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 triaxially stabilized earth oriented satellite is provided with an attitude control system. The attitude control system contains a controller, final control elements for generating control torques around each of three axes of a system of coordinates (principle axes x, y, z) fixed relative to the satellite. Additionally, the attitude control system comprises, as measuring transducers, exclusively a direction vector measurement device, like a magnetometer or star sensor, a sun sensor array and an earth sensor. The process of the invention makes use of the apparatus as noted above and includes the steps of:A. determining an estimate for the rotating velocity of the satellite from a direction vector measurement;B. searching for the sun, depending on the size of the field of view of the sensor, by one or two search rotations at right angles to the optical axis;C. adjusting the direction of the sun relative to the optical axis; andD. adjusting the direction vector measurement to a reference direction S.sub.

Abstract: A method and apparatus for satellite station keeping is disclosed in which six thrusters are mounted on the anti-nadir face of a satellite with their direction of thrust passing through the center of mass (CM) of the satellite. The thrust lines of the north trio of thrusters and the south trio of thrusters make an angle .theta. with the satellite north-south axis in a northern and southern direction, respectively. The four outer thrusters are laterally separated and slewed by an angle .alpha. about the north-south axis. Each outer thruster produces three components of .DELTA.V, i.e. normal, tangential and radial (toward the Earth), thereby providing complete control of the three orbit vectors, inclination, eccentricity and mean motion. The north and south center thrusters are positioned in the Local-Vertical North-South plane and produce only normal and radial .DELTA.V components, providing control for momentum dumping and, in the event of failure, inclination control.

Abstract: For controlling the attitude of the body of an earth satellite placed on a low orbit, values of components of the geomagnetic field of the earth are measured along three axes of a frame of reference bound to the body. The values are derivated with respect to time, and multiplied by a gain. Currents responsive to the multiplicated derivatives are passed through magnetic torquers located along the three axes of the body to create magnetic torques that bias the body to a fixed angular position relative to the field lines of the geomagnetic field. Such steps are continuously carried out during eclipse periods. Out of eclipse periods, the pitch of the body is controlled by modifying an internal momentum in response to a signal provided by a solar sensor, so as to maintain solar generators carried by the body of the satellite oriented towards the Sun.

Abstract: A spacecraft is provided which has an environmental disturbance torque filter that predicts the environmental disturbance torque acting on the spacecraft, and an unloading control section 5 which establishes, from the predicted environmental disturbance torque 4 and the angular momentum residual 2, a magnetic moment to be generated, by considering the environmental disturbance torque, thereby providing a reduction in magnetic unloading power consumption and a reduction in the size and weight of related hardware.

Abstract: An attitude or orientation control system for spacecraft that provides for yaw axis error reduction and momentum dumping using a simple calculation involving only the momentum wheel speeds as input and which outputs two axis (roll and yaw) torque commands. The system utilizes the spacecraft 1) momentum wheels that store spacecraft momentum, 2) processing means for the calculation of control torques, and 3) torque actuators, e.g., magnetic torquers, that can provide a torque axis anywhere in the spacecraft roll/yaw plane, and involves three fundamental steps, 1) the input of the momentum wheel speed mearsurements, and 2) the calculation of the required torques, using the wheel speed measurements, and and 3) the outputting of the required torques in a form appropriate for controlling the available actuators.

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: This invention discloses a method for compensating for the disturbance torque on an orbiting satellite (10) from the interference of the earth's magnetic field lines with particular current loops associated with the satellite (10). The current flow in these current loops on the satellite (10) is measured by a current sensing device (42). A scaling factor (44,46) is applied to the measured currents to derive an estimated disturbance torque in particular axes of the satellite's attitude. The estimated magnetic disturbance torques are applied, along with other control torques, to a satellite actuation device (38). The actuation device (38) actuates the satellite (10) in the particular axes to compensate for the disturbance torques being applied to the satellite (10).

Abstract: A magnetic torquing system for a spacecraft using conducting coils is disclosed. In one embodiment, the conductors (44, 46) on a spacecraft's (10) solar wing (28, 30) which connect the solar cell strings (38) to the voltage controller (48) are wired to produce a magnetic torque which can be regulated by shunting individual strings to ground (52) or by opening a string circuit. This embodiment does not require the extra weight of a coil because the panel's solar string produces the torque normally produced by an additional coil. In another embodiment, a coil (44, 46)is wired between a shunting switch (54) in the spacecraft voltage controller (48) and a ground (32) so that shunting a string to ground (52) supplies current first to the coil to generate a magnetic torque in the desired direction.

