Abstract: Provided are various spacecraft propulsion systems, and associated methods of operation. A spacecraft comprises an ion propulsion system and an ion blocker suspended from the spacecraft via one or more electrically insulated tethers. The ion propulsion system is configured to generate a first propulsive force by emitting a charged ion beam in a direction with an ion velocity vector comprising an ion vector component that is perpendicular to a magnetic field of a planet, such as Earth. The magnetic field causes the ion beam to curve toward the ion blocker at a trajectory such that ions within the ion beam are blocked by the ion blocker to generate a second propulsive force on the ion blocker. The ion blocker blocks the ions by contacting or deflecting the ions. The ion blocker is positioned approximately twice the gyroradius of the ion beam trajectory.

Abstract: A space vehicle includes one or more magnetorquers operable to change an attitude of the space vehicle in an external magnetic field, each magnetorquer comprising a coil, and a switching circuit for short-circuiting the coil of at least one of the magnetorquers so that a closed electric circuit comprising said coil is formed, for damping tumbling motion of the space vehicle in the external magnetic field. The switching circuit is configured to short-circuit the coil of the at least one magnetorquer upon occurrence of a condition indicative of end-of-life or failure of the space vehicle. The application further relates to a corresponding method of operating a space vehicle.

Abstract: A method for desaturating reaction wheels of a spacecraft having a magnetic dipole is provided. The method includes orienting the spacecraft relative to an external magnetic field to apply a torque to the spacecraft via the magnetic dipole in a direction opposing momentum stored in the reaction wheels; and using the applied torque to unload at least some of the momentum stored in the reaction wheels. A corresponding spacecraft and non-transitory computer-readable medium are also provided.

Abstract: The invention relates to small satellites capable to fly in formation (10), in particular nano- or picosatellites with a mass of 10 kg or less, for LEO applications, comprising a housing (12) and at least one plug-in board (14) arranged in the housing (12) with a predetermined functionality and a propulsion system (16) for generating a directed pulse in the direction of the flight trajectory Tk. It is proposed that the small satellite (10) comprises an independent and autonomously working collision avoidance system (18), which is capable of adapting a trajectory correction Tkk of the trajectory Tk by the propulsion system (16), when a collision with an object (30) is expected. In a further independent aspect, the invention relates to a formation (100) composed of several small satellites capable to fly in formation (10), wherein a relative position and flight trajectory Tk of each small satellite (10) is modifiable via the independently and autonomously working collision avoidance system (18).

Abstract: This disclosure relates to a satellite attitude control apparatus, comprising a spherical inertial body (1), which is arranged to be rotated in space in a desired rotational mode, one or more rotational feedback sensors. The apparatus further comprises a set of supports (3) and/or actuators (2), whereas one or more actuator is used to create rotational momentum to the spherical inertial body through friction. A set of three or more actuators and supports is used in order to keep the inertial body in place. Three or more supporting elements (supports and actuators) are used in combination with magnetic attraction or four or more supporting elements are used in case magnetic attraction is not in use. Supports can be passive or semi-active in order to reduce friction.

Abstract: A multi-degree-of-freedom electromagnetic machine includes a spherical armature, a first coil, a second coil, a third coil, a structure, a stator, and a bracket. The spherical armature has a first axis of symmetry, a second axis of symmetry, a third axis of symmetry, and the first, second, and third axes of symmetry are disposed perpendicular to each other. The spherical armature and structure are rotatable together, relative to the bracket and the stator, about the first axis of symmetry; the spherical armature, structure, and bracket are rotatable together, relative to the stator, about the second axis of symmetry; and the spherical armature, structure, bracket, and stator are rotatable together about the third axis of symmetry.

Abstract: A combined system for imaging and communication by laser signals including a telescope (10) in which the primary (1) and secondary (2) mirrors are shared between an imaging function and a function involving the emission of laser transmission signals. The accuracy required for the direction of emission (DE) of the laser transmission signals is obtained by additionally placing a geolocation device on board a satellite carrying the combined system. No coarse pointing assembly is used and it is also possible for the system to be devoid of any fine pointing assembly and target acquisition and tracking detector.

