Abstract: A momentum wheel (14, 142) for compensating for the adverse effects of disturbing torques on stabilized spacecraft is mounted on a base (12, 124) on the spacecraft by three circumferentially spaced supports which include at least two linear actuators (26, 118) that tilt the wheel (14, 142) in any desired direction. One end of each actuator (26) is mounted to a base (38, 154) supporting the wheel by a spherical connection (32, 34, 36) which allows freedom of motion about multiple axes. The other end of each actuator (26) is mounted to the base (12) by either a pair of spaced hinges (22, 24, 46) and a support bracket (18) which is secured to the actuator (26) or a flexible bracket (78). The wheel base (12,124) may also be mounted on the actuators by flexible members (150) which allow tilting of the wheel.

Abstract: The present invention is a method and apparatus to control a pointing trajectory of an aim point with respect to one or more targets for a momentum bias spacecraft in an orbit inclined relative to an equatorial geosynchronous orbit wherein the spacecraft is provided with an onboard computer capable of using orbit information available to the spacecraft and spacecraft hardware to move the aim point according to a predetermined tracking model which may be either time-varying or invariant. The tracking model allows the controller to remove short-term nutational and long-term orbit rate dynamical disturbances, beyond simple-order sinusoidal functions having a single frequency.

Abstract: A geostationary telecommunication satellite has a platform pointed towards the sun and a payload that is rotatable so as to point always toward the Earth. A method of placing this satellite in its final orbit entails using the inertia wheel of the satellite together with its sensors and possibly its two-axis rate gyro to secure attitude control of the satellite during the transfer, that is while in the transfer orbit and during the apogee burn.

Abstract: An essentially passive and "fuel-less" method for inverting the orientation of a preferably nutationally stable, dual spin spacecraft disposed in an inclined orbit, includes the steps of increasing the rotational speed of (i.e., "spinning up") the spacecraft's despun platform and decreasing the rotational speed of (i.e., "spinning down") the spacecraft's rotor, to thereby generate, via product of inertia coupling, a transverse torque of sufficient magnitude to temporarily destabilize the spacecraft and cause the spacecraft spin axis, which is the minimum moment of inertia axis of the spacecraft, to diverge and precess through a flat spin orientation and towards a final, inverted orientation, e.g., disposed at a precession angle of 180.degree. relative to the initial orientation of the spacecraft spin axis.

Abstract: The system for stabilizing a support comprises a working rotor mounted on the support by frictionless bearings and fixed to a working member for which at least one parameter associated with the rotation of the working rotor is variable as a function of time. A nested reaction rotor is mounted concentrically relative to the working rotor by frictionless bearings such as magnetic bearings or gas bearings and it supports a reaction inertia member. A single servo-controlled electric motor is interposed between the working rotor and the reaction rotor in order to drive the working rotor and the reaction rotor in opposite directions in such a manner that the torque disturbances generated on the support by the motion of the working rotor are cancelled by the simultaneous reaction motion of the reaction rotor, with the ratio of the speeds of the working rotor and the reaction rotor being such that the total angular momentum of the working rotor and of the reaction rotor is kept equal to zero.

Abstract: A first orientation control system controls the orientation of the main body of a satellite so that the main body is oriented toward the center of the Earth, and a second orientation control system controls the orientation of an antenna so that the antenna is continuously directed toward a target (e.g., a satellite revolving around the Earth). Between these two control systems, there is inevitably mutual interference, due to dynamic effects. To prevent this mutual interference, the first control system controls the orientation of the main body of the satellite on the basis of antenna driving information of the second orientation control system, while the second control system corrects the antenna driving information on the basis of orientation error information of the first control system.

Abstract: The balance of a spin-balanced spacecraft may be dynamically adjusted by moving a liquid between or among containers spaced about the spin axis. The transfer of liquid is accomplished by controllable heaters associated with each container, for heating that container from which liquid is to be transferred to increase the pressure and drive out liquid. A spacecraft for accomplishing this method includes a balance sensing arrangements such as a gyroscope and logic for determining which container or containers are to be emptied and for controlling the heaters. Weight is minimized by accomplishing the transfer of liquid among fuel containers by way of manifolds which also supply fuel to thrusters.

