Abstract: An attitude control system using magnetic torquing is enhanced using a momentum wheel which is pivoted if the sensed error signal about the roll axis exceeds a given threshold. The momentum wheel is pivoted under control of a stepping motor through a first angle proportional to the detected average roll error, and one-half nutation period later through an equal angle in the same direction. Alternatively, the momentum wheel is pivoted over first and second angles of different values wherein the difference is related to nutation.

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: A system for actively damping nutation of a satellite and/or aligning the pitch axis of a momentum bias satellite with some external reference includes a magnetic torquing means which is energized with currents of appropriate polarity and magnitude to effect the necessary torque to reduce nutation. A roll sensor produces signals proportional to the roll error of the satellite. A control signal is generated a predetermined time period after the roll error exceeds a predetermined threshold. The predetermined time period is a fraction x of the nutation period t.sub.n of the satellite or x+n(t.sub.n) where n is an integer.

Abstract: The angular momentum of a tumbling orbiting satellite is changed to reduce the spinning of the spacecraft such that the spacecraft can operate in an essentially stabilized condition. The change in spin rate is accomplished by the development of a magnetic torque by commutation of magnetic air coils in response to earth sensor signals controlled by spacecraft logic. The coils are disposed to produce magnetic dipoles transverse to the satellite spin axis such that by interacting with the earth's magnetic field a torque is developed in a direction opposite the spin vector. The logic selects the proper magnetic dipole based upon data from sensors detecting the presence of the earth for the development of the desired magnetic torque.

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: For a spacecraft in an equatorial or low inclination orbit predictable roll disturbance torques due principally to solar radiation pressure are minimized by interacting a magnetic dipole oriented parallel to the spacecraft's yaw axis with the earth's magnetic field. Signals either from an on board sun sensor or from an earth station are utilized by the satellite mounted magnetic switching and driver circuit to control the compensating torques.

Abstract: Control of an active three-axis multiple-wheel attitude control system is provided with a bias momentum to achieve stiffness about the pitch axis. With this stiffness, the system overcomes the inherent inability of a zero momentum system employing the sun as its yaw reference to be provided with this yaw reference during the periods encompassing satellite high noon and midnight (i.e., approximate coalignment of the sun line and the local vertical). The satellite attitude is normally maintained by angular momentum exchange developed by three or more reaction wheels positioned on the satellite. The control mechanism is arranged to provide net angular momentum in the pitch axis wheels to achieve the momentum bias or stiffness along the pitch axis, whereas the time average of the angular momentum about the roll and yaw axes is substantially zero.

Abstract: Attitude of a pitch momentum biased satellite in a low inclination orbit is controlled by a closed loop magnetic torquing control system. A magnetic dipole oriented perpendicular to the satellite's pitch axis in the roll/yaw plane and at a predetermined azimuth angle relative to the roll axis minimizes roll and yaw errors by interaction with the earth's magnetic field. Signals from an on-board roll attitude sensor are used by satellite mounted detection and driver circuits to control the compensating torques. Magnetic torquing is initiated whenever the roll error as measured by the roll attitude sensor exceeds a predetermined threshold.

Abstract: Spacecraft nutation caused by applying propulsive forces to a spacecraft along an axis which does not pass through the spacecraft's center of gravity is eliminated by a signal responsive control system which operates the propulsive forces for a predetermined time period which corresponds to an integral number of nutation periods.

Abstract: A system for autonomously unloading accumulated angular momentum on space craft reaction wheels in the attitude control systems with external torques generated by the interaction of spacecraft magnetic dipoles with the earth's magnetic field.An onboard computer receives reaction wheel speed information, wheel speed threshold and geomagnetic latitude information and controls magnetic dipoles that will create a spacecraft torque and desaturate reaction wheels in a closed loop system.

Abstract: Three-axis attitude control of a substantially zero momentum satellite is achieved by a control system which includes three mutually orthogonal reaction wheels and a fourth backup reaction wheel. The backup wheel is mounted on an axis which is skewed with respect to the axes of the three mutually perpendicular wheels. Spacecraft attitude is normally maintained by torques developed by the mutually orthogonal wheels. If one and only one of the orthogonal wheels should fail to correct the attitude error on its associated axis, the backup wheel is caused to rotate and develop a torque to maintain spacecraft attitude.

Abstract: A pair of reaction motors or thrusters connected to spin-stabilized spacecraft are used to reorient spacecraft attitude without affecting spacecraft velocity. The spacecraft reaction motors are arranged to produce a couple in response to an attitude reference signal without applying a spacecraft velocity changing force to the spin-stabilized spacecraft.

Abstract: The orientation maneuver of a bias momentum stabilized spacecraft is achieved without the need of attitude determination by sensors and gyros and the like and without the conventional spin-down and spin-up procedures. The spacecraft is provided with a momentum wheel mounted in perpendicular relation to the axis of maximum moment of inertia. The wheel is initially de-energized during the time the spacecraft is initially launched from the ground launching platform until it is desired to orient the spacecraft in the final orbit. The momentum wheel, when energized from zero rotation to increasing rotation speeds, causes the rotation of the spacecraft from spinning about the maximum moment of inertia axis to an axis parallel to the momentum wheel axis with the final convergence of the wheel axis to the momentum vector being effected by energy dissipation in a nutation damper.