Abstract: Attitude and/or attitude rate of a vehicle may be controlled using jet paddles and/or movable masses. Thrust direction generally may also be controlled using jet paddles. The jet paddles may be moved into and/or sufficiently close to the exhaust flow, and out of the exhaust flow, to change the thrust direction. Movable masses may also be used in addition to, or in lieu of, jet paddles. Movement of the movable masses alters a center-of-mass of the vehicle, generating torque that changes vehicle attitude.

Abstract: A method for monitoring position of and controlling a second nanosatellite (NS) relative to a position of a first NS. Each of the first and second NSs has a rectangular or cubical configuration of independently activatable, current-carrying solenoids, each solenoid having an independent magnetic dipole moment vector, ?1 and ?2. A vector force F and a vector torque are expressed as linear or bilinear combinations of the first set and second set of magnetic moments, and a distance vector extending between the first and second NSs is estimated. Control equations are applied to estimate vectors, ?1 and ?2, required to move the NSs toward a desired NS configuration. This extends to control of N nanosatellites.

Abstract: An optical meter includes a force member to receive a force and a reflector disposed on the force member to receive radiation and to communicate a pressure of the radiation to the force member. The reflector includes a reflective surface, and the force member is configured to be displaced in response to receiving the force comprising the pressure. The optical meter is configured to measure a power of the radiation, an energy of the radiation, or a combination thereof based on the pressure. A process for measuring a property of radiation includes receiving radiation by the reflector, reflecting radiation from the reflective surface, communicating a pressure from the reflector to the force member, and displacing the force member.

Abstract: A method and system of using a common set of coils to provide at least two of magnetic flux, heat and degaussing in a mobile platform are provided. In accordance with one embodiment, the method involves oscillation of current in the coils at a frequency higher than a defined pointing requirement to provide heat. In accordance with another aspect, the coil functions as a degausser by energizing the coil with an oscillating current that decreases in amplitude over time.

Abstract: A method of controlling inertial attitude of an artificial satellite in order to perform a navigation function and to maximize terrestrial coverage of the Earth by the satellite. The method includes deploying the artificial satellite in an orbit about the poles of the Earth; applying gyroscopic precession to the artificial satellite spin axis to precess and maintain the satellite near the ecliptic pole; deploying the artificial satellite so that the spin axis is initially perpendicular to or substantially perpendicular to sun lines; and applying gyroscopic precession to the artificial satellite spin axis to precess the spin axis away from an initial deployed attitude at a selectively-variable precession rate and to maintain the spin axis perpendicular to or substantially perpendicular to the sun lines.

Abstract: Space tug vehicles and methods for providing space tugs and moving target vehicles are provided. More particularly, a space tug utilizing electrostatic or Coulomb force for acting on target vehicles and for moving the target vehicles into new orbits or altitudes are provided. The space tug may establish an attractive electrostatic force by controlling the electrical potential of the space tug so that it is opposite the electrical potential of a target vehicle. The target vehicle may acquire an absolute electrical potential due to its interaction with the space plasma and photoelectrons or the space tug may impart additional charge to the target vehicle. After establishing the attractive electrostatic force, a propulsion system of the space tug is operated to provide thrust. Thrust is directed to change the orbit and/or position of the target vehicle that is being pulled towards the space tug by the electrostatic force.

Abstract: An electromagnetic field-producing device is used for non-contact type high elevation, elevation maintenance, or manipulation of objects. This type of elevation, elevation maintenance, or manipulation of objects contributes to the present modes of flight and people and larger object transport including but not limited to rotary-wing or non-rotary-wing modes of flying and mechanical/electronic modes of object movement.

Abstract: A satellite comprises a main body and, in orbit, at least one free object floating in space including a portion of the payload transferred from the main body. The free object is restricted to a precise position and attitude relative to the main body. An electromagnetic force and an electromagnetic torque are produced for controlling the position and the attitude of the free object in accordance with a master-slave relationship.