Abstract: A 3-axis stabilized spacecraft includes roll and yaw magnetic torquers, and a momentum wheel oriented with its spin axis orthogonal to, and pivotable about, the roll axis. Roll control may be applied by pivoting the wheel. Secular increases in pivot angle may result in loss of control authority when the mechanical limits of the pivot are reached. The pivot angle is sensed, and an unloading control loop is closed, by which magnetic torquers are energized to torque the spacecraft, to return the pivot angle toward zero. The unloading control loop includes a bandpass filter, which eliminates constant components of pivot angle offset. This prevents the unload control loop from attempting to maintain the pivot at a position in which the wheel axis is offset from the desired normal to the orbit plane. Consequently, the magnetic torquers do not expend system energy attempting to maintain an undesirable attitude.

Abstract: An attitude control device for a three-axis stabilized satellite in terrestrial orbit embodies an attitude sensing system adapted to deliver roll, yaw and pitch attitude signals, a set of at least three non-parallel axis momentum wheels and a set of at least two magnetic coils oriented at least approximately in two non-parallel directions in the roll/yaw plane. A primary control loop embodies a processor and controller connected between the attitude sensing system and the set of at least three momentum wheels. It is adapted to determine from attitude signals and attitude set point signals attitude correction signals for the momentum wheels so as to apply to the satellite primary attitude correction torques. A secondary control loop embodies a coil controller connected between the processor and controller, the set of at least three momentum wheels and the set of at least two magnetic coils.

Abstract: A roll/yaw attitude control system for a three-axis stabilized satellite in an at least approximately Equatorial orbit embodies a processor circuit connected between a roll, yaw and pitch attitude sensing device including a terrestrial sensor and a stellar sensor adapted to detect the Pole Star and an actuator device having a kinetic moment system substantially oriented along the pitch axis and a magnetic dipole generator system disposed at least approximately in a roll/yaw plane. The processor circuit embodies a short-term roll/yaw control loop adapted to estimate the roll and yaw attitude angles and angular speeds and to determine set point signals for some elements of the actuator device and a long-term roll/yaw control loop adapted to estimate roll/yaw attitude angles and external disturbances and to determine a dipole signal to be applied to the magnetic dipole generator system or to backup actuators.

Abstract: A spacecraft includes an attitude control system using one or more reaction wheels, the speed of which from time to time lie near and pass through zero angular velocity. When operated for extended periods of time at low speeds, the lubrication films are not distributed uniformly on the wheel bearings, leading to reduced lifetime. Reliability is maintained by a threshold comparator coupled to compare wheel speed with a lower limit value, for operating a torquer associated with the spacecraft body when the wheel speed drops below the lower limit, in a manner which tends to raise the wheel speed. In a particular embodiment of the invention, the lower limit is integrated with a wheel overspeed unloading.

Abstract: A satellite adapted to be stabilized in terrestrial orbit by rotation about an axis oriented at least approximately NORTH-SOUTH has on its external wall a solar generator composed of a plurality of columns of solar cells divided into groups adapted to supply at least one given electrical voltage. In each of a plurality of groups equi-angularly distributed in the circumferential direction, the cells straddle two columns electrically connected by at least two electrical connections forming with the group(s) to which the connection belong an electrical loop. The clockwise or counterclockwise direction of current flow is the same for least the majority of the electrical loops and is so chosen that, at any time, the loops which are illuminated at this time by the Sun, by virtue of interaction with the terrestrial magnetic field, produce a compensating torque in the opposite direction to the solar radiation pressure torque.

Abstract: A satellite comprising a deployable solar generator that can be oriented relative to the satellite body is kept facing the sun at all times. It includes an attitude stabilization device comprising a superconductive material closed loop carried by a support coupled to the solar generator and disposed on a side of the support facing substantailly away from the sun. A secondary loop parallel to the closed loop is substantially the same shape and size as the closed loop, and is carried by the support near the loop. A resistive circuit is carried by the support near at least part of the closed loop. A selective electrical power supply circuit is connected electrically to the secondary loop and to the resistive circuit.

Abstract: A method for maintaining the attitude of a three-axis controlled satellite by use of magnetic torquers includes using magnetometers for measuring the direction of the ambient geomagnetic field. The direction of the net reaction wheel momentum is also determined. The angle between the direction of the geomagnetic field and the net reaction wheel momentum is determined. The angle is compared with a threshold value. Magnetic torquer power consumption is reduced by operating the magnetic torquers only when the angle exceeds the threshold value.

