Abstract: A bearing assembly for stabilizing rotation of a rotatable element includes a gas bearing acting as a bearing for the rotatable element. An ultrasonic vibration pump is coupled to the gas bearing for providing pressurized gas to the gas bearing, and a sealed housing encloses the gas bearing and the ultrasonic vibration pump. Also described is a reaction wheel assembly for a spacecraft. The reaction wheel assembly includes a reaction wheel and the bearing assembly for stabilizing rotation of the reaction wheel.

Abstract: Methods and systems for multi-hypothesis object tracking for automated driving systems. One system includes an electronic processor configured to receive environment information and generate pseudo-measurement data associated with an object within an environment of the vehicle. The electronic processor is also configured to determine, based on the environment information and the pseudo-measurement data, a set of association hypotheses regarding the object. The electronic processor is also configured to determine, based on the set of association hypotheses, an object state of the object. The electronic processor is also configured to control the vehicle based on the determined object state.

Abstract: A method (50) of acquiring images of a terrestrial region Z using a spacecraft (10) in non-geostationary orbit around the Earth (30), the spacecraft includes an observation instrument associated with a ground footprint of length L along the direction of travel, the method includes: a step (51) of observing a portion P1 of the terrestrial region Z, including a step of controlling the attitude of the spacecraft (10) during which the ground footprint is kept stationary during the entirety of the step of observing portion P1, and a step of acquiring an image of portion P1, a step (52) of modifying the pitch attitude of the spacecraft (10) so as to place the ground footprint over a portion P2 of the terrestrial region Z, and a step (53) of observing portion P2 of the terrestrial region.

Abstract: Cell-based systems may interlock in a reconfigurable configuration to support a mission. Space systems, for example, of a relatively large size may be assembled using an ensemble of individual “cells”, which are individual space vehicles. The cells may be held together via magnets, electromagnets, mechanical interlocks, etc. The topology or shape of the joined cells may be altered by cells hopping, rotating, or “rolling” along the joint ensemble. The cells may be multifunctional, mass producible units. Rotation of cell faces, or of components within cells, may change the functionality of the cell. The cell maybe collapsible for stowage or during launch.

Abstract: An attitude control apparatus for a satellite includes: at least three electric motors, wherein the at least three electric motors are arranged in such a way that a torque may be generated with any orientation of an associated torque vector, and a controller, wherein the controller is configured to drive the at least three electric motors based on a torque controller. The torque controller is adapted to operate the at least three electric motors outside a rest state only when an acceleration torque and a braking torque are required to execute an agile attitude change maneuver. There is also described an associated method.

Abstract: A control system for reducing disturbance torque of a spacecraft is disclosed. The spacecraft revolves around a celestial body surrounded by an atmosphere. The control system includes processors in electronic communication with one or more actuators and a memory. The memory stores data into a database and program code that, when executed by the one or more processors, causes the control system to instruct the spacecraft to enter a safing mode. In response to entering the safing mode, the control system instructs the one or more actuators to align a principal axis of the spacecraft with a vector that is normal to the orbit around the celestial body. The control system also instructs the actuators to rotate the spacecraft about the principal axis, where a rotational orientation of the spacecraft relative to the celestial body is shifted by about one-half a rotation about the principal axis.

Abstract: A system includes a motor configured to be coupled to a non-rigid load and a control system disposed within, or communicatively coupled to, a drive system configured to control an operation of the motor. The control system includes a processor and a memory accessible by the processor. The memory stores instructions that, when executed by the processor, cause the processor to generate a smooth move input profile to control the operation of the motor based on inputs specifying a desired operation of the motor, apply a notch filter having a notch filter frequency to the smooth move input profile to produce a filtered smooth move input profile, and send a command to the drive system based on the filtered smooth move input profile, wherein the command is configured to adjust the operation of the motor.

Abstract: A bonding layer 3 is formed over a piezoelectric material substrate, and the bonding layer is made of one or more materials selected from the group consisting of silicon nitride, aluminum nitride, alumina, tantalum pentoxide, mullite, niobium pentoxide and titanium oxide. A neutralized beam is irradiated onto a surface of the bonding layer and a surface of a supporting body to activate the surface of the bonding layer and the surface of the supporting body. The surface of the bonding layer and the surface of the supporting body are bonded by direct bonding.

