Abstract: The present invention relates to an elastic metamaterial for reducing vibrations of a flexible structure such as a main cable of a tether system for controlling an orbit of a satellite revolving around a planet, and a method for improving a vibration reduction performance thereof, and more particularly, to an elastic metamaterial having an improved precision, in which a ratio of a cross-sectional area of a pendulum ring may be adjusted to maintain a frequency characteristic other than a band gap generated due to the elastic metamaterial even in a state where a mass of the pendulum ring is not changed, and a band gap (R_ring) generated due to the pendulum ring of the elastic metamaterial and a band gap (R_beam) generated due to the elastic beams may be combined into one band gap to expand a vibration damping range, and a method for improving a vibration reduction performance thereof.

Abstract: A method for demonstrating a new scientific discovery made by the inventor about the nonlinear instability of vehicles, like aircrafts, automobiles and ocean vehicles. Said method includes a model and a three-gimbaled framework that permits said model to respond to inertial moments about the axes of which the moments of inertias are the smallest and the largest, wherein said model has restoring and damping capabilities along these two axes. Said method also comprises how to use a variable motor or a crank for controlling said model rotational motions about the intermediate principal axis of inertia with closed form formulas for the external driven frequencies and amplitudes to be used to excite the nonlinear instabilities of said model. Said model could be an aircraft, an automobile, a ship, or even a rectangular block.

Abstract: Techniques for performing an orbital maneuver on a spacecraft by firing a thruster include executing a first double pulse, double coast firing (DPDC) sequence with the thruster prior to the orbital maneuver, firing the thruster for a duration of the orbital maneuver, and executing a second DPDC firing sequence with the thruster subsequent to the orbital maneuver.

Abstract: An electrodynamic machine includes a rotor assembly, a stator assembly defining an annular core receiving the rotor assembly, the rotor assembly rotating within the stator assembly based on electromagnetic fields generated by the stator assembly and the rotor assembly, a plurality of structural components mechanically supporting the rotor assembly and stator assembly, and an active damping element providing forced vibrations that counteract structural vibrations of the electrodynamic machine caused by magnetic forces based upon the electromagnetic fields or caused by mechanical forces based upon the plurality of structural components.

Abstract: A magnetic bearing apparatus for supporting a rotational shaft includes a magnetic bearing, a rotation sensor, a controller which controls an electric current flown to an electromagnet, and at least one vibration sensor. The controller controls the electric current flown to the electromagnet as follows: in a first control mode, a first electric current is flown to the electromagnet for levitating and supporting the rotational shaft; and in a second control mode, a second electric current is flown to the electromagnet in place of the first electric current at a first periodic timing which is detected by the rotation sensor so that vibration is reduced to a lower level than a first vibration value detected in the first control mode in at least one of setting location of the at least one vibration sensor.

Abstract: A device used for providing dynamic isolation and damping of dynamic vibrations, in a passive way, originated in the launch vehicle of a space shuttle and reaching the payload or satellite. The device comprises a plurality of identical elementary unit elements, such that the device is designed in a modular way, allowing the individual modularity of each of the elementary unit elements. Each of the elementary unit elements is tailored and designed individually, and the complete device can be designed for each particular application and payload needed as a function of each of the elementary unit elements allowing an easy design and lower costs, for a wide range of payload applications. Each elementary unit element comprises a spring component and a damping component, such that the functionalities provided for each component are separated and can be individually tailored, thus providing a device having a wider range of adaptation capabilities.

Abstract: A system for damping nutation and removing wobble of a spacecraft spinning about a given axis is provided. The system includes a sensor configured to determine three dimensional attitude measurements of the spacecraft, a processor operatively coupled to the sensor and configured to execute a process that facilitates aligning the spin axis with a spacecraft momentum vector. The processor, when executing the process, is programmed to receive spacecraft attitude data from the sensor, determine a torque command using the received attitude data, and control a momentum storage actuator on the spacecraft using the determined torque command such that an angular deviation about the given axis is reduced.