Abstract: The magnitude and direction of the residual magnetic field or dipole of an orbiting spacecraft is estimated from the attitude-perturbing effects of its interaction with the magnetic field of the heavenly body about which it orbits. The measurements are performed over a significant portion of an orbit. The estimate uses h, the total spacecraft momentum in the body frame, and its body angular rate .omega.. The Earth's magnetic flux density B is determined either by magnetometer measurements, or by recourse to stored historical and ephemeris information.

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: Agile (electronically steerable) beam sensing with associated on-board processing, previously used exclusively for positioning of antennas for beam formation and tracking in communications systems, is now also used for satellite active attitude determination and control. A spinning satellite (100) is nadir oriented and precessed at orbit rate using magnetic torquing determined through use of an on-board stored magnetic field model (520) and attitude and orbit estimates (212). A Kalman filter (211) predicts parameters (202, 203) associated with a received signal (204) impinging on the satellite's wide angle beam antenna (201). The antenna system measures the error between the parameter predictions and observed values and sends appropriate error signals (207) to the Kalman filter for updating its estimation procedures.

Abstract: The attitude of a spacecraft can be effectively controlled in three axes using the combination of a gravity gradient boom and three electromagnets. The electromagnets are controlled according to a control law that takes into consideration signals from a magnetometer and a signal indicating the orbital angular velocity of the spacecraft.During the acquisition phase of the flight, the gravity gradient boom is not extended and electromagnets are operated according to a control law that only takes into account the output of the magnetometer. After the acquisition phase is complete, the gravity gradient boom is extended and the spacecraft enters the libration phase of the flight. In the libration phase of the flight, electromagnets are controlled according to a control law that takes into account both the output of the magnetometer and a signal which indicates the orbital angular velocity of the spacecraft.

Abstract: A method for determining the earth's magnetic field and, selectively, additionally, for determining a satellite orientation as a premise for the attitude control of an earth satellite in a known orbit. The satellite is equipped with magnetic coils 5, 6, and 7, gyros, sun sensors and reaction and/or flywheels 1, 2, 3, 4. The x-axis of a coordinate system integral with the satellite is continuously oriented towards the sun. The satellite can be rotated through a variable angle of rotation .alpha. around the x-axis. At three instants t.sub.0, t.sub.1, t.sub.2, spaced from each other by small time intervals, respective angular momentum components H.sub.y, H.sub.z are measured. Herein, a constant magnetic moment M.sub.x is generated throughout the second time interval. From these domains, the components of the earth's magnetic field B.sub.y, B.sub.z are calculated for t.sub.0. From this, subsequently, the angle of rotation .alpha. can be determined.

Abstract: Electrical machine including a base member, a rotor and a stator. The base member has a first magnetically permeable body with a first winding arrangement. The rotor has a second magnetically permeable body and is rotatable relative to the base member around first and second axes. The stator is adjacent to the base member, surrounds the rotor and has a stator core and second and third winding arrangements. The stator core is wound with the second winding arrangement held inward near the rotor, and wound with the third winding arrangement exterior to the stator core. The second winding arrangement is energizable to rotate the rotor around the first axis, and the first and third winding arrangements are jointly energizable to rotate the rotor around the second axis. Other electrical machines are also described.

Abstract: In the disclosed satellite and arrangement therefor, roll or yaw of a satite with pitch momentum bias and travelling on a substantially polar orbit with the aid of a coil interacting with the earth's magnetic surface includes a timing arrangement for dividing the orbit of the satellite into four successive quadrants P1, P2, P3, and P4, with the quadrants being inclined at an angle of 45.degree. to the equator and the quadrants P2 and P4 being polar quadrants and the quadrants P1 and P3 being equatorial quadrants. A first control unit produces satellite control signals in response to the timing means which are arranged to use the first control while the satellite is in the polar quadrants and the second control unit when the satellite is in the equatorial quadrants. The sign of the control unit is inverted in the quadrants P3 and P4, but maintained in the quadrants P1 and P2. Both controls use or generate internal signals corresponding to the roll angles and roll velocity to produce the control unit signals.

Abstract: A closed-loop magnetic roll/yaw control system for high inclination satellites where the predominant component of the geomagnetic field lies in the orbit plane of the satellite. The system includes a torquing means that forms a magnetic dipole moment parallel to the pitch axis. The torquing means is energized to minimize roll or yaw error only in two spaced portions 1 and 3 of its orbit. Each of the portions lie between the extension of the earth's magnetic equatorial plane and a second plane normal to the orbit and passing through the magnetic poles of the earth. Portions 1 and 3 are those parts of the orbit in which the satellite is traveling toward the second plane. The torquing means is de-energized after the satellite crosses the second plane and is in either of the remaining spaced portions 2 and 4 of its orbit.