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 management system is provided. The management system monitors the wheel speeds and generates a wheel speed error vector. The error vector is integrated, and the error vector and its integral are combined to form a correction vector. The correction vector is summed with the attitude control torque command signals for driving the reaction wheels.

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: The system for stabilizing a support to be stabilized comprises a working rotor mounted on the support via bearings and fixed to a working member having at least one parameter associated with the rotation of the working rotor and which is variable as a function of time. A servo-controlled first electric motor is interposed between the support and the working rotor and serves to rotate the working rotor in a first direction. A reaction rotor is mounted on the support by means of bearings and independently from the working rotor, while being coaxial therewith.

Abstract: An improved fluid actuating system for imparting motion to a body such as a spacecraft is disclosed. The fluid actuating system consists of a fluid mass that may be controllably accelerated through at least one fluid path whereby an opposite acceleration is experienced by the spacecraft. For full control of the spacecraft's orientation, the system would include a plurality of fluid paths. The fluid paths may be circular or irregular, and the fluid paths may located on the interior or exterior of the spacecraft.

Abstract: A reaction-wheel stabilized spacecraft reduces attitude errors at wheel reversals by application of a dither component to the wheel torque command signal.

Abstract: A spacecraft incorporating momentum sources such as spinning momentum wheels having momentum vectors or components thereof oriented along the desired axis of rotation of the spacecraft. The rotation axis of the spacecraft is nominally any one of the three principal moment of inertia axes of the body. Proper adjustment of the spin rate of the main body relative to the magnitude of the moment sources angular momentum will insure that the spacecraft will spin about the desired axis in a passively stable way even though the rotational axis is not the principal axis of maximum moment of inertia.

Abstract: In an attitude control system for a spacecraft having a frame which is to be maintained in a desired orientation relative to a selected set of reference axes, which system includes an electromechanical actuator mechanically coupled to the spacecraft frame for applying to the frame a mechanical torque which opposes deviations from the desired orientation relative to at least one reference coordinate, the actuator including an electric motor and an element connected to be rotated by the motor, and a torque command signal generator connected to the motor for applying a torque command signal which is a function of such deviations and which drives the motor in a manner to control the torque being produced by the actuator, there is included an energy desaturating system composed of: a unit coupled to the element for producing a signal representative of element motion; a signal modifying circuit connected to produce a control signal having a magnitude which is a function of the signal representative of element motio

Abstract: Method and apparatus for controlling the attitude of a satellite in space by thruster reaction forces without loss of thruster fluid into space. A jet of gas, preferably compressed air, comprising reaction mass stored in a reservoir on board, is discharged in a predetermined direction into an endless enclosed evacuated volume. The molecules of discharged gas are scavenged from the volume and recompressed and restored to the reservoir. The discharged gas is scavenged sufficiently rapidly from the volume to prevent appreciable increase in the static pressure encountered by repeated bursts from the jets. The apparatus comprises an endless duct network circumferentially surrounding the two axes of the satellite, with solenoid operated thrusters selectively to discharge the gas jets and a vacuum-pressure pump to scavenge the air molecules, recompress them and pump them back into the reservoir tank.

Abstract: A system and method for attitude control in a geosynchronous satellite to compensate for roll and yaw pointing errors consequent to orbit inclination variation from the nominal equatorial orbit plane provides for an inertially fixed momentum vector coupled to the satellite through a gimbal system providing a one and preferably two degree-of-freedom relationship with the momentum vector. In a two degree-of-freedom embodiment, the momentum vector is established by spinning the satellite about an axis or providing an independent momentum wheel with the gimbal axes provided along the roll and yaw axes. Gimbal torquers torque the satellite about the inertially fixed momentum vector in a time-varying manner to effect correction of the roll and yaw pointing errors. Roll and yaw pointing errors consequent to orbit inclination drift from the nominal equatorial orbit are corrected in a fuel-efficient manner to extend the operating life of the satellite.

