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 technique for maintaining a satellite in an assigned orbit without control or intervention from the ground. Autonomously obtained navigational data provide a measurement of the actual orbit in which the satellite is traveling. So long as the measured orbit conforms to a desired orbit to within a preselected tolerance, periodic corrections of equal magnitude are made to the satellite's velocity, based on a prediction of the effect of atmospheric drag on the orbit. Measurement of the orbit is made by observation of the time that the satellite passes a reference point in the orbit, such as by crossing the ascending node. If the measured orbit departs from the desired orbit by more than the preselected tolerance, a velocity correction of a magnitude different from the one based on prediction is applied to the satellite. For a decaying orbit, the magnitude of the velocity correction is increased above the correction value based on prediction.

Abstract: Earth acquisition from the Sun pointing attitude starts with angular displacement of the satellite so that, in the field of view of a Sun sensing system, the direction of the Sun is brought into an orientation S' such that subsequent rotation of the satellite about the orientation S' brings the Pole Star into the field of view of a star sensing system whose optical axis is substantially parallel to the pitch axis. During this rotation the stars sensed are compared with those in a catalog containing, in addition to the Pole Star, stars likely to be encountered upon such movement. At least two of these stars are identified and then the Pole Star is captured. The satellite is then rotated in pitch until the Earth is sensed and captured.

Abstract: Satellite and method for placing said satellite in orbit by lunar gravitational assistance.After placing the satellite into standard orbit (01) within a quasi-equatorial plane, the satellite is transferred onto a circumlunar orbit (02) so that the satellite traverses the sphere of influence S1 of the moon. On leaving this sphere, the orbit (03) is inside a plane inclined with respect to the equatorial plane. The orbit is then connected to the definitive orbit (04).

Abstract: In an apparatus for orbit control of at least two co-located geostationary satellites a single interface is provided in the form of an orbit database for exchanging orbit and maneuver data between four independent decoupled function blocks of an orbit control system having a matched function scope, i.e. between a first function block for orbit determination and maneuver estimation/maneuver calibration, a second function block for maneuver planning, a third function block for monitoring the relative movement and a fourth function block for predicting specific events and ephemerides (FIG.

Abstract: A method and system (12) for controlling the attitude of a spacecraft during its transfer orbit using an on-board, stand-alone, three-axes attitude determination and control system. The system utilizes a set of on-board sensors to define two independent angular measurements, which will initially identify the z-axis orientation of the spacecraft from an arbitrary attitude after launch vehicle separation. A set of three-axis gyros are then bias calibrated in order to measure the transverse rates of the spacecraft. The three-axis attitude of the spacecraft is continously determined by integrating the gyro outputs even if the Earth or Sun is not visible by an on-board sensor. A state estimator model is provided to determine the three-axis attitude of the spacecraft in the presence of large wobble and nutation. The system also utilizes a linear combination of the estimated attitude, rate and acceleration states to generate commanded rate increments with a pulse-width frequency modulator.

Abstract: A method and apparatus for satellite station keeping is disclosed in which four thrusters are mounted on the anti-nadir face of the satellite with their direction of thrust passing through the center of mass of the satellite. The thrust lines of the north pair (20, 22) of thrusters and the south (24, 26) pair of thrusters make an angle .theta. with the satellite north-south axis in a northern and southern direction respectively. The thrusters are laterally separated and slewed by an angle .alpha. about the north-south axis. Each 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. In the event of failure of a thruster, the thruster diagonally opposite the failed thruster is shutdown and the remaining diagonal pair of thrusters is used to maintain station keeping.

Abstract: This device is used for controlling the attitude of a spacecraft to be rotated about a body's axis of rotation. Actuators generate torques about the axis of rotation and two lateral axes. Angular velocity signals .omega..sub.X, .omega..sub.Y with respect to the lateral axes are in each case fed to first and second signal paths. The latter contain an integrator. Modulators, which each comprise a variable dead zone, supply control signals for the actuators. In order to limit the nutation amplitude to a constant value in a reliable manner, multiplication elements are provided in the first and second signal paths. The lateral-axis angular velocity signals .omega..sub.X, .omega..sub.Y or the angular position signals .PHI., .theta. are acted upon by factors which are proportional to the rotation axis angular velocity signal .omega..sub.Z or its square in the multiplication elements. Furthermore, the thresholds of the dead zones of the modulators are varied proportionally to .omega..sub.Z.sup.2.

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 method of attitude measurement for an artificial satellite (100) utilizes one or more star trackers (12) together with an earth sensor (30). Periodic updates of satellite orbital information, either propagated onboard or from a ground station are combined with earth and star position coordinate data to provide a continuous and accurate measurement of the spacecraft body 3-axis attitude. The method can be used for ground-based attitude determination or onboard closed loop control systems.