Abstract: The aim of the method of controlling at least one mechanical system exhibiting at least one flexible structure element and at least one actuator or group of actuators is to reduce the level of vibrations of said element while controlling the actuator or group of actuators in such a way as to achieve at least one objective assigned to the mechanical system. The control consists of increments having to be carried out at a sampling period T, each of the increments being constant in amplitude in a sampling time interval extending between two successive sampling instants and each of the increments being applied during a duration which is less than or equal to the sampling period T. In each sampling interval, at least one of the initial and final instants of application of the increment is modified by adding a variable, random or pseudo-random, temporal deviation dT.

Abstract: A system for controlling an aerospace vehicle using on-line inertia estimation may include an attitude sensor to measure an attitude of the aerospace vehicle. The system may also include a processor on board the aerospace vehicle. An inertia estimator operable on the processor may generate an on-line inertia estimate of the aerospace vehicle. A rate and attitude estimator operable on the processor may determine an angular position and angular velocity of the aerospace vehicle using the attitude measurement of the aerospace vehicle and the on-line inertia estimate for controlling movement and orientation of the aerospace vehicle without any rates of rotation of the aerospace vehicle being required.

Abstract: Methods and apparatus are provided for active motion damping of a spacecraft whose attitude control system (SACS) has experienced a partial failure. In a preferred embodiment, the apparatus comprises, an attitude control processor (ACP), a motion damping controller (MDC) coupled to the ACP, a control moment gyro (CMG) comprising a gimbal loop controller (GLC) coupled to the MDC and a gimbal motor with input coupled to the GLC and gyroscopic torque output coupled to the satellite. Normally, a gimbal rate command (GRC) from the ACP passes to the GLC for execution to rotate the CMG gimbal and the coupled satellite. If the ACP output becomes invalid, the MDC assumes control and can modify CMG operation in several ways to safely decelerate a moving satellite without hazardous mechanical stress. Following such active motion damping, the CMG remains quiescent until a RESET is received whereby further valid GRCs are then executable.

Abstract: A system for damping nutation and removing wobble of a spacecraft spinning about a given axis is provided. She system includes a sensor configured to determine three dimensional attitude measurements of the spacecraft, a processor operatively coupled to the sensor and configured to execute a process that facilitates aligning the spin axis with a spacecraft momentum vector. The processor, when executing the process, is programmed to receive spacecraft attitude data from the sensor, determine a torque command using the received attitude data, and control a momentum storage actuator on the spacecraft using the determined torque command such that an angular deviation about the given axis is reduced.

Abstract: The invention provides a novel, low-cost, spin-stabilized lander architecture capable (with appropriate system scaling tailored to the attributes of the target) of performing a soft-landing on a solar-system body such as Earth's Moon, Mars, Venus, the moons of Mars, Jupiter, Saturn, Uranus and Neptune, selected near-Earth and main-belt asteroids, comets and Kuiper belt objects and even large human-made objects, and also moving about on the surface of the target solar-system body after the initial landing in movement akin to hopping.

Abstract: The invention concerns an observation satellite (1) which is intended to be placed in orbit around a celestial body (2), and which comprises a reflecting device (5), a receiving device (6), a linking mechanical system (19), the whole forming a capture system (3) which is suitable to be able to orient the capture system (3) by gravity gradient in an aiming position in which the electromagnetic radiation corresponding to the information to be captured is received. The invention extends to an observation method and a fleet of such satellites (1).