Abstract: A magnetic bearing arrangement for rotor for controlling the attitude of an artificial satellite wherein the rotor has a rotating speed and an axis of rotation which are adjusted in accordance with a change in the attitude of the satellite. The arrangement includes a stator defined by at least three segments, and each stator segment is provided with a permanent magnet to eliminate the necessity of any electromagnet for controlling radial rotor displacement and any sensor for detecting such radial displacement. The apparatus is free from any eddy currents that may cause great resistance to rotation.

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: Electromagnetic method and apparatus to accurately control the orientation of a platform according to two axes wherein pivoting of the platform about two orthogonal axes passing through its center is effected by electromagnetic means, with the servo-control being obtained by using a detector responsive to the angle of deviation from a reference direction, an amplifier, a low-pass filter, a phase lead filter and power amplifiers supplying currents to the electromagnetic devices. The method and apparatus is particularly adapted to the orientation of platforms carrying means such as aerials or inertia wheels intended for piloting a space vehicle.

Abstract: The invention relates generally to an improved system for sensing and compensating for external disturbance forces acting on a satellite while in orbit. A proof mass member is housed within an enclosure and shielded from external, non-gravitational forces. The proof mass is electromagnetically levitated to move in a purely gravitational orbit, along an axis aligned with the satellite's velocity vector. The proof mass is subjected to a controlled magnetic biasing field and is caused to have a constant reaction to the resultant biasing force, by means of a thermal control system which maintains constant resistivity of the proof mass, during operation. The position of the proof mass with respect to its axis is detected optically and is utilized to control the firing of spacecraft thrusters. As a result, the satellite is caused to maintain a substantially constant position relative to the proof mass and thereby also is caused to follow a purely gravitational orbit.

Abstract: A nutation damper apparatus capable of damping small nutation angles in a satellite. A permanent magnet mounted on the free end of a pendulum arm oscillates back and forth parallel to an adjacent conductive plate in the form of a circular segment having a cross-sectional area that increases from the center to its lateral ends, and induces in the plate a magnetic flux which is proportional to the deviation angle of the pendulum from its rest position.

Abstract: The invention relates generally to a method of regulating the onboard voltage of a spacecraft and of steering the same by at least two contrarotating motor-generator dynamoelectrical machines.According to said method, there are provided two phases:The energy storage mode, during which each machine receives electrical energy from solar generators through a velocity controlling generator commanded from a synthesis of the data relating to position and rotor velocity;The energy restitution mode, during which each generator supplies energy to a load consisting of the onboard equipment of the satellite. Output monitoring systems effect a synthesis of the data relating to position rotor velocity and additionally control the power switches.

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: A momentum wheel wherein friction has been substantially eliminated by replacing mechanical bearings with a magnetic suspension. The rotor of the momentum wheel, in the form of a circular disk whose diameter greatly exceeds its thickness, is maintained in suspension, free of physical contact with the housing, by means of magnetic fields applied between the circular flat surfaces of the rotor disk and a pair of stator plates. The rotor is passively restrained in the radial direction by means of permanent magnets, while an active axial restraint is provided by a combination of permanent magnets and electromagnetic windings.

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: A bar magnet is suspended by torsion wires and a ring inside a spherical frame rigidly attached to the space vehicle. The frame is filled with a damping fluid and damping domes are attached to each end of the bar magnet. The damping fluid and damping domes cooperate in damping or resisting angular motion between the bar magnet and the frame.

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: The invention provides a satellite momentum wheel with passive magnetic centering means, between a stator and a rotor comprising at least one passive radial magnetic centering device, at least one axial magnetic centering device slaved to an axial velocity sensor and to an axial position sensor, at least one centering motor-generator unit having a magnetized rotor and a stator with toroidal windings switched by an angular rotation position sensor, at least one damping device, two emergency taper pivots for safe centering of the axial mounting of the rotor one of which is possibly in contact with a doped-polyamide step bearing on the stator, and interconnecting means between the wheel rim and the axis of said pivots employing a so-called coated arrayed-thread type cycloprofile structure.

Abstract: A satellite is compensated for torques caused by solar pressure on a solar panel extending from the satellite by an air core coil energized to develop a magnetic torque in opposition to the solar pressure torque. The coil is disposed generally in the periphery of the solar panel mounted externally to the satellite. The energizing current is fixed or varied by ground command or automatically in a closed loop system to compensate for variations in the earth's magnetic field.