Abstract: A triaxially stabilized satellite having a Cartesian axis system with mutually perpendicular X, Y and Z axes, two electric propulsors for orbital maneuvers and attitude control of the satellite and having respective versors including a right angle between them and lying in an X-Y plane, the versors being orientable in the plane about the Z axis to include a variable angle a with the X axis to impart two degrees of freedom of movement to the electric propulsors, the electric propulsors generating thrust vectors in line with the versors away from a center of gravity of the satellite at an intersection of the versors, the electric propulsors being firable independently and for respective angles of arc in their displacement with the degrees of freedom to maneuver the satellite orbitally, control attitude of the satellite and dissipate angular momentum of the satellite. The electric, preferably ionic, pulsars can be used with triaxially stabilized satellites with an LEO or GEO.

Abstract: Apparatus and methods of Gravity Guidance and Propulsion of Geosynchronous Satellites, other satellites and space vehicles using net kinetic energy PUSH of Gravity of the electromagnetic spectrum particles which continually irradiate the earth from space, based on the Oppositely Charged Twin Monopole (OCTM) Theory of Matter. Specifically Fully Stabilized Geosynchronous Satellites can be made with the same lift-off weight as Spin Stabilized Geosynchronous Satellites by using GG&P Methods and Rules of mass distribution in the satellite.

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: A spacecraft (201) maintains its north-south positioning by using one of two pairs of single-gimballed throttled thrusters (221-224) on a face of the spacecraft (201). The throttles (118) and gimbals (116) of the thrusters (221-224) are controlled to produce torques on the spacecraft (201) that will maintain a desired attitude for the spacecraft (201) while simultaneously desaturating the momentum stabilizing wheels ( 120, 121 ) of the spacecraft (201).

Abstract: A pitch momentum stabilized spacecraft includes an earth sensor for generating roll signals. The roll signals are applied to torquers for proportional (P) or proportional-integral PI control of roll and yaw. According to the invention, the sensed roll signals are bandpass filtered at the orbit rate, phase adjusted and applied to the roll and yaw torquers for closing a high gain degenerative feedback loop at the orbit rate, which reduces orbit-rate components of roll and yaw. A second bandpass filter at the second harmonic of the orbit rate is also connected to the earth sensor, for producing second harmonic signals, which are phase controlled and summed with the fundamental orbit rate signals, for degeneration of attitude perturbations occurring at twice the orbit frequency.

Abstract: This invention discloses a method for determining when an orbiting satellite (10) is eclipsed from the sun in order to remove control torques to the satellite (10) which compensate for the disturbance of solar pressure on the satellite (10). A current measuring device (46) measures the current traveling through a particular circuit associated with the satellite (10) which is indicative of the satellite batteries being discharged, as would occur during an eclipse. The measured current is applied to a threshold logic circuit (48) which sends a signal to a control compensator (36) if the measured current exceeds a predetermined threshold level. Consequently, the compensation provided by the control compensator (36) removes the compensation for compensating for solar pressure when the satellite (10) is in an eclipse. In a second implementation, the threshold logic circuit (48) is replaced with a proportionality logic circuit to compensate for the effects of partial eclipses.

Abstract: A vibration suppression system senses undesired mechanical motion of a body, and continuously drives a proof mass actuator system to reduce disturbances. An inertial or relative displacement sensor is processed to determine the energy content of the motion, and a correction signal is generated which is summed with the actuator drive signal to minimize the sensed motion energy. The position of the proof mass is sensed, and processed by differentiating and scaling, and the processed proof mass position signal is also summed with the actuator drive signal. The system causes the proof mass actuator loop natural frequency to tend to track the frequency associated with the maximum vibrational energy.

Abstract: A unified spacecraft attitude control system includes a memory aboard the spacecraft, in which a linear transformation matrix [.alpha.] is stored, which includes information identifying pseudo-complementary pairs of thrusters, and the characteristics of each pseudo-complementary pair. During each control cycle of the spacecraft attitude control system, the error signal is multiplied by a gain representing a desired slew rate to form pulse-width signals {pw} for the pseudo-complementary paired thrusters. An augmented pulse-width vector matrix {PW} is formed by transformations, to eliminate negative values of pulse width. The actual thruster pulse widths {.DELTA.t} are calculated as {.DELTA.t}=[.alpha.]{PW}. The thrusters are energized by limited values of {.DELTA.t}.

Abstract: A method of attitude control for spacecraft with flexible structures utilizes an estimator/state controller pair with on-board time-varying gain scheduling. The control method includes an attitude estimator (100) for each axis, which uses rate input from inertial reference sensors (4, 5, 6) to produce estimates (37, 38, 39) of each of the state variables. The estimator employs a predictor-corrector structure which computes initial rate and position estimates for each sample interval and corrects these values by weighing them with iteratively-calculated time-varying gains according to equations 35 and 36. The state controller (40) for each axis operates on these inputs, combining them with position and rate commands (41, 42) and weighing the results with time-varying gains calculated iteratively for each sample period according to equations 46, 47, and 48. The final result is a commanded control acceleration (50) which is forwarded to a thruster modulation logic.