Abstract: An spacecraft control system is provided that includes a tube and a plurality of microwheels. The plurality of microwheels is disposed within the tube. Each microwheel has a first stator wafer, a second stator wafer, and a rotor wafer disposed therebetween. The first and second stator wafers are configured to spin the rotor wafer.

Abstract: An apparatus and method that make use of electromagnetic energy to maneuver an object, such as stop, slow, and/or divert a vessel or projectile in low and zero-gravity environments. The apparatus comprises an element capable of generating a magnetic field in the zero or low-gravity environment, and an object capable of electromagnetically interacting with the magnetic field so that the object's speed and/or trajectory is altered when moving in proximity to the magnetic-field generating element. As such, the method entails maneuvering an object in a zero or low-gravity environment by generating a magnetic field in the zero or low-gravity environment, and then moving the object in proximity to the magnetic field such that the magnetic field alters the trajectory and/or speed of the object.

Abstract: The present invention improves the orbital maneuvering and power generation capabilities of a system of a satellite(s) connected with a conducting tether(s) by spinning the system about its mass center at an angular rate which is relatively high compared to the average orbital rate. An improvement in tether performance is achieved because at many times during rotation the tether is positioned at much better angles with the magnetic field and significantly higher currents are driven through the tether without destabilizing the system. The current can flow either in the direction of the EMF induced in the tether, or in the reverse direction, depending on the tether orientation with respect to the magnetic field and the mission goals. The reverse current is driven by the onboard power sources. Spinning electrodynamic tether systems can also be lighter and simpler in design and more flexible in operation.

Abstract: The present invention describes a communication system comprising a primary radio station (PS) and at least one secondary radio station (SS), intended to be in motion (MOT). The secondary radio station has at least one controllable structure (CS) for communicating with the primary radio station, and control means (CONT) for controlling the controllable structure depending on the movements of the secondary radio station. The control means of the controllable structure comprise magnetic field sensors (MFS) and gravitational field sensors (GFS) for providing measurements of the earth magnetic (H) and gravitational (G) fields, and computing means (COMP) for computing control information from these measurements. The computing means read the outputs of each sensor and make the calculations required to control the controllable structure at appropriate time intervals depending on the motion state of the secondary radio station.

Abstract: For controlling the attitude of a satellite placed on a low earth orbit, components of a vector Bm of the earth's magnetic field along three measurement axes of a frame of reference bound with the satellite (typically by means of a three-axis magnetometer) are measured. The orientation of the earth's magnetic field in the frame of reference is computed and a derivative {dot over (B)}m of the vector is also computed. Magneto-couplers carried by the satellite are energized to create a torque for spinning the satellite at an angular frequency &ohgr;c about a determined spin axis of the satellite, where &ohgr;c is greater than an orbital angular frequency 2&ohgr;0 of the satellite.

Abstract: The invention relates to a combined propulsion system for a spacecraft, combining an electric propulsion system and a tether propulsion system, characterized in that the electric propulsion system is an electrostatic system (EP) which has, on the one hand, a generator (G) for ejecting an ionized plasma, comprising at least one grid for electrostatically accelerating the ions in the plasma and, on the other hand, an electron generator (NM) generating electrons to maintain the electrical neutrality of the system, and in that the electron generator (NM) is coupled to a first end of the tether which is the opposite end to the end at which the plasma-ejection generator (G) is arranged, generating a compensating current (Ic) flowing along said tether.

Abstract: A three-axis attitude control for a low-orbiting satellite. A sensor which measures in two axes is a magnetometer, and a satellite is positioned on an inclined orbit at an inclination angle of approximately 25 degrees to approx. 90 degrees. The satellite has been provided with an overall spin by spin wheels, and actuators generate torque on all three axes of the satellite. Accordingly, this is applicable to three-axis attitude control of a low-orbiting satellite that includes a sensor which measures in two axes and that includes one or several spin wheels.