Abstract: Device and method for aiming a space probe toward a celestial body. It comprises a solar sail (30, 31) having an asymmetry about an axis (Z) and which subjects the probe (S) to a tilting torque of solar pressure. A kinetic wheel (20) turns about a perpendicular axis (X). This results in a rotation of the probe about the axis perpendicular to the two preceding axes (Y). The rotational speed of the wheel (20) is varied by a control system (23, 24) in order to change the rotational speed of the probe.Application of the invention is to space probes that lie in the plane of the ecliptic.

Abstract: An active vibration damper system, for bending in two orthogonal directions and torsion, in each of three mutually perpendicular axes is located at the extremities of the flexible appendages of a space platform. The system components for each axis includes: an accelerometer, filtering and signal processing apparatus, and a DC motor-inertia wheel torquer. The motor torquer, when driven by a voltage proportional to the relative vibration tip velocity, produces a reaction torque for opposing and therefore damping a specific modal velocity of vibration. The relative tip velocity is obtained by integrating the difference between the signal output from the accelerometer located at the end of the appendage with the output of a usually carried accelerometer located on a relatively rigid body portion of the space platform. A selector switch, with sequential stepping logic or highest modal vibration energy logic, steps to another modal tip velocity channel and receives a signal voltage to damp another vibration mode.

Abstract: This method serves for calibrating gyros forming part of the attitude control system of a satellite stabilized in three axes and which have an unknown drift. A two-axis sensor measuring the direction of a radiating reference object in the coordinate system x, y, z fixed with respect to the satellite is used. Two reference attitudes each different with respect to the reference object are commanded to the satellite in the course of two consecutive time intervals spaced from each other. The values measured by the sensor representing the respective actual direction of the reference object are recorded at the start and at the end of the time intervals. The respective time integrals of the gyro output signals are computed during the time intervals, which represent the gyro drift plus the respective instantaneous satellite deviation from the commanded reference attitude. Finally the gyro drift is determined from these time integrals and the values measured by the sensor.

Abstract: A reacquisition process is used in the event that an earth orientation of the yaw axis Z of an earth satellite is lost, in that the angular velocity .omega. of a flywheel, the rotational axis of which is oriented in the direction of a pitch axis Y, falls below a predetermined value as a result of an insufficient energy supply. After restoration of the energy supply, the pitch attitude of the earth satellite and, accordingly, the earth orientation of the yaw axis Z, are to be restored. For this purpose, constantly alternating increases and decreases in speed are commanded for the flywheel in a determined manner in correlation with the zero passages of the pitch deviation, whereby the satellite is subjected to a constant reversal of its direction of rotation. This occurs until the angular momentum of the satellite falls below a predetermined value during a zero passage of the pitch deviation.

Abstract: Large angular control moments and torques are developed by controllably circulating a relatively small mass of liquid (22) through small diameter pipes (20) describing a large diameter loop (15f,a,b). The loop (15), by thus generating and storing angular momentum, can thereby provide efficient cancellation of periodic, non-accumulating, externally induced rotational disturbances. The loop (15) is preferably located on or near the periphery of a structure (10) which is to be thus stabilized.

Abstract: Apparatus for autonomously performing stationkeeping maneuvers for three-axis stabilized spacecraft (1) such as geosynchronous satellites. For each of one or more spacecraft axes (y, z) the invention autonomously performs desaturation of a momentum/reaction wheel (31-32, 41, respectively) associated with that axis, while simultaneously accomplishing the preselected compensation of the spacecraft's east-west position. Thrusters (35-38, 45-48) having a polarity corresponding to the desired desaturation polarity are fired in a particular sequence: when a preselected east-west firing bias is present, the thrusters fired are solely from that face of the spacecraft (1) needed to counteract the east-west deviation. After the bias has been worked down, thrusters are fired from alternating spacecraft faces.