Abstract: A method and system for generating extended satellite ephemeris data for a satellite for use by a terminal apparatus is described. Historical orbital parameter data (e.g. from historical TLEs) are used to build a predictor for future values of the orbital parameters. The orbital parameter prediction model is then stored on the terminal apparatus (or an existing model can be updated), and at a desired epoch the pre-stored orbital parameter predictor is used to produce estimates of the orbital parameters for the satellite for that epoch (ie synthetic or predicted TLEs). These predicted orbital parameters (predicted TLE) are then used as input to a standard orbit predictor to produce an estimate of the satellite's location, for example to determine satellite pass times. This method allows generation satellite ephemeris data which is valid out to a year or more as compared to the normal validity of ephemeris data of a few hours or days.

Abstract: A method, for providing spatial stability and electrical power with a hybrid power source and control moment gyroscope (HPCMG), includes producing spatial stability force for the HPCMG by spinning a central mass within a first transverse gimbal assembly about a first axis of rotation of a control moment gyroscope (CMG). The CMG includes the first transverse gimbal assembly, the central mass, and a second gimbal assembly rotationally connected to the first transverse gimbal assembly. The first transverse gimbal assembly is rotationally connected to the central mass at a first position of the first transverse gimbal assembly and at a second position of the first transverse gimbal assembly along the first axis of rotation. The method includes producing a voltage potential with the central mass. The method includes charging or discharging the central mass through conductive bearings.

Abstract: Proposed method and device could be applied as controlling system facilitating the maneuverability and stabilization parameters of UAV, as well as various flight objects and small satellites. The aim of the invention is to reduce the occurrences of imbalances of UAV in strong wind or atmospheric turbulence situations and simultaneously rehabilitation of maneuverability and stabilization parameters without increasing flight speed. Thus, in order to upgrade the above mentioned parameters of UAV along with the gyroscopic momentum compensation it's necessary to increasing of the kinetic moment by means of gyroscope.

Abstract: An apparatus for pointing wireless communication antennas may include (1) a mount that secures (A) a wireless communication antenna that transmits wireless communication signals to a remote wireless communication antenna that is secured on a remote mount and (B) an array of pointing antennas that receives, from a remote array of pointing antennas mounted on the remote mount, a beacon signal that indicates a location of the remote array of pointing antennas, (3) a motorized drive that physically orients the mount, and (4) a processing device that (A) determines, based at least in part on the beacon signal, an orientation of a boresight axis of the wireless communication antenna relative to a boresight axis of the remote wireless communication antenna and (B) directs the motorized drive to orient the mount such that the wireless communication antenna's boresight axis aligns with the remote wireless communication antenna's boresight axis.

Abstract: A method for initial alignment of an inertial navigation apparatus, comprising the following steps: providing an apparatus loaded with a sensor, and preprocessing the sensor; carrying out relative alignment to calculate an installation error angle of the sensor; carrying out absolute alignment to calculate an installation attitude angle error of the sensor to increase an accuracy of an error attitude angle calculated during the relative alignment. The relative alignment process calculates a relative error attitude angle, the relative error attitude angle being used as the initial value for attitude error in a stat vector in the absolute alignment process, thereby accelerating convergence of the Kalman filter. Alignment precision is further enhanced by the absolute alignment process.

Abstract: Unmanned aerial vehicles (1), and methods of flying such, comprising at least four rotors (2) arranged such that the plane of rotation of each rotor (2) is co-planar with a face of a notional polyhedron, and wherein each face of the notional polyhedron is co-planar with the plane of rotation of at least one rotor (2). Such methods comprise: a first step of flying the vehicle (1) using a first rotor set (2a-c) to provide lift; and, a second step using a second rotor set (2d-f) to provide lift; wherein, the second rotor set (2d-f) includes at least one rotor (2) that is not used to provide lift in the first step or that operates so that airflow through the rotor (2) is in the opposite direction to that through the rotor (2) during the first step; and, wherein at least one of the first and second sets (2a-c, 2d-f) comprises a plurality of rotors (2).

Abstract: A storage unit holds context information including a minimum energy state, a current state, a minimum energy, a current energy, a temperature, and a bias about each of a plurality of trials. A control unit includes a counting unit counting update processing repeated in each trial and another counting unit counting trials performed. The control unit repeats update control about the update processing a certain number of times in each trial on the context information, based on spin information and an update target energy change. A calculation unit calculates energy change candidates based on a generated random number, a stored weight, the spin information, and the current state and temperature updated by the update control. A selection unit selects the update target energy change from the candidates calculated by the calculation unit based on a random number.