Abstract: An inventive autonomous active manoeuvring method and system 1 for spinning spacecraft is provided having a capability to enhance the AOCMS performance of passive spinning satellites and to fulfil the emerging autonomy requirements applicable to new generation satellites. In broad terms, the invention resides in (a) the overall concept of providing autonomous execution of spin axis re-orientation manoeuvring for spinning spacecraft designed and executed autonomously on-board the spacecraft by the AOCMS and (b) in the proposed strategy set in place to execute the re-orientation manoeuvres with respect to the handling of residual nutation. Advantageously, the provision of coupling nutation avoidance manoeuvres with active nutation damping 8 on board the spacecraft reduces/minimises the manoeuvre settling time required to return the spacecraft to the steady state pointing performance, while not imposing constraints upon the particular spacecraft inertia sensor properties.

Abstract: A pod (34) containing sensors and/or targeting systems such as laser designators or range finders are hung on a cable (30) below an aircraft (12). The line (30) can be reeled in or out similar to towed decoys. This allows the aircraft (12) to operate above or in the cloud cover while the sensors/targeting equipment (34) are operating below the cloud cover to find the enemy forces.

Abstract: Recovery of a towed body, in one embodiment in the form of a decoy which is initially stowed in a receptacle or canister and which is allowed to pay out behind an aircraft on a towing cable wrapped around a spindle, is accomplished by snaring or lassoing a portion of the towing cable and by dragging it to a further spindle which is driven so as to cause the lasso and a portion of the towing cable to wind up around the driven spindle. When a sufficient amount of the cable is wound around the driven spindle, the cable end secured to the canister is severed to allow all of the rest of the towing cable to be wound up. In one embodiment, a telescoping saddle or docking cradle is provided which extends from the canister to receive the retracted towed body so that it may be secured to the moving vehicle from whence it can be recovered, refurbished, and redeployed.

Abstract: A power plant for propeller aircraft has two propulsion devices, each having a propeller, an engine for powering the propeller, and a gear transmission interposed between the engine and the propeller. The transmission has a power input shaft connected directly to the engine, and a power output shaft supporting the propeller and rotated with respect to the input shaft about a hinge axis perpendicular to the output shaft. The gear transmission also has a single gear fitted to the output shaft and having two opposed sets of teeth meshing with a pair of pinions. One of the pinions extends coaxially with the hinge axis, and is connected to the input shaft by a single pair of gears.

Abstract: A satellite attitude control system includes a programmed processor system which includes a gyroscopic actuator first control stage for changing the attitude of the satellite and a reaction wheel second control stage for assuring that pointing of the satellite is accurate and stable. The method is intended to be used for a satellite including the two control stages indicated above, which it uses selectively for the operations indicated above.

Abstract: A pod (34) containing sensors and/or targeting systems such as laser designators or range finders are hung on a cable (30) below an aircraft (12). The line (30) can be reeled in or out similar to towed decoys. This allows the aircraft (12) to operate above or in the cloud cover while the sensors/targeting equipment (34) are operating below the cloud cover to find the enemy forces.

Abstract: An automatic, on-board system for orbiting spacecraft that controls and reduces short-term transients caused by roll momentum unloads by manipulating the states of the nutation damper through resetting one of the states in the polynomial transfer function to a value in proportion to the magnitude of the expected roll momentum unload and creating a transient equal and opposite in magnitude to that expected by the roll unload. The compensator will, of its own accord, automatically damp out the resulting transient. Thus the short-term transient is reduced without biasing the spacecraft in roll. No angle bias need be changed or removed.

Abstract: A method for estimating motion of a dual-spin spacecraft having a gimballed momentum wheel. The method disengages the gimbal from its drive train in anticipation of a short disturbance. Then, the gimbal slip resulting from the disturbance is measured. The method may be used to stabilize any dual-spin spacecraft that uses a gimballed momentum wheel. To stabilize the dual-spin spacecraft, the gimbal angle and gimbal rate are measured during and after the disturbance to provide an indication of the inertial spacecraft motion along gimbal axes. The magnitude and direction of the disturbance are determined by comparing motion of the gimbal before and after the disturbance. Then, torques are applied to the gimbal to counteract the spacecraft motion resulting from the disturbance.