Abstract: An improved optical rotary joint is achieved by transmitting information in the form of an optical signal from a part of the de-spun portion outside the rotational axis to create on the spun portion, or vice versa, 360.degree. about the rotational axis, a continuous communication path having a configuration described by a series of overlapping elliptical areas. The detection area on the spun portion is less than that of the detection path thereby providing continuous optical communication with the transmitted optical signal.

Abstract: A space craft cellular energy generating and storage device for generating electrical power in space and storing the power for extended periods of time. The space craft generating and storage device characterized by having a pair of annular rotating housings having a plurality of individual compartments therein. The compartments receiving material which is compartmentalized to accurately control the mass distribution during the rotation of the housing for maintaining speed as energy is withdrawn or generated.

Abstract: A three axis attitude control system for an orbiting body, such as a spacecraft, comprises a motor driven flywheel supported by a torque producing active magnetic bearing. Free rotation of the flywheel is provided about its central axis and together with limited angular torsional deflections of the flywheel about two orthogonal axes which are perpendicular to the central axis. The motor comprises an electronically commutated DC motor, while the magnetic bearing comprises a radially servoed permanent magnet biased magnetic bearing capable of producing cross-axis torques on the flywheel. Three body attitude sensors for pitch, yaw and roll generate respective command signals along three mutually orthogonal axes (x, y, z) which are coupled to circuit means for energizing a set of control coils for producing torques about two of the axes (x and y) and speed control of the flywheel about the third (z) axis.

Abstract: A nutation damping control for a dual-spin satellite is provided by tilting the axis of the rotor relative to the normal spin axis of the satellite and back a half nutation period later. The tilting is in the direction of the angular momentum vector component that is transverse to the rotor's spin axis.

Abstract: A three-axis controlled spacecraft (1), typically a satellite, is spun up about its roll axis (20) prior to firing a motor (2), i.e., a perigee kick motor, to achieve the requisite degree of angular momentum stiffness. Thrusters (21) for imparting rotation about the roll axis (20) are activated in open-loop fashion, typically at less than full duty cycle. Cross-axis torques induced by this rotational motion are compensated for by means of closed control loops for each of the pitch and yaw axes (30, 40, respectively). Each closed control loop combines a prebias torque (72) with torques (75, 74) representative of position and rate feedback information, respectively. A deadband (52) within each closed control loop can be widened during the spinup, to conserve fuel. Position feedback information (75) in each of the control loops is disabled upon saturation of the gyroscope associated with the roll axis (20).

Abstract: An energy storage attitude control and reference system for a craft including at least two flywheels with their angular momenta balanced to produce zero net angular momentum and at least two motor/generator units each including one of the flywheels in its rotor structure. The system further includes one pair of bearings for supporting each of the flywheels, means for measuring the position of each flywheel relative to its bearings, means for resolving current to and from said motor/generator units, means for applying torques to each flywheel through its respective bearings for controlling attitude of the craft about two mutually perpendicular axes, and means for exerting torques to each flywheel through its respective motor/generator to provide attitude control of the craft about a third axis perpendicular to the first two and for controlling the energy storage level of the flywheels.

Abstract: In a satellite or space station having at least one stabilized service module and a payload module connected thereto in different orientations, an interconnecting and coupling structure includes latchable lift-and-turn mechanisms having some components in the service module and remaining components in the payload module; relative lift is provided such that the components are interconnected to obtain the interconnection of the modules; the modules are independently and releasably interconnected for latching and locking them together while the lift-and-turn mechanism is disconnected.