Abstract: This invention relates to a spacecraft which generates its own magnetic moment and magnetic field gradient in order to produce lift on the hull. The magnetic moment is generated by a large area solenoid located in the hull. A toroidal core wrapped with electrical bobbins at intervals along said core produces a traveling magnetic wave along its surface. This magnetic wave creates a spacetime curvature, similar to a tilted plate, which causes the formation of a magnetic field gradient. Power is not critical because the system uses a magnetic vortex wormhole generator to lower the speed of light in order to efficiently create highly relativistic fields due to Lorentz transformation.

Abstract: For controlling the attitude of a satellite placed on a low earth orbit, components of a vector Bm of the earth's magnetic field along three measurement axes of a frame of reference bound with the satellite (typically by means of a three-axis magnetometer)are measured. The orientation of the earth's magnetic field in the frame of reference is computed and a derivative Bm of the vector is also computed. Magneto-couplers carried by the satellite are energized to create a torque for spinning the satellite at an angular frequency &ohgr;c about a determined spin axis of the satellite, where &ohgr;c is greater than an orbital angular frequency 2&ohgr;0 of the satellite.

Abstract: Spacecraft maneuvers (e.g., orbit transferring, stationkeeping and attitude controlling) of a first spacecraft are realized with conventional force and torque generators (e.g., thrusters and momentum wheels). Spacecraft maneuvers of a second spacecraft are realized through magnetic interaction between the first and second spacecraft using a closed loop control system. In particular, a magnetic moment vector m1 of a magnetic system of the first spacecraft and a magnetic moment vector m2 of a magnetic system of the second spacecraft are adjusted to apply selected force vectors and torque vectors to the first and second spacecraft using a closed loop control system.

Abstract: Spacecraft maneuvers (e.g., orbit transferring, stationkeeping and attitude controlling) of a first spacecraft are realized with conventional force and torque generators (e.g., thrusters and momentum wheels). Spacecraft maneuvers of a second spacecraft are realized through magnetic interaction between the first and second spacecraft using a closed loop control system. In particular, a magnetic moment vector m1 of a magnetic system of the first spacecraft and a magnetic moment vector m2 of a magnetic system of the second spacecraft are adjusted to apply selected force vectors and torque vectors to the first and second spacecraft using a closed loop control system.

Abstract: An attitude control system uses a first order partial derivative of an error signal, representing a difference between a commanded pitch attitude and a sensed, actual pitch attitude. The first order partial derivative forms a cross product with the Earth's magnetic field to provide a X axis torque rod dipole command, which powers the torque rod in a way that causes a satellite to move toward its commanded attitude.

Abstract: A rotation stabilizing device in a microgravitational rotating apparatus is provided by which vibration is suppressed and rotation is stabilized. Recess portions 10a, 10b are provided on a casing 10 so that bearings 14, 15 are arranged therein for supporting a rotary shaft 30. Experimental boxes 17a to 17h fixed to arms 16a to 16b rotate together with the rotary shaft 30. A fin 33 is fixed to outer circumferential surfaces of the experimental boxes 17a to 17b projecting outwardly therefrom. Pairs of electromagnetic coils 31a, 31b, in which the fin 33 is interposed with a predetermined gap maintained therebetween, and gap sensors 32 close thereto are fitted on the casing 10 side. Displacement of the fin 33 is detected by the gap sensors 32 to be inputted into a control unit to thereby control exciting current of the electromagnetic coils 31a, 31b. The gap is brought into the position within a demand value and rotation is stabilized.

Abstract: The present invention comprises apparatus and methods for using and controlling electrodynamic tethers. The apparatus taught by the present invention uses an interconnected multiwire (compared to the long, narrow single wires of the prior art) conductive tether whose area maximizes electrodynamic drag while simultaneously minimizing the Area-Time-Product swept by the tether during its operating life. The preferred tether length is two kilometers to five kilometers. The preferred tether mass is one percent to five percent of the spacecraft mass. The methods of control taught by the present invention comprise orienting the tether structure at a preferred angle to the local vertical to maximize electrodynamic drag and minimize tether instabilities. The angle of 35.26 degrees to the local vertical line is preferred.

