircraft. Upstream processing elements receive the observed states, convert dimensional aircraft state parameters into nondimensional aircraft state parameters, and provide the data to neural networks. The neural networks receive combinations of observed parameters and nondimensional aircraft state parameters and predict flight envelope limiting parameters. The neural networks are trained to predict the limiting parameters. A downstream processing element determines the most limiting flight envelope limiting parameter provided from the neural networks, therefore providing a tactile cueing position for a flight control input apparatus.

Abstract: A method for correcting spacecraft pointing errors. An orbit frequency distortion angle between a rate sensor and at least one attitude sensor is estimated. An estimated attitude of the spacecraft is adjusted to correct for the distortion angle.

Abstract: An attitude detection system for an artificial satellite includes an interpolator for interpolating an angular-velocity signal to generate an interpolated angular-velocity signal, a sequential Kalman filter for generating a low-frequency attitude-angle signal, and an adder for adding the low-frequency attitude-angle signal and a high-frequency attitude-angle signal generated by a high-frequency angular sensor to generate a broad-band attitude-angle signal.

Abstract: A deployment method maintains a cluster of subsatellites within a rigid formation relative to and orbiting about a real or fictitious center satellite that in turn orbits the earth in a frozen inclined eccentric reference orbit by determining the subsatellite reference orbits of the each of the subsatellites based upon position and velocity vectors of the center satellite. The determined subsatellite reference orbits can then be maintained through fuel efficient microthrusting so as to maintain the rigid formation for coherently combining communication signals within a very large sparse antenna array formed by the cluster having improved beam directionality for improved signal communications.

Abstract: The present invention relates to a method for calculating the instantaneous characteristics of a satellite in orbit equipped with a GNSS receiver, capable of performing measurements of pseudo-distance, when fewer than four transmitters are simultaneously in visibility.

Abstract: A system and method for detecting if a UAV is operating within its flight envelope. The invention includes defining the flight envelope of the UAV. The operating state of the UAV is identified from a present time to a future elapsed time. The invention determines if the UAV will be within its flight envelope at the end of the elapsed time in the future. It can then command the UAV to maintain operation within the flight envelope if the operating state determined is outside the flight envelope.

Abstract: A method and apparatus for estimating the orbits of spacecraft or other objects relative to primary celestial bodies, given applicable measurements. The present invention uses the variation of parameters in universal variables for spacecraft trajectory propagation during operation of its optimal sequential orbit estimation. This method provides a universal capability for all three conics: elliptic, parabolic, and hyperbolic, forward-time and backward-time numerical integration for filtering and smoothing, respectively. A method for determining local and global estimates of atmospheric density is also provided. A method for realistic orbit simulation provides a means for validating an orbit estimate.

Abstract: Method and apparatus for creating a compact orbit model is described. Satellite tracking data is obtained having a first set of orbit terms that define a first orbit model. The satellite tracking data is formatted to form formatted data having a second set of orbit terms that define a second orbit model. A number of terms in the first set of orbit terms is greater than a number of terms in the second set of orbit terms.

Abstract: A stratospheric flying object which is placed in the stratosphere of the earth has an airship-shaped flying object body, a despun platform unit mounted on the flying object body by a mount, and a controller for detecting an azimuthal angle of the despun platform unit in a horizontal plane and controlling the platform unit to face in a constant direction at all times within the horizontal plane. Preferably, directional antennas and telescopes for optical communications are mounted on the despun platform unit.

Abstract: A spacecraft emulation system that can emulate both the attitude control subsystem and the non-attitude control subsystem is integrated into a single compact unit. The unit includes an emulated spacecraft control processor for processing attitude control information and an emulated central command and telemetry unit for interfacing simulated spacecraft data. The inlet also includes a first simulation engine that is operative to simulate the spacecraft attitude control system and a second simulation engine that is operative to simulate the spacecraft power, thermal, propulsion and payload subsystems. Both the first and second simulation engines are connected to the emulated spacecraft control processor via a respective bus. A host computer provides the command data and receives the telemetry data from the emulated spacecraft control processor.

Abstract: Methods and structures are provided that enhance attitude control during gyroscope substitutions by insuring that a spacecraft's attitude control system does not drive its absolute-attitude sensors out of their capture ranges. In a method embodiment, an operational process-noise covariance Q of a Kalman filter is temporarily replaced with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. The error effect of the substituted gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range. In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.