Abstract: A rotator apparatus comprising at least one channel capable of containing a flowable material and capable of assuming a first configuration and a second configuration at least one pump capable of causing the flowable material to flow in the channel; and a configuration change motor capable of moving the channel between the first configuration and the second configuration. An acceleration mode and a deceleration mode are described useful to control the direction and/or speed of rotation of the present apparatus. A further rotator apparatus is disclosed and comprises a plurality of motor/flywheel assemblies arranged so as to have at least one and preferably two assemblies per each of the three axes of rotation. This apparatus also has an acceleration mode useful to control the direction of rotation of the present apparatus.

Abstract: An improved system and method for ejecting and stabilizing a satellite/perigee stage assembly from a shuttle or orbiter type craft in space where the satellite includes gyros aligned parallel to three orthogonal axes, a computer, and spin and attitude control thrusters is described. The method comprises the steps of orienting the shuttle such that the perigee motor is pointing in the desired firing attitude, initializing said gyros and initializing the attitude determination algorithm in the spacecraft's computer, ejecting the spacecraft/perigee stage assembly without spin, after ejecting, slowly spinning the assembly about a desired spin axis, damping the assembly's nutation and precessing the assembly to the initial attitude while slowly spinning, and when at the initial attitude, spinning up the assembly for perigee motor firing.

Abstract: An improved system and method of ejecting a spacecraft from a shuttle or orbiter type craft in space includes a momentum wheel on the spacecraft that has a spin axis parallel to the ejection direction, spinning-up this momentum wheel prior to ejection and after ejection de-spinning this wheel whereby the whole spacecraft spins without the need of a spin table on the shuttle or orbiter craft.

Abstract: A torque orientation device (10) selectively orients a space craft without the use of jets or the ejection of any material. Pedestals (12, 14) support a gimbal (16) which is selectively driven by a motor (20). The gimbal (16) includes a frame (24) which supports an axle (30), a motor (26) and a commutator (28). Mass members (36, 38) are driven about the axle (30) by the motor (26). The moment of inertia about the primary axle (18) of the gimbal (16) is a function of the angular position of the mass members (36, 38) about the axle (30). The motor (20) applies selective torque impulses at periods (T1, T2) to the primary axle (18) in synchronism with the angular position of the mass members (36, 38) about the axle (30). The resulting counter torque from the motor (20) is passed to the space craft to cause the space craft to rotate toward the desired angular position.

Abstract: Equatorial orbit satellite with solar panels having blades with a medium line inclined at a certain angle relative to the equatorial plane. Thus, the field of vision of the antennaes is freed and disturbing torques become acceptable.

Abstract: The nutation damper comprises a closed flow path for a dense liquid, which may be a propellent and a pump for setting the liquid in periodic motion, thereby to dissipate or generate rotational energy controlled by a detector sensitive to a parameter relating to the nutation angle, according to the law of conservation of angular momentum. The principal application of the device is to the nutation damping of spinning spacecraft, in which the ratio between the moment of inertia about the spin axis and the lateral moment inertia is greater or less than 1.

Abstract: An agile payload orientation system for a space satellite comprises a three-axis gimbal (30) for coupling an appended body (20) containing the payload (24, 25, 26, 27) to a main body (10) of the satellite. The appended body (20) also contains moment control gyroscopes (29), which are responsive to external command for changing the orientation of the appended body (20) relative to an inertial reference. The gimbal (30) comprises three separate torquer-resolver units (31, 32, 33), which are connected to each other in series so that their axes are mutually orthogonal to each other and coincide at a single point (34) located at the center of mass of the appended body (20). Since the center of mass of the appended body (20) does not move when the appended body (20) is rotated about the axis of any of the torquer-resolver units (31, 32, 33), a change in orientation of the appended body (20) can be made without disturbing the orientation of the main body (10).

Abstract: An orientation maneuver for a bias momentum stabilized spacecraft is carried out on a spacecraft initially spinning about a minimum moment of inertia axis. With the spacecraft spinning about that axis, it is precessed until the angular momentum vector points to the south along the orbit normal. The spin rate is then reduced until the angular momentum remaining in the spacecraft is substantially equal to the nominal angular momentum of the momentum wheel in an in-orbit operation. The momentum wheel is then energized and gradually spun up until it contains its nominal angular momentum. As the wheel accelerates, the angular momentum will be redistributed between the wheel and the body, with the total system angular momentum remaining constant. When the wheel has reached its final speed, the spacecraft will have re-oriented itself such that the body of the spacecraft is spinning about the positive pitch axis, coning about the angular momentum vector in a nutational motion.