Abstract: The excitations from the engines of launch vehicles and the aerodynamics of flight produce large vibrations which are highly detrimental to spacecraft during launch. Significant dynamic loads often exist in all three translations and for rotations as well, therefore, complete six degree-of-freedom vibration load isolation is often desired. The disclosed device utilizes a previously disclosed passive axial vibration isolation device to simply and effectively create a passive three-axis vibration isolation device suitable for effecting a six degree-of-freedom whole-spacecraft passive vibration isolation system. The vibration isolation system design can be simply tuned to address various dynamic load frequency bands of concern, including isolation for pyrotechnic shock. The resulting system is compact and lightweight and can be easily utilized with existing launch vehicle-to-payload support structures.

Abstract: A star tracker coupled to the spacecraft having a star catalog associated therewith. A sun sensor is coupled to the spacecraft. A control processor is coupled to the star tracker and the sun sensor. The processor obtains star data using a star tracker and an on-board star catalog. The processor generates a coarse attitude of the spacecraft as a function of the star data, and establishes a track on at least one star in the on-board star catalog. The processor calculates a sun tracking rate, and obtains a normal phase attitude as a function of the star data and the coarse attitude. The information is used to slew the spacecraft to a desired attitude.

Abstract: Excitations from engines of the launch vehicles and the aerodynamics of flight produce large vibrations which are very detrimental to spacecraft during launch. Whole-spacecraft vibration isolation systems that greatly reduce the transmitted structure-borne vibrations to the spacecraft are disclosed. The disclosed devices, when combined together into an isolation system, provide vibration isolation for the complete spacecraft and are the only connection between the spacecraft and the launch vehicle. Devices that provide isolation in the axial (thrust) direction are disclosed. These devices consist of an arrangement of damped metallic flexures and machined springs, which provide relative flexibility in the desired direction. Passive damping is provided by compressing and shearing viscoelastic materials.

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: A method of damping nutation and removing wobble of a spacecraft provided with first and second momentum sources having linearly independent momentum components in the plane transverse to the given axis. The method includes sensing angular velocities of the spacecraft along orthogonal first and second axes in the transverse plane, and generating velocity signals representative of the angular velocities. The velocity signals are processed to form control signals representative of desired torques for driving the momentum sources. The desired torques have a first additive component proportional to the angular velocities for removing the wobble and a second additive component proportional to the first derivatives of the angular velocities for damping the nutation.

Abstract: One or more deployable thermal panels (24, 28, 70) are actively controlled throughout the orbit of a satellite (20) to provide thermal dissipation. Adjusting the incident angle between the panel (24, 28, 70) and the sun and controlling the flow of fluid through optional flexible heat pipes loads and unloads heat to provide thermal stability for components (62) which have special thermal requirements. An optional antenna panel (70) on a nadir side (64) of the satellite (20) offers an antenna side (74) on one surface and a thermal radiating side (72) on an opposing surface. In addition, thermal panel movements are controlled (96) to provide counter-disturbance torques (140).

Abstract: A feedback motion compensation (FMC) component for use in a spacecraft having a payload. The FMC component includes a controller having sensor data inputs and an open loop dynamics model that is driven by actuator commands to produce a high accuracy attitude estimate suitable for payload motion compensation.

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: In a satellite (20), two-axis solar panel drives (56, 58) allow two degrees of rotational freedom of movement for solar panels (22, 24). Regulation of electrical current loops (46, 48) within the solar panels (22, 24) generates a magnetic field. When the solar panels (22, 24) are tilted about an overturning axis (28), the magnetic field provides a component of a control torque about a pitch axis (30). An on-board processor (52) commands the panel drives (56, 58) and a magnetic torque actuator (60) to instantaneously and simultaneously damp disturbance torques about a windmill axis (26), an overturning axis (28), and a pitch axis (30). The processor (52) allows disturbance torque to be corrected continuously rather than periodically over the course of an orbit. Counter-disturbance torques (90) are applied (88) in synchronism with a satellite nutation frequency (76) but out of phase with the nutation so that nutation is damped.