Abstract: Orientation system and process of an artificial satellite in low orbit. Aerodynamic control surfaces producing rotations of the satellite by upper-layer atmosphere are combined along two axes to magnetic couplers (1, 2, 3) producing rotating movements along the three principal axes of pitch, roll and yaw thanks to the Earth's magnetic field. Great precision may be obtained in orienting the satellite without substantial energy consumption. Command laws are provided.

Abstract: A magnetic torquer (10) for spacecraft attitude control, providing a torque that is controllable over an extended range greater than a normally used linear range, and thereby providing a higher ratio of maximum torque to device weight. In one form of the torquer, a feedback loop including a sensor (20), a signal subtraction circuit (24), a controller (28) and an adjustable power supply (16), continuously determines a corrected command signal that results in the generation of a desired magnetic moment and a corresponding torque, over the extended range, regardless of non-linearity across the extended range. In an alternate form of the torquer, the corrected command signal is generated from a mathematical model (56) of the magnetic torquer (10). Use of the invention also minimizes the effect of any residual magnetic moment that might by present when the actuating current is reduced to zero.

Abstract: A method for use on a satellite that automatically inhibits scanning of an Earth sensor to handle sensor intrusions by the sun, moon, or other celestial bodies. In implementing the method, a predicted state vector for the satellite, derived from an orbit propagator, is generated. An attitude profile for the satellite is generated. Then, the satellite state (predicted state vector) and attitude profile are processed to determine Earth, sun, and moon vectors in a satellite body frame of reference at any instant. The Earth, sun, and moon vectors are compared to the Earth sensor field of view and sensor field of view limit boxes to determine if scan inhibiting or Earth sensor switching should occur. The affected Earth sensor is inhibited or switched if an intrusion of the sun and/or moon into the field of view of the Earth sensor is predicted.

Abstract: Apparatus and a method is presented for controlling yaw and roll excursions in a spacecraft having on-board components such as magnetic torquers, a roll thruster, momentum wheels, a wheel controller, an earth sensor, and a Digital Integrating Rate Assembly (DIRA). An observer module on the spacecraft receives inputs containing information comprising the unbiased roll error from the earth sensor, yaw momentum measured from the wheel speeds, and commanded yaw and pitch momentum output from the wheel controller, and produces therefrom output signals indicative of yaw estimate, yaw momentum estimate and torque disturbances estimate.

Abstract: A smart docking surface consisting of closely spaced cantilevered sensor/actuator structures capable of precisely repositioning an object having a ferro-magnetic surface in contact with the smart docking surface. It is designed for use in a micro gravity environment for the final docking sequence of two small (<100 kg) satellites. Its purpose is to reduce the complexity of the docking process where a precise mating is required.

Abstract: A pseudo gyro emulates mechanical gyros by software processes by processing space system appendage measurement data and reaction wheel tachometer data within reference and control systems of a satellite using principals of conservation of momentum to compute the vehicular bus angular velocity rate data by accounting for external torque, the momentum transfer between the satellite, the bus, and the appendages for providing accurate relative vehicular position and angular velocity rate data as an integral part of attitude reference and control systems now having higher reliability, longer life times, lower power consumption, and more accurate vehicular angular velocity rate data generated within high bandwidth operations. A specific improvement is the use of a bias torque estimator for accounting for uncertainties in external torque estimations for accurate total system momentum computations.

Abstract: A method and apparatus are provided that effect spacecraft wheel desaturation by controlling the magnetic torquers aboard the spacecraft. An attitude control system energizes a magnetic torquer coil or set of magnetic torquer coils using a power source having an initial polarity and then monitors the spacecraft's response to the torque created by the coil's magnetic field. By measuring the spacecraft's response to the torque created by the energized magnetic torquers, the direction of the earth's magnetic field is determined. Once the direction of the earth's magnetic field is known, the proper polarity may be applied to the coil to achieve the desired torque. The control system of the present invention synchronously demodulates the craft's angular velocity, estimated from signals obtained from the attitude sensors, using a small dither signal that is superimposed on the control current used to energize the magnetic torquer coil.