Abstract: In an aerospace vehicle in which a satellite is detachably adapted to the first stage rocket and the second stage rocket, an electronic device 65 is installed in the satellite 60, wherein the electronic device 65 controls the first stage rocket 10 and the second stage rocket 20 before the satellite is detached from the first stage rocket 10 and the second stage rocket 20. Thereby, it is unnecessary to provide sensors, radio devices and electronic devices for exclusive use with respect to the first stage rocket 10 and the second stage rocket 20 so that a manufacturing cost of the rockets 10 and 20 can be reduced and a total weight of the rockets 10 and 20 can become lighter by omitting these equipments. Thus, an aerospace vehicle system within the rockets 10 and 20 can be simplified and a launch of the aerospace vehicle can be prepared within a short period.

Abstract: Systems and methods for maintaining and stabilizing the position and orientation of a payload attached to a high-altitude balloon are provided. A payload may be attached to a powered gimbal. The powered gimbal may be configured to orient and position the payload in a plurality of directions corresponding to a first, second, and third rotational axis of the balloon-mounted payload system. After the payload is positioned by the powered gimbal, the position and orientation of the payload may be maintained and stabilized by one or more rotational stabilization devices. The stabilization by the one or more rotational stabilization devices can occur along any one, or combination of, the first, second, and third rotational axes.

Abstract: A method for controlling each rotor on a multirotor aircraft includes receiving one or more sensed states of the multirotor aircraft, and determining a control scheme dynamically using the sensed one or more states of the multirotor aircraft. Receiving one or more sensed states can include receiving one or more of a rotational rate or acceleration of the aircraft, or any other suitable aircraft state.

Abstract: A method of controlling the attitude of a spacecraft in spinning around itself with a non-zero total angular momentum HTOT. The spacecraft includes a set of inertia flywheels configured to form an internal angular momentum HACT. The axis of the total angular momentum HTOT is aligned with a principal axis of inertia of the spacecraft, in the course of which the inertia flywheels are controlled to form an internal angular momentum HACT. The following expression, in which J is the inertia matrix of the spacecraft: Hact×J?1(Htot?J?1Htot) is negative if the principal axis of inertia targeted is the axis of maximum inertia of the spacecraft and is positive if the principal axis inertia targeted is the axis of minimum inertia of the spacecraft.

Abstract: When the number of CMGs is represented by n (n is an integer of 4 or more), (n?3) gimbal angles out of n gimbal angles corresponding to the n CMGs are set as free parameters, and an algebraic equation representing a relationship among three gimbal angles out of the n gimbal angles, the free parameters, and an angular momentum of all the CMGs is used to solve the algebraic equation while changing the free parameters within set ranges, to thereby obtain solutions of the gimbal angles of the plurality of CMGs required for achieving a given angular momentum.

Abstract: An ADCS module may be configured to use coordinate data from 2D photodiodes in one or more sun sensors to determine a sun vector. The ADCS module may then use the sun vector in reference to its own body faced (BF) coordinate system to calculate a change in the orientation of the space vehicle. The change in orientation mechanism may be accomplished by reaction wheels, ion thrusters, or other orientation altering mechanisms. A miniature, intelligent star tracker may be included that improves satellite attitude determination and pointing accuracy. An improved reaction wheel assembly may be included that is more robust and suitable for inclusion in small space vehicles.

Abstract: Apparatus, such as can be included in an electronic device, performs a method for operating apparatus configured for combined vibratory and projection functions. The apparatus includes a mirror configured to reflect a projected image and a motor. The mirror is coupled as a counterweight to the motor to configure the apparatus for a vibratory function. The motor is further configured to translate and rotate the mirror about a first axis to align the mirror with a projector to reflect the projected image during a projection function.

Abstract: An unmanned vehicle including a body and a frame structure extending from the body and supporting a plurality of propeller assemblies, each propeller assembly including at least one propeller and a corresponding motor with the motor housed in a watertight housing or coated and made corrosion resistant. The propellers comprise a first subset of propellers of the propeller assemblies and a second subset of propellers of the propeller assemblies which rotate in a plane positioned below a plane in which the first subset of propellers rotate, wherein said first and second subset of propellers are configured for independent operation of one another as the vehicle transitions from an air medium to a water medium.