Abstract: A plurality of sensors provide spacecraft attitude signals to reaction wheel assemblies which provide converted signals to a flight controller that controls the reaction wheel assemblies. The reaction wheel assemblies include a secondary power supply to power the sensors, and power can be regenerated from the reaction wheels. A microcomputer in the reaction wheel assemblies can control the reaction wheel assemblies in place of the flight controller.

Abstract: Methods and structures are provided that enhance the accuracy of the service attitude of an inclined-orbit spacecraft and, thereby, facilitate reduction of service error between a communication service area and the spacecraft's payload beam. The enhancement is realized by configuring a beacon-receiving antenna to have a beacon-receiving field-of-view that substantially matches a beacon-station window. Preferably, the beacon-receiving field-of-view is elongated and tilted to enhance its match with the beacon-station window in both size and orientation. The goals are also realized by configuring the beacon-receiving antenna to have a beacon-receiving field-of-view that is substantially smaller than the beacon-station window and successively steering a beacon-receiving boresight to successive beacon-receiving attitudes that maintain the beacon station within the beacon-receiving field-of-view over each solar day.

Abstract: A spacecraft embedded in a reference frame rotating relative to inertial space. The spacecraft generally includes actuators for maneuvering the spacecraft with respect to the reference frame, an attitude measurement device that measures the pitch and roll attitude of the spacecraft with respect to the reference frame, a control device adapted to keep the roll and pitch angles of the spacecraft close to the commanded roll and pitch angles, and a harmonic torque estimator adapted to read the commanded angular velocity of the spacecraft relative to an inertial frame, read momentum wheel speeds, read known predicted external torques and combine angular velocity, measured wheel speed and known external torque to produce an estimated observable periodic torque.

Abstract: A load relief system for a launch vehicle including a wind sensing system responsive to wind speed and direction at selected locations for providing an output of sensed wind speed and wind direction at the selected locations, a plant model responsive to the sensed wind speed and direction within a finite horizon, a current state of the launch vehicle, and control commands of the launch vehicle to predict the trajectory of the launch vehicle, an error circuit responsive to the predicted trajectory and a reference trajectory to produce a trajectory error, and an optimizer responsive to the trajectory error and configured to provide control commands to compensate for wind load over the finite horizon of the launch vehicle by reducing the total angle of attack.

Abstract: The present application discloses plug and play modular mission payloads in the context of aerial vehicles, and a supporting open system architecture that moves the control function of mission payloads away from the ground station and into the aerial vehicle. The plug-and-play (PnP) modular mission payloads and web-based payload interface software resides in a payload computer in the vehicle, and this is networked via a uniform resource locator (URL) addressing scheme to a ground control station. Consequently, when new payload types are added to the system, integration issues and costs are minimized.

Abstract: A video guidance sensor system for use in automated docking of a chase vehicle with a target vehicle wherein the chase vehicle includes a laser rangefinder that uses pulse or phase time of flight measurement to measure distance. The laser rangefinder includes a diode laser pulse or phase driver that produces an output signal to a timing element and simultaneously operates a laser diode. The laser diode produces an intense light beam of a predetermined wavelength which is directed to retroreflectors that are positioned on a passive target. The laser rangefinder includes an avalanche photodetector that produces a corresponding output signal when detecting light reflected from the retroreflectors. The timing element measures a time interval between the output of the laser diode and the detection of light and supplies a corresponding output signal to a computer in order to determine the range of the target vehicle relative to the chase vehicle.

Abstract: The invention provides a method of providing protection against the effects of protons of solar origin, so as to make it possible for a star sensor to be reconfigured.

Abstract: An attitude control system (ACS) for a spacecraft includes an attitude control assembly (ACA) interface, a control moment gyro (CMG) array and a reaction wheel assembly (RWA) control unit. The ACA interface converts torque commands received from the RWA control unit into CMG gimbal rates for the CMG array. The ACA interface receives CMG gimbal angles from the CMGs and converts the CMG gimbals angles into RWA speeds, which are then provided to the RWA control unit.

Abstract: A control system for a UAV includes control translations which maximize operational employment of the UAV payload. By determining spatial references, and then using the references to transform the control stick commands, the operator treats the UAV as a point source. For control through imagery from onboard mission sensors, the transformations provide for the UAV to move itself and achieve payload orientation.