Abstract: A torque/reactive momentum wheel control system for use in satellites for dynamic attitude maintenance and alteration where the flywheel of each momentum wheel is levitated by a high-temperature superconducting element repulsively interacting with permanent magnets in the flywheel. The spin rate (rpm) of the flywheel being controlled by either an active magneto or electromagneto drive system. Each momentum wheel is cooled by a cryo-cooler and can have a total weight of about 10 Kg to a fraction of 1 Kg and delivering 3.5 Js with less than 1 W loss.

Abstract: A navigation system for spinning projectiles using a magnetic spin sensor to measure the projectile roll angle by sensing changes in magnetic flux as the projectile rotates through the earth's magnetic field is disclosed. The magnetic spin sensor measurements are used to despin a body reference frame such that position, velocity, and attitude of the projectile can be determined by using a strapdown inertial navigation system (INS) algorithm. More particularly, a multisensor concept is used to measure pitch and yaw angular rates, by measuring Coriolis acceleration along the roll axis and demodulating the pitch and yaw rates therefrom.

Abstract: A spacecraft attitude control system uses at least four reaction wheels In order to minimize reaction wheel speed and therefore power, a wheel speed control means system is provided. The wheel speed control means monitors the wheel speeds and controls wheel speed nullspace components to keep the wheel speeds as small as possible.

Abstract: A method for controlling orbiting spacecraft uploads a plurality of parameters to an onboard processor, the parameters including a plurality of quaternion elements describing an inertial to orbit reference frame, and including radial position, radial velocity and angular rate of the spacecraft. Parameters are input for processing one of a plurality of propagation algorithms, preferably calculating said propagation algorithms with the uploaded parameters, computing enhanced time of day vectors, and computing ground time of day angles from the enhanced time of day vectors for adjusting the payload pointing angle, such as an antenna pointing angle, of the satellite. Preferably, the calculating step comprises integrating an orbital state vector with a second order Runge-Kutta integration algorithm. In addition, enhanced time of day fault processing provides an internal check on integration accuracy and is checked every time of day update.

Abstract: Control moment gyros in an array are rotated to reorient a satellite. Gyro angle is selected from one of three values based upon whether the stored angular momentum for the desired angle is below a singularity free value, greater than the singularity free value or is greater than the maximum available stored angular momentum.

Abstract: The present invention comprises an electrodynamic tether structure and a method of use. The structure of the tether taught by the present invention is a short, wide, interconnected multiwire (compared to the long, narrow single wires of the prior art) conductive tether whose area maximizes electrodynamic drag while simultaneously minimizing the Area-Time-Product swept by the tether during its operating life. The preferred tether length is two kilometers to five kilometers. The preferred tether mass is one percent to five percent of the spacecraft mass. The method of operation comprises orienting the tether structure at an angle to the local vertical to maximize electrodynamic drag on the host spacecraft and minimize tether instability. The angle of 35.26 degrees is preferred.

Abstract: The attitude of a physical platform (62) is altered by a summation of forces from an attitude actuator (74), disturbance forces (80) and random noise (82). Attitude sensors (64) sense the attitude of the physical platform (62) and output measured attitude parameters to a disturbance force effects (DFE) filter (66) and an attitude filter (68). The DFE filter (66) outputs an attitude perturbation command to the attitude filter (68) and a command translator (72). Effects of the attitude perturbation are used by the DFE filter (66) to estimate the effects of disturbances on the physical platform (62). A trigger signal (76) from the attitude actuator (74) is used by DFE filter (66) to select an appropriate gain matrix used in the estimation process.

Abstract: A method and apparatus for estimating attitude sensor bias in a satellite system uses attitude sensors and a spacecraft control processor. Attitude sensors provide output signals, which may contain bias. The present invention interprets the signals, determines the bias present in the signals, and generates an output signal to offset the bias in the signals from the attitude sensors, thereby leading to more accurate positioning of the satellite employing the present invention.