Abstract: An attitude control system is provided for a body-stabilized orbiting space vehicle having roll, pitch and yaw axes. Sensors carried by the vehicle provide signals representative of angular deviations about at least two of said orthogonal axes. An actuator for the system comprises two wheels rotatable about respective non-aligned spin axes fixed with respect to the vehicle. An electronic control circuit receives signals from the sensors and controls the rotation speeds of the wheels for attitude correction about two of said orthogonal axes. Short term stabilization about the third of the orthogonal axes is provided by the gyroscopic stiffness of the actuator.

Abstract: A satellite, especially a geostationary satellite which may be stabilized in its three major axes and which has a spin wheel or momentum wheel supporting a position control system and which also has an engine or propelling system, is controlled by actuating the spin or momentum wheel at the latest when the satellite is in a phase of being transferred from one orbit to another orbit. Thereafter, the satellite is rotated about its axis which extends in parallel to the axis of the spin wheel or momentum wheel whereby the rotational impulse of the satellite is smaller than that of the momentum wheel. The stabilizing of the satellite relative to the three major axes is then made at the point in time when the satellite is to change from one orbit to another orbit whereupon the engine system for reaching the new orbit is switched on. The thrust of the engine system is so dimensioned that the position control system intended for a subsequent mission retains its orientation even during the phase of changing orbits.

Abstract: A preselected axis other than the axis of maximum moment of inertia of a momentum stabilized spacecraft is oriented for alignment with the total angular momentum vector of the spacecraft by rotating one or more rotors to a critical rate of rotation. The rotation of the rotor or rotors is maintained at the critical rate to effect the desired rotation. The axis of the rotor or the effective angular momentum vector of the resultant momentum of a plurality of rotors is aligned with the axis of the desired spacecraft orientation. Energy dissipation assures stability by dissipating energy on the spacecraft body at a faster rate of dissipation than the rate of energy dissipation on the rotor.

Abstract: Apparatus and a related method for detecting failures of control elements, such as skewed-axis reaction wheels with closed-loop speed control, used in satellite attitude control systems, and for compensating for such failures and thereby maintaining continuity of operation. The apparatus includes a wheel loop model for simulating operation of each of the reaction wheels and speed control loops in the satellite, and for deriving an estimated wheel speed from a command speed for each wheel. By comparing the estimated speed of each wheel with the corresponding actual or observed speed, the apparatus determines whether a failure was occurred in the wheel, or in its control loop.

Abstract: An attitude control and stabilization system for momentum biased satellites utilizing a pair of counterrotating flywheels mounted parallel to its yaw axis. The speed of one of the wheels of the pair is controlled by sensing and correcting the roll axis errors of the satellite.

Abstract: A set of two two-degrees-of-freedom rate gyroscopes and three linear accelerometers are assembled in a single module adapted to be mounted within a single aircraft electronics control unit, the unit comprising a strapped-down attitude and heading reference system. The module base provides a common keyed support for the two pre-calibrated gyros and three accelerometers in intimate mechanical and thermal association. The two gyros are oriented in the base and the base oriented in the aircraft so that the spin axis of one gyro is oriented parallel to the aircraft Z axis and that of the other gyro parallel to the aircraft Y axis while the gyro pickoffs and torquers (input and output axes, respectively) are rotated or skewed forty-five degrees about the spin axes to positions such that the input and output axes lie along the slant heights of a forty-five degree half angle right circular cone, the axis of which lies along the aircraft X axis.