Abstract: The three axes thruster modulation (8) of the present invention accepts three axes of input torque commands or angular acceleration commands and generates thruster selection and thruster timing (40) information which is used to fire thrusters (48) for the purpose of spacecraft attitude control and velocity change maneuvers. The modulation logic (8) works in all three axes simultaneously and is suitable for use with an arbitrary thruster configuration, including a configuration in which individual thrusters or thruster groups do not produce torques about mutually orthogonal axes. After thruster selection and on-times have been determined, the modulation logic (8) uses this information to compute a best estimate of the actual rate change (42) which is then compared to the commanded rate change (44) to develop a residual unfired rate change.

Abstract: A space transfer vehicle has a plurality of thrusters for maneuvering in six degrees of freedom about three orthogonal vehicle axes. The space transfer vehicle includes four box-like thruster modules which each house an identically arranged plurality of thrusters. The thrusters are selectively energized in pairs to produce torque about the three orthogonal axes with a first magnitude for maneuvering the space transfer vehicle without a payload, and of a second magnitude for maneuvering the space transfer vehicle with a payload.

Abstract: A propellant immobilizing system and method in which propellant motion in the propellant storage tanks (22) is reduced or eliminated during velocity change maneuvers to reduce disturbance torques acting upon the spacecraft (10) and improving the spacecraft attitude pointing performance. The thrusters (14, 16, 18 or 20) are fired to produce a small impulse in the direction of the desired Velocity change to begin motion of the propellant within the fuel tank in the opposite direction. After the fuel has moved within the tank to a location in which the propellant center of mass is aligned with the propellant tank (22) center of curvature in the direction of the desired velocity change, thrusters (14, 16, 18 or 20) are again fired to produce a force in the direction of the current velocity of the propellant center of mass to stop the propellant center of mass relative to the propellant tank (22).

Abstract: To reacquire the attitude of a satellite wholly or partially stabilized on three axes a test is executed to determine if a terrestrial sensor is sensing the Earth (test 1) and if a star sensor is sensing a star whose magnitude is at least approximately equal to that of a given reference star (test 2).* Phase a. If the results of tests 1 and 2 are positive, the Earth and the star are captured and the consistency of roll information supplied by the Earth and star sensors is checked: if the information is not consistent phase (b) is carried out.* Phase b: If the results of test 1 only is positive, the Earth is captured and the satellite is caused to rotate in yaw until the result of test 2 is positive. The reference star is captured and the phase (a) consistency test is carried out.* Phases c and d: If the result of test 1 is negative, the pitch speed is reversed for at most a given time. If the result of test 1 becomes positive, test 2 is carried out.

Abstract: Roll and yaw attitude control method for a satellite stabilized about its roll, yaw and pitch axes embodying a momentum wheel system generating a continuous angular momentum substantially parallel to the pitch axis and having a variable component at least approximately parallel to the roll-yaw plane and a continuously acting actuator system in which the roll and/or yaw attitude of the satellite is sensed. Control signals are applied to the momentum wheel system that are produced by a fast control loop using a known fast control law and second control signals are applied to the continuously acting actuator system that are produced by a slow control loop using a known slow control law. The continuously acting actuator system is loaded in fixed direction of the satellite parallel to the variable component if the latter has a fixed direction.

Abstract: A method and system for formationfinding and formationkeeping in a constellation of satellites (20) orbiting about a primary body (34) is provided. A set of satellite (20) destination position (22) and time goals is selected corresponding to the satellite's (20) station in a desired constellation. The satellite (20) senses actual destination arrival positions (32) and times relative to reference positions independent of the constellation and compares these values to values from the set of goals. Corrective maneuvers are selectively implemented responsive to this comparison.

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 scanning sensor having a radiation detector is mounted on a spacecraft or satellite orbiting the earth. The scanner is pointed in such a way with respect to the orbit plane of the satellite that the instantaneous field of view of the detector crosses the region between the lower and upper limits of the travel of a celestial body in a year in order for the radiation detector of the earth sensor to encounter the celestial body at least once per orbit. Electrical signals based on the horizon crossing and the presence of a celestial body in the field of view of the detector are generated and used to derive Yaw, Pitch and Roll attitude information for the satellite with respect to the earth.

Abstract: A method is disclosed wherein a flexible space craft may be slewed by the application of positive and negative acceleration forces applied about a slewing axis. The method is such that at the end of the application of the slewing forces there is no residual energy in the excited modes. By examination of the response of an undamped and then damped structural mode to a sequence of step forces--entirely within the premises of structural dynamics discipline--a minimum-time zero-residual-energy torque profile with unequal intervening pulses is arrived at heuristically. Rigorous yet simple relationships are then established among the maneuver angle of a rest-to-rest slew, slew time, widths of the intervening pulses, and natural frequency and damping of a critical mode whose energy at the end of slew must be zero.

Abstract: A method for controlling reorientation of a spacecraft's spin from a minor axis spin bias to a desired