Abstract: A method for operating a space vehicle equipped with at least one device cooperating with a celestial body, wherein rotation of the space vehicle takes place around at least one axis of the space vehicle. The rotation of the space vehicle is slowed when the device cooperating with the celestial body is oriented substantially toward the celestial body. The device can be a solar panel intended to face the sun and produce electrical energy.

Abstract: A fuel efficient technique for changing the inclination, with respect to the Earth's equator, for a satellite includes first maneuvering the satellite to the moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so-called fuzzy boundary or weak stability boundary. A negligibly small maneuver can then bring it back to the Earth on a reverse BCT to the desired Earth inclination. Another maneuver puts it into the new ellipse at the earth. In the case of satellites launched from Vandenberg AFB into LEO in a circular orbit of an altitude of 700 km with an inclination of 34°, approximately 6 km/s is required to change the inclination to 90°. The previous flight time associated with this method was approximately 170 days. A modification of this method also achieves a significant savings and unexpected benefits in energy as measured by Delta-V, where the flight time is also substantially reduced to 88 or even 6 days.

Abstract: A propellant utilization system for a space vehicle having at least a first thrust source and a second thrust source. The propellant utilization system utilizes a common set of algorithms to generate mixture ratios for each thrust source as each thrust source becomes active. The propellant utilization system includes a processing system comprising sequencer logic, propellant logic, and mixture ratio logic. The sequencer logic determines when a thrust source is active, e.g. one of the first and second thrust sources, and provides flight parameters for the active thrust source to the propellant logic and the mixture ratio logic. The propellant logic processes information from propellant sources connected to the active thrust source, using the flight parameters for that thrust source, to determine an amount of remaining propellant in each source.

Abstract: A method for correcting differences in measurements between sensors. Misalignment of a roll error of a sun sensor alignment is estimated with respect to a reference roll value measured by an earth sensor. Sun sensor assembly elevation angle residual is calculated utilizing the estimated sun sensor assembly roll misalignment. A yaw attitude of the spacecraft is updated based upon the calculated sun sensor assembly elevation residual.

Abstract: The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and which requires a shorter period of setting time for obtaining a target attitude. The maneuver device is provided with: a feed forward torque instruction signal generator 8 which outputs a feed forward torque instruction signal 11 based on a maneuver plan; a thruster 10 which outputs control torque based on the feed forward torque instruction signal 11; and an attitude control signal calculator 6 to which an attitude angle and an angular velocity of the artificial satellite as well as a target attitude angle and a target angular velocity are input and which outputs an attitude control signal 13. The maneuver device is further provided with a disturbance compensating signal calculator to which the feed forward torque instruction signal 11 and a detected angular velocity signal 16 are input, and which generates and outputs a disturbance compensating signal 12.

Abstract: Methods and structures are provided that enhance attitude control during gyroscope substitutions by insuring that a spacecraft's attitude control system does not drive its absolute-attitude sensors out of their capture ranges. In a method embodiment, an operational process-noise covariance Q of a Kalman filter is temporarily replaced with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. The error effect of the substituted gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range. In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.

Abstract: A method and apparatus for calibrating rate sensor measurements to compensate for rate sensor scale factor non-linearities. An outer loop compensator compares current rate sensor scale factor estimates with the current scale factor non-linearity compensation, and deviations from the current non-linearity compensation are corrected in an updated compensation.

Abstract: An inertial sensor includes an inertial reference for determining the attitude and position of a satellite or satellite parts. A test mass is situated in a space which is essentially free of electric and magnetic fields and is enclosed by a housing, and optical measuring sections are set up between reference elements on the housing and on the test mass, for determining the attitude and/or position of the test mass relative to the reference elements on the housing. A measuring arrangement is situated outside the housing for the optical measuring sections. The optical measuring sections are constructed as optical interferometric measuring elements; and the attitude and position of the test mass can be adjusted by means of the pressures of light exerted on the optical interferometric measuring sections upon the test mass.

Abstract: An autonomous navigation system for an orbital platform incorporating a global positioning system based navigation device optimized for low-Earth orbit and medium-Earth orbit applications including a 12-channel, GPS tracking application-specific integrated circuit operating in concert with a microprocessor implementing an extended Kalman filter and orbit propagator which autonomously generates estimates of position, velocity, and time to enable planing, prediction and execution of event-based commanding of mission operations.