Abstract: A spacecraft (10) body (20) located in an external magnetic field (12) has a substantially unbalanced electrical power bus (18) that generates a composite local magnetic field The electrical power bus (18) substantially surrounds an external surface (22) of the spacecraft (10). Current in the electrical power bus (18) is solar panel (14and 16) and/or battery (34) generated and can flow in different directions through different current paths. Electrical loads (58) are coupled to power converters (56) and current path switches (54) to couple the power converters (56) to the various current paths. A controller (28) takes sensor data (82), determines a required attitude adjustment torque (84), translates that adjustment torque into a composite magnetic field (86), determines an unbalanced bus current profile (88) to generate that composite magnetic field, selects a current source (90) and regulates current path switches (92) to achieve a desired attitude adjustment torque for the spacecraft (10).

Abstract: Spacecraft maneuvers (e.g., orbit transferring, stationkeeping and attitude controlling) of a first spacecraft are realized with conventional force and torque generators (e.g., thrusters and momentum wheels). Spacecraft maneuvers of a second spacecraft are realized through magnetic interaction between the first and second spacecraft. In particular, a magnetic moment vector m.sub.1 of a magnetic system of the first spacecraft and a magnetic moment vector m.sub.2 of a magnetic system of the second spacecraft are adjusted to apply selected force vectors and torque vectors to the first and second spacecraft.

Abstract: Attitude-control methods are provided which eliminate or reduce structural resonances in spacecraft that are excited by momentum wheel imperfections. The methods are preferably practiced with an over-determined attitude-control system which is typically available because spacecraft often carry backup momentum wheels to insure against failure of primary momentum wheels. In a method embodiment, a set of angular-velocity waveforms is selected that substantially realizes a commanded momentum vector when applied to a corresponding set of momentum wheels wherein none of the selected angular-velocity waveforms dwells at a resonant frequency of the spacecraft. The angular velocity of the corresponding set of momentum wheels is then conformed to the selected set of angular-velocity waveforms to realize a spacecraft attitude that corresponds to the commanded momentum vector without exciting the resonance.

Abstract: A pseudo gyro emulates mechanical gyros by software processes by processing space system appendage measurement data and reaction wheel tachometer data within reference and control systems of a satellite using principals of conservation of momentum to compute the vehicular bus angular velocity rate data by accounting for the momentum transfer between the satellite, the bus, and the appendages for providing accurate relative vehicular position and angular velocity rate data as an integral part of attitude reference and control systems now having higher reliability, longer life times, lower power consumption, and more accurate vehicular angular velocity rate data generated within high bandwidth operations.

Abstract: A low earth orbiting satellite control arrangement according to the invention has an earth sensor for measuring roll and pitch of the satellite relative to an earth oriented moving reference system, a spin wheel mounted along an axis perpendicular to the orbital plane of the satellite for measuring spin of the satellite, and two magnet coils for generating control. A first of the two magnet coils is aligned substantially in parallel to a roll axis of the satellite, and the second is aligned substantially parallel to the yaw axis. Alternative control algorithms are disclosed for controlling attitude, nutation and spin of the satellite based on information from the earth sensor, spin wheel and an observer which is also mounted on the satellite.

Abstract: An intriguing and educational display structure (1) appears to be spinning upon itself without any apparent drive mechanism, power supply or bearing. The structure comprises two concentric hollow spheres (3, 5) spaced apart by a transparent fluid (6). The outer sphere is totally transparent, immobile, and may rest on a tripod (2) or other type of support. The inner sphere may be partially transparent or translucent and carries over its surface, a design such as a map of the world (7). The inner sphere rotates within, and independently from, the first outer one. The internal, that is self-contained, drive mechanism uses either a conventional electrical motor (8) with its own field winding and commutator, or one made of a rotor comprising a cross-array of electromagnets (AC, BD) that interact with an ambient magnetic field such as the earth magnetic field.