Abstract: A method for simultaneously controlling satellite nutation, sun angle and attitude walk using a dual-slit sun sensor and thrusters. The sun angle and nutation angles are computed from sun sensor data, and the thrusters are fired when their use improves both sun angle and nutation. Precession around the sun line is controlled by constraining firing, in the simplest case, firing only when the angular momentum is precessed directly toward or away from the sun. Precession maneuvers are implemented by commanding a new sun angle and precession plane spin phase.

Abstract: Methods for performing attitude maneuvers of a spinning satellite without the use of thrusters, or with a minimal number of thrusters, are disclosed. The attitude maneuvers are primarily achieved through the use of gimballed momentum wheels and solar wing drives. Various maneuvers can be performed depending on whether the satellite has near-zero net momentum or significant net momentum. The maneuvers include sun acquisition, sun hold, Earth acquisition and inversion.

Abstract: The invention provides a method for controlling the precession of a spinning spacecraft (20) which allows the spacecraft body to respond to an input torque without the nutation normally attendant when an input torque is applied about one transverse axis to accelerate a spinning spacecraft about that one axis. Dynamic decoupling eliminates nutation through the impression of additional derived feedback torques (44,46) to the input torque control of a spinning spacecraft to oppose or cancel the intrinsic cross-coupling terms (34,36) of the spinning spacecraft's gyrodynamics that give rise to the nutation.

Abstract: A method and system of damping nutation of a spacecraft (20) having a desired spin axis along a first principal inertia axis utilizes a momentum source (28) oriented along a second principal inertia axis perpendicular to the first principal inertia axis. An angular rate of the spacecraft (20) is sensed along an axis transverse to both the first principal inertia axis and the second principal inertia axis. An angular rate signal representative of the angular rate is generated. The angular rate signal is processed to form a control signal representative of a desired torque to drive the momentum source. The desired torque has a first additive component proportional to a derivative of the angular rate to critically damp the nutation under an at least second order model of the spacecraft (20). The momentum source (28) is driven in dependence upon the control signal.

Abstract: An automatic, on-board system for orbiting spacecraft that controls yaw excursions caused by solar torques and thruster firings, which system combines inputs indicative of roll and yaw momentum increases and inputs containing information comprising the unbiased roll error from the earth sensor, yaw momentum measured from the wheel speeds, and commanded yaw momentum output from the wheel controller, and produces therefrom output signals indicative of the yaw estimate and the yaw momentum estimate. These output signals are combined and processed in a controller with a mimimum yaw error and roll thrust yaw controller gain and a miminmum yaw error and roll thrust yaw momentum controller gain, and a signal is produced therefrom for commanding roll thruster firings to change roll momentum, and, in turn, control yaw attitude and yaw excursions.

Abstract: The invention relates to a method for determining the angular momentum vector of a satellite, located in an external magnetic field, relative to a satellite-integral orthogonal system of coordinates. The satellite has a torque generating system to generate pulse-shaped torques around all three axes. A magnetometer measuring on three axes is used, whose measured values are fed to a computer. In the latter, at successive points in time and using the current measured values, an observer equation is integrated, with new estimated values being determined for the angular momentum vector. Using these as well as the current measured values, an inequality is continuously checked; when it is not fulfilled, the torque generating system operates briefly, so that finally the fulfillment of the inequality is forced and identity of the last estimated value that has been determined with the rotary pulse vector to be determined is achieved.

Abstract: The invention provides an improved process and apparatus for controlling the attitude of a three-axis stabilized spacecraft. Separate dead zone elements are used to define threshold values for limiting nutation amplitude and deviation of the spin direction of the spacecraft from a desired orientation, and a continuous check is performed to determine whether the nutation amplitude and deviation of the spacecraft exceed the threshold values. A control intervention is initiated to reduce the nutation, even when the spin direction is within the range limited by the applicable threshold value.