Abstract: An inertia wheel more particularly adapted to space applications. It includes the combination of a peripherical type of motor with permanent magnet on the rotor and ironless winding on the stator. The current of said winding is switched electronically by an amplitude modulation system, associated to a reactance coefficient varying circuit, and reversal of the direction of rotation of which is achieved by permutation of the control circuits. There are also provided bearings formed by a passive radial magnetic centering device and a redundant active axial magnetic centering device slaved to an axial rate detector.

Abstract: An apparatus for controlling nutation motion in a spinning body, comprising an angular accelerometer having its input axis perpendicular to the spin axis of the body, a flywheel having an axis of rotation perpendicular to the axis of the accelerometer and to the spin axis of the body, and a motor for driving the flywheel to attenuate or build nutation. The motor is controlled by circuitry that monitors the output of the angular accelerometer and drives the motor clockwise or counterclockwise during predetermined nutation angles synchronized to the zero crossover points of the accelerometer signal and centered about the nutation peaks. In one embodiment, the motor drive is phased to damp nutation motion to zero for stabilization.

Abstract: A circuit for applying rate feedback to the pitch and yaw canard commands a guided projectile or missile to stabilize the airframe. The circuit demodulates and differentiates the signal from the gyro cage coil to obtain a signal representative of the airframe movement at its natural frequency. Phase detectors provide a phase reference by comparing the cage signal carrier with the output of hall effect elements in the speed sensing circuitry. The rate amplitude from the differentiator is then multiplied by the phase references to resolve the rate into pitch and yaw command components. Summation circuits then scale and subtract the rate commands from the guidance commands to provide negative rate feedback.

Abstract: The invention concerns stabilized platform arrangements of the kind utilized, for example, to carry a rotatable ship's radio antenna, which requires to be stabilized against pitch and roll. A control system is provided which includes means for relating error signals representative of deviations in pitch and roll to orthogonal axes fixed with respect to the antenna, by conversion to angular errors in two orthogonal planes, the elevation and cross elevation planes, containing the reference axis about which the antenna rotates.

Abstract: An inertial guidance system including a true space stable platform is provided while retaining a simple two-degree-of-freedom gimbal system. An added third degree of freedom is provided by motions of the support itself which maintains desired platform orientation. The output of a third axis instrument is used to maneuver the frame itself about the axis of the frame which most nearly will null any rotation about the platform roll or longitudinal axis. A rotational rate gyroscope is positioned on the space stable platform, and actuates the movement of the frame about the longitudinal or roll axis directly from the reading of the rotational movement along the roll axis of the platform.

Abstract: Roll and yaw attitude errors introduced by orbit inclination deviations from the nominal orbit plane are minimized by sinusoidally varying the momentum produced by a transverse wheel mounted on the spacecraft. The wheel is mounted on the spacecraft such that its axis is parallel to the spacecraft's yaw axis. Sinusoidal variation of wheel momentum is obtained by sinusoidally varying wheel speed in response to a sine wave signal periodically updated from an earth station. In response to the sinusoidal variation of transverse wheel momentum, the spacecraft is rolled to minimize thereby the roll error introduced by the orbit deviation from the nominal orbit plane. Yaw error is minimized by providing sufficient transverse wheel momentum so as to maintain the total spacecraft momentum vector perpendicular to the nominal orbit plane.

Abstract: This invention relates to an attitude stabilizing system for a rod-shaped satellite in which stabilization is achieved by use of a spinning energy dissipator. The system can operate by being mounted on the rotor of a dual spin satellite or by being connected to the main body of the satellite through a motor and bearing assembly. In the latter situation, the stabilizer acts both as a nutation damper and a momentum source and thereby eliminates the need for an additional rotor element to provide the spacecraft momentum. The spinning energy dissipator may be designed to use four identical wheels mounted coaxially with and at the ends of the arms of a cruciform structure. The structure may then be spun about the axis perpendicular to the arms, to provide angular momentum along the spacecraft spin axis. In the event of any spacecraft nutation, the wheels experience the torques produced by inertial forces.