Abstract: A pod assembly which includes a pod and an inertial measurement unit positioned within the pod. The pod assembly further includes deceleration hardware associated with the pod. A method for using a pod assembly, includes a step of securing the pod assembly, which includes a pod and an inertial measurement unit contained within the pod, to an aircraft and a step of operating the aircraft with the inertial measurement unit operating sensing movement and alignment of the pod and storing data corresponding to the sensed movement and alignment of the pod on a data collection assembly connected to the inertial measurement unit. The method further includes a step of releasing the pod assembly from the aircraft and a step of utilizing deceleration hardware associated with the pod to reduce rate of descent of the pod assembly released from the aircraft.

Abstract: A hybrid power source and control moment gyroscope (“HPCMG”) is disclosed. The HPCMG includes a control moment gyroscope (“CMG”), a first conductive bearing, and a second conductive bearing. The CMG includes a first transverse gimbal assembly, a central mass that produces a voltage potential, and a second gimbal assembly rotationally connected to the first transverse gimbal assembly. The first transverse gimbal assembly is rotationally connected to the central mass along a first axis of rotation and the central mass is configured to spin about the first axis of rotation and the first transverse gimbal assembly is configured to rotate about a second axis of rotation of the second gimbal assembly. The first conductive bearing rotationally connects the central mass with the first position of the first transverse gimbal assembly along the first axis of rotation.

Abstract: A rotor assembly for deployment within a momentum control device that enables near-zero revolutions per minute (RPM) sensing, and method for making same, are provided. The provided rotor assembly utilizes a magnet coupled to the rotor shaft and a stationary sensor element to detect magnetic flux from the magnet and derive reliable near zero RPM therefrom.

Abstract: Two methods of combining multiple response types into a single flexible command model are provided and include receiving a pilot stick input, generating an aircraft response to the pilot stick input that is a continuous blend of response types by including calculable time-varying coefficients set as a function of a magnitude of the pilot stick input and other aircraft states such as airspeed, imposing at least an angular acceleration command limit and using other non-linear elements to optimize the aircraft response to the pilot stick input.

Abstract: Provided is a spatial stabilization apparatus that realizes high spatial stability with a single inertial sensor.

Abstract: A device for determining relative orientation between two locations including an imager, an inertial-orientation-sensor firmly attached to the imager for determining information relating to the orientation thereof and which exhibits drift and a processor coupled with the imager and with the inertial-orientation-sensor. The processor determines a first orientation-measurement and first time-tag when the device is oriented with a first-orientation-indicator located at a first location. The processor determines a second-orientation-measurement and second time-tag when the device is oriented with a second-orientation-indicator located at a second location. The processor determines a third-orientation-measurement and third time-tag when the device is oriented again with the first-orientation-indicator.

Abstract: An energy storage system is provided with one or more local energy storage systems and a procedure to operate such an energy storage system. The local energy storage systems are connected to a non-local power supply grid and/or one or more local power supply grids each compromising at least one local control unit, which is intended at least for control of the respective energy storage system for localized regulating and system tasks (LRS) for the one or more respective local power supply grids. The energy storage system furthermore comprises a central control unit connected via the communication network to the respective local control units, which is intended for control (SNL) of absorbing (En) and emitting (Ep) energy of the one or more local energy storage systems to the non-local power supply grid, with the central control unit being equipped to dispose of all portions of the local storage capacities (LSKg).

Abstract: A geometry-based flight control system is disclosed. The geometry-based flight control system receives a set of inputs associated with a requested set of forces and moments to be applied to the aircraft and computes an optimal mix of actuators and associated actuator parameters to achieve to an extent practical the requested forces and moments. Computing the optimal mix of actuators and associated actuator parameters includes taking into consideration dynamically varying effectiveness of one or more actuators based on a dynamic state of the aircraft. For example, the dynamic state may comprise an angle of a tiltwing of the aircraft or an angle of a tilt rotor of the aircraft.

Abstract: An attitude estimator that uses sun sensor outputs as the only attitude determination measurements to provide three-axis attitude information. This is accomplished by incorporating the Euler equation into the estimator. An unscented Kalman filter is employed to accommodate various nonlinear characteristics and uncertainties of the spacecraft dynamics and thus improve the robustness and accuracy of the attitude estimate.

Abstract: Various embodiments provide an anti-jamming method in a satellite communication (SATCOM) system. A waveform modulation is performed on a waveform of a source data to provide a modulated waveform transmitted to a transmitter and then to a receiver. The modulated waveform transmitted from the receiver is demodulated to provide a demodulated waveform. In response to interference, a game reasoning process is performed on the demodulated waveform to provide an anti-jamming adaptive waveform. By performing a game-optimal waveform modulation on the anti-jamming adaptive waveform, a frequency hopping process is used to guide the data transmission, such that the source data is transmitted with a minimal probability of being interfered.