Abstract: A dynamic balance mechanism for balance control of a gyro stabilized (spinning) satellite is disclosed. An elongated gear rack is attached to the spacecraft. A movable mass is mounted by guide rollers on the gear rack and translates along the gear rack according to requisite electronic commands. The movable mass includes a housing, a stepper motor, a rotary potentiometer, a cable reel, a pair of gear heads and a pinion gear. The pinion gear meshes with the rack and is driven by the stepper motor. The potentiometer measures the position of the movable mass on the gear rack. The cable reel saves space and minimizes harness jamming conditions. The invention secures better weight efficiency.

Abstract: For using, on a spin-stabilized satellite, an actuator having an active axis substantially parallel to the spin axis wherein the actuator is activated in a predetermined mode and the resulting nutation period is measured. The actuator is activated in this predetermined mode for at least part of a preliminary step whose duration is equal to a predetermined fraction of the nutation period. This brings the satellite into a configuration in which triggering the service mode of operation induces minimal nutation. The actuator is then activated in this service mode of operation for any predetermined duration. The actuator is then activated under conditions which are the reverse of those of the preliminary step.

Abstract: A spacecraft uses monopropellant arcjets for velocity change such as for north-south stationkeeping. It has been discovered that, while an arcjet cannot be modulated by pulsing the fuel supply, the amount of thrust can be varied by modulating the power applied to the arc, without extinguishing the arc. While the specific impulse (I.sub.SP) of the arcjet is thereby reduced from the maximum I.sub.SP of which the arcjet is capable, the resulting I.sub.SP may still be larger than the combined I.sub.SP of an unmodulated arcjet in conjunction with a modulated chemical thruster in a typical scenario. According to the invention, attitude control is provided in conjunction with north-south stationkeeping or other velocity change by, in response to an error signal generated by an attitude control system, modulating the arc power(s) of an arcjet thruster(s), which provides the velocity change. The arc is not extinguished during the stationkeeping maneuver, but is varied in magnitude.

Abstract: A control system and method for causing a spacecraft (10) initially spinning about a principal axis of inertia (H) to transition to spin about an arbitrary command axis (3) is presented in which the capture maneuver is performed in a way to preserve spacecraft attitude knowledge without sensor or attitude propagation by using a simple rate feedback loop. The capture is accomplished by applying a step torque about axes (1) and (2) transverse to the desired spin axis (3) creating an initial nutation. The nutation is subsequently damped actively by closing a rate feedback loop with a low pass filter and applying transverse torques proportional to the transverse nutational rate.

Abstract: An improved method for transitioning a spacecraft from a thruster (5) controlled, station keeping mode to an operational, on-orbit mode, in which attitude control is maintained using internal momentum wheels (3). The method first utilizes a thruster compensation technique to supplement the conventional transition mode control system. This compensation uses empirical thruster data to derate thruster (5) efficiency for extremely short pulse durations necessarily used in transition mode operation. Secondly, a three step deadbeat thruster (5) sequence is employed in which a first pulse linearizes spacecraft (1) motion prior to orienting and stopping nutation. This linearization provides improved accuracy in computing and shortens the waiting time for subsequent thruster (5) firings.

Abstract: A spacecraft attitude control system uses one or more momentum or reaction wheels. Wheel bearing viscous (velocity-dependent) friction reduces the actual torque imparted to the spacecraft in response to a torque command signal. Friction compensation is provided by applying the torque command signal to a model of an ideal, friction-free wheel, and calculating the speed which the ideal wheel achieves in response to the torque command. An error signal is generated from the difference between the ideal wheel speed and the actual wheel speed. The error signal is summed with the torque command signal to produce the wheel drive signal. This results in a closed-loop feedback system in which the actual wheel speed tends toward the ideal wheel speed, thereby causing a torque on the spacecraft which is substantially equal to that commanded.