Abstract: A GPS positioning system that has: at least one mobile station to measure a positional coordinate by receiving radio wave from a satellite; a plurality of fixed stations that have predetermined positional coordinates and receive radio wave from the satellite; and arithmetic processing means, which is connected with at least one mobile station and a plurality of fixed stations via communication, and that transmits correction data suitable for the mobile station to the mobile station based on positioning data transmitted from the mobile station.

Abstract: A system for alerting the occupant of a vehicle that the vehicle is in or approaching a zone of awareness. For instance, the system may be used to alert the pilot or flight crew of an aircraft that the aircraft is on or approaching a runway, and may identify the specific runway. Thus, the invention also provides a method of reducing unintentional incursions of taxiing aircraft onto runways. The system generally includes a storage device or memory, a positioning system such as a GPS that typically repeatedly determines the location of the vehicle, an alerting device such as an alarm or display, and a processor, all of which are typically located on the vehicle. The processor typically compares the location of the vehicle with stored location information for various zones of awareness and initiates an alert when appropriate. The processor may take into consideration the velocity and direction of travel of the vehicle. Zones of awareness may be defined by coordinates, such as the end points of line segments.

Abstract: A method for maintaining three axis control of a geosynchronous spacecraft without body rate measurements using reaction wheel assemblies. Earth sensor assembly angle measurements are utilized for high-bandwidth roll and pitch control. A positive pitch momentum bias is stored in the reaction wheel assemblies. A gyroscopic feedforward torque is applied to rotate reaction wheel assembly momentum in a yaw/roll plane of the spacecraft at orbit rate. A dynamic mode that couples yaw and roll axes and that results from applying the gyroscopic feedforward torque and the high-bandwidth roll control is damped based on earth sensor assembly roll measurements.

Abstract: Attitude determination and control systems are provided that combine attitude measurements from all spacecraft payloads to determine a master attitude estimate for a master payload and relative slave attitude estimates for the remaining slave payloads. These estimates are then used to control the attitudes of spacecraft elements that correct the absolute and relative attitude errors. These systems significantly enhance attitude accuracy when compared to systems that realize independent payload estimates, determine payload attitudes. These systems also provide significant processing advantages (e.g., simpler algorithms, reduced data throughput and slower processing rate).

Abstract: The present invention is a GPS system for navigation and attitude determination, comprising a sensor array including a convex hemispherical mounting structure having a plurality of mounting surfaces, and a plurality of antennas mounted to the mounting surfaces for receiving signals from space vehicles of a GPS constellation. The present invention also includes a receiver for collecting the signals and making navigation and attitude determinations. In an alternate embodiment the present invention may include two opposing convex hemispherical mounting structures, each of the mounting structures having a plurality of mounting surfaces, and a plurality of antennas mounted to the mounting surfaces.

Abstract: A vehicular position and attitude can be accurately determined from GPS signals by positioning a master and two slave orthogonally disposed receiver antennas rigidly mounted together enabling the rotation of the slave receiver antennas about the master antenna. The slave antennas are rotated and then dither back and forth so that differential phase measurements between the slave antennas and the master antenna is nulled in carrier phase alignment of the GPS signals for determining azimuth angles and elevation angles to the GPS satellites for determining the position and attitude of the vehicle.

Abstract: The present invention provides methods of and apparatus for determining the inertial attitude of an aerospace vehicle. In one embodiment, the invention provides a rotational astronomical object-sighting concept to determine the inertial attitude of an axis of the aerospace vehicle without the star identification or dragback. In another embodiment, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor.

Abstract: A fuel efficient technique for changing the inclination, with respect to the Earth's equator, for a satellite includes first maneuvering the satellite to the moon on a BCT (Ballistic Capture Transfer). At the moon, the satellite is in the so-called fuzzy boundary or weak stability age boundary. A negligibly small maneuver can then bring it back to the Earth on a reverse BCT to the desired Earth inclination. Another maneuver puts it into the new ellipse at the earth. In the case of satellites launched from Vandenberg AFB into LEO in a circular orbit of an altitude of 700 km with an inclination of 34°, approximately 6 km/s is required to change the inclination to 90°. The previous flight time associated with this method was approximately 170 days. A modification of this method also achieves a significant savings and unexpected benefits in energy as measured by Delta-V, where the flight time is also substantially reduced to 88 or even 6 days.