Abstract: A method for controlling an attitude control system (30) for a space vehicle (10), the attitude of the space vehicle being controlled during at least one preparation phase followed by an observation phase during which an image capture is performed. In an attitude control system that includes a maneuvering subsystem (300) which includes at least one reaction wheel, the method includes, during the at least one preparation phase, a preparation step (50), during which commands are issued to the maneuvering subsystem in order to control the attitude of the space vehicle, followed by a step (55) of stopping the at least one reaction wheel prior to the observation phase. In addition, during the observation phase, a fine control subsystem (310) having a lower vibration signature than that of the maneuvering subsystem, is sent commands in order to control the space vehicle's attitude. An attitude control system for a space vehicle is also described.

Abstract: An improved spacecraft actuator wheel is provided which can be operated as a momentum wheel, a reaction wheel or a gimbal. The actuator wheel has a central cavity. One or more battery modules are located within the actuator wheel's central cavity. The battery modules supply power to one or more electronic components affixed to the actuator wheel or mounted on the spacecraft frame via an electrical harness. In addition, the actuator wheel's central cavity is pressurizeable for storing spacecraft propellant which can be controllably diverted to the spacecraft's thrusters through conduits and flow valves.

Abstract: In one embodiment, a method comprises: establishing a first broadband data link between a first mobile narrowbeam transceiver positioned on a vehicle and a first fixed narrowbeam transceiver mounted along a prescribed path of the vehicle; and switching from the first broadband data link, by the first mobile narrowbeam transceiver, to a second broadband data link with a second fixed narrowbeam transceiver mounted along the prescribed path after the first fixed narrowbeam transceiver, enabling the vehicle to maintain continuous broadband access to a wide area network via a prescribed sequence of the fixed narrowbeam transceivers along the prescribed path.

Abstract: One embodiment is directed towards an inertial measurement unit (IMU) for measuring an input rate of rotation about an input axis. The IMU includes a first three dimensional gyroscope disposed such that a first axis of its three axes is oriented at a skew angle in degrees away from a reference plane, wherein the reference plane is normal to the input axis. The IMU also includes one or more processing devices coupled to the first gyroscope. The IMU also includes one or more data storage devices coupled to the one or more processing devices, the one or more data storage devices including instructions which, when executed by the one or more processing devices, cause the one or more processing devices to calculate the input rate of rotation based on dividing a sensed rate of rotation about the first axis by the sine of the skew angle.

Abstract: A method of providing automatic collision avoidance in a vehicle with a front wheel electric power steering (EPS) system and rear wheel active rear steering (ARS) system and an automatic collision avoidance system are described. The method includes generating a vehicle math model including the control variables, designing a steering control goal as a criterion to determine the control variables, and implementing a model predictive control to solve the steering control goal and determine the control variables. The method also includes providing the control variables to the EPS system and the ARS system to respectively control a front actuator associated with front wheels and a rear actuator associated with rear wheels.

Abstract: A method of deploying a modular space station comprises placing an initial space station module in space in a first deployment, the initial space station module including a first control law and momentum component that provides an initial solution for guidance, navigation, and control (GNC) during the first deployment. A first space station modular segment is joined with the initial space station module in a second deployment to produce a first joint configuration of the space station. A second control law and momentum component provides a second solution for GNC of the first joint configuration during the second deployment. A second space station modular segment is joined to the first joint configuration in a third deployment to produce a second joint configuration of the space station. A third control law and momentum component provides a third solution for GNC of the second joint configuration during the third deployment.

Abstract: A system for managing momentum accumulation of a spacecraft in orbit may include a reaction wheel assembly for controlling an attitude of a body of a spacecraft, the body defining at least one face, and absorbing momentum, a plurality of arcjet thrusters coupled to the face to generate thrust, and a control processor coupled to the plurality of arcjet thrusters for controlling the thrust, wherein actuation of each arcjet thruster of the plurality of arcjet thrusters is configured to produce a net momentum accumulation in the reaction wheel assembly that is below a momentum saturation point of the reaction wheel assembly.

Abstract: A space vehicle has a frame with a configuration of fiber optic gyroscopes (FOGs) and control moment gyroscopes (CMGs) at the outer perimeter of the frame. The FOGs and CMGs provide guidance and control for the space vehicle. This arrangement results in the largest possible FOG and CMG diameters, and therefore yielding the highest signal sensitivities and precision inertial control of vehicle orientation and pointing. Because the configuration places these guidance and control system components at the perimeter of the vehicle, it also provides a platform for multi-aperture signal channels in the interior of the vehicle by freeing up volume within the vehicle.

Abstract: Various embodiments of the present invention include an attitude control system for use with small satellites. According to various embodiments, the system allows rapid retargeting (e.g., high slew rates) and full three-axis attitude control of small satellites using a compact actuation system. In certain embodiments, the compact actuation system includes a plurality of single-gimbaled control moment gyroscopes (SGCMG) arranged in a pyramidal configuration that are disposed within a small satellite.

Abstract: A method for characterizing a transmitting antenna of a satellite in orbit comprising a payload comprising signal amplification means, includes: configuring the said amplification means to generate thermal noise at the input of the transmitting antenna; acquiring, by means of a ground station, the signal transmitted by the transmitting antenna on the downlink of the satellite over a predetermined time period; during the said predetermined time period, controlling the said satellite in orbit in order to impart on it an angular bias of predetermined variation and recording this variation; correlating the measurement of the signal transmitted on the downlink and the variation of angular bias of the satellite in order to derive therefrom the gain variations of the transmitting antenna as a function of the angular bias of the satellite.

Abstract: A method for reducing the angular momentum of a spacecraft comprises a propulsion device able to generate a torque along a Z axis corresponding to an axis of maximum inertia or minimum inertia of the spacecraft, an X axis and a Y axis forming with Z an orthogonal frame; an angular momentum accumulating device, able to generate an angular momentum and a torque along the three axes; and a set of sensors able to measure the angular velocities and estimate the angular momentum of the spacecraft. The method comprises a first step of aligning the angular momentum of the spacecraft along the Z axis, consisting in slaving the angular momentum of the angular momentum accumulating device to the angular velocity of the spacecraft, and a second step of reducing the angular momentum of the spacecraft using a torque generated by the propulsion device.

Abstract: Methods and apparatus are presented for spacecraft operation using non-Eigen axis attitude transitions via control momentum gyroscopes (CMGs) to avoid or mitigate singularities by providing a time-varying attitude command vector including a plurality of time-varying attitude command signals or values representing a plurality of spacecraft states and control trajectories as a guidance command input to an attitude controller of the spacecraft without modifying the spacecraft feedback control law.

Abstract: Spacecraft momentum management techniques are coordinated with station-keeping maneuvers or other delta-V maneuvers. A body stabilized spacecraft attitude is controlled, the spacecraft including at least one momentum/reaction wheel, and a set of thrusters. A first momentum storage deadband limit is adjusted, the adjustment being related to a first delta-V maneuver window. A momentum management strategy is executed with the adjusted first momentum storage deadband limit such that a first thruster firing that performs desaturation of the momentum/reaction wheel also provides velocity change beneficial to the first delta-V maneuver.

Abstract: Methods and apparatus are presented for spacecraft operation in which a control problem is formulated using a control law or steering law as a path constraint or as a dynamic constraint, and the control problem is solved to provide a guidance command trajectory for use in operating spacecraft control momentum gyroscopes to guide the spacecraft from an initial state to a desired final state.

Abstract: Various embodiments of the present invention include assemblies and methods for minimizing the amplitude of attitude jitter. In one embodiment, a split flywheel assembly includes a plurality of independent concentric flywheels axially aligned and in operable engagement with one another such that each flywheel is configured to be independently controlled in order to manipulate the phase difference therebetween.

Abstract: A space vehicle has a frame with a configuration of fiber optic gyroscopes (FOGs) and control moment gyroscopes (CMGs) at the outer perimeter of the frame. The FOGs and CMGs provide guidance and control for the space vehicle. This arrangement results in the largest possible FOG and CMG diameters, and therefore yielding the highest signal sensitivities and precision inertial control of vehicle orientation and pointing. Because the configuration places these guidance and control system components at the perimeter of the vehicle, it also provides a platform for multi-aperture signal channels in the interior of the vehicle by freeing up volume within the vehicle.

Abstract: A method and apparatus for estimating gyro scale factor during normal spacecraft operations, using small attitude motions that are compliant with mission pointing accuracy and stability requirements and a signal processing method that specifically detects the intentionally induced motions. This process increases operational availability by avoiding the need to take the spacecraft offline for large calibration maneuvers.