Abstract: A stabilizer comprised of a pair of toroidal fluid passages filled with a viscous fluid orthogonally mounted with respect to one another on a spin stabilized projectile. One torus is rigidly attached to the projectile and aligned with the x axis of the projectile and the other torus is similarly attached and aligned along the y axis.

Abstract: An improvement to a variable-pressure solid-propellant-powered maneuvering ystem for spacecraft, enabling the system to operate for longer than one propellant burn time. At least two solid propellant gas generators are connected via a manifold to a plurality of nozzle valve clusters; at least one of the gas generators is connected to the manifold via a sequence valve which isolates that generator from the manifold until such time as it is desired to put that generator into operation.

Abstract: A device is disclosed for regulating a set value and/or for stabilizing elastic objects subject to spin and free to move around their axis of rotation, in particular aircraft and spacecraft. For that purpose, the regulating signals required to regulate the set value or to stabilize the freely moving objects can be obtained by means of a regulator having a transfer function with a denominator degree about three orders higher than the numerator.

Abstract: A combination vibration generator and dampener having magnetic suspension and servo-control along three axes includes a hollow outer body fixed to a structure whose vibration is to be controlled, and an inner body which is essentially spherical in shape and which is inside the hollow outer body. Six active magnetic bearing elements are diametrically disposed in pairs along three orthogonal axes and serve to hold the inner spherical body inside the hollow outer body without direct contact between the inner and outer body. At least three position detectors detect the position of the inner body relative to the hollow outer body while at least one vibration detector device detects vibration forces on the structure whose vibration is to be controlled.

Abstract: A method for controlling transition from thruster control to momentum wheel control to minimize nutation angle and transition time in a three-axis stabilized spacecraft employing an internal body fixed-momentum wheel as an attitude stabilizer. The method employs direct, substantially real-time measurement of momentum components and angular rates and generates paired roll thrust impluses and paired yaw thrust impulses in synchronization with the nutation period of the spacecraft to drive the spacecraft to a state of substantially zero nutation. The controller employs a deadbeat principle in a closed-loop real-time feedback system. In a preferred embodiment, the impulsive thruster firing commands in roll and yaw are simultaneous, spaced one-half nutation period apart in time with the effect that the target momentum is achieved in one to three pairs of impulse firings.

Abstract: In order to shift the moment of inertia H.sub.i of a body in rotation initially in rotation approximately about its principal axis of inertia Z into any given orientation H.sub.F one applies transversely to the axis Z three torque impulses in a plane containing H.sub.i and H.sub.F, the first for a time T.sub.1, such that, at the end of one rotation of a satellite about itself, the axis Z shifts to Z.sub.1 in the plane bisecting H.sub.i and H.sub.F, whilst the kinetic moment moves to H.sub.1, the second impulse at the instant when the principal axis of inertia is at Z.sub.1 for a time T.sub.2 appropriate to shift the kinetic moment into a position H.sub.2 symmetrical with H.sub.1 with respect to this bisecting plane, and the third impulse at the instant when the principal axis of inertia is parallel to H.sub.F for a duration equal to T.sub.1.

Abstract: A high precision linear actuator for compensating short stroke, high frequency vibrations in a space deployed structure comprises a set of cylindrical magnetic pole pieces affixed to respective ones of a pair of graphite flexures, that are mounted to a thermally conductive cylindrical housing, which is to be affixed to the space deployed structure. Reaction against the "proof mass" pole pieces is achieved by a magnetic field produced by set of stationary magnetic coils that surround the pole pieces and are energized by a vibration compensation reaction current supplied from a sense/control processor. Coaxial with and mounted integrally with the cylindrical pole pieces of the proof mass is an internal accelerater, which generates an output signal representative of the force to which the proof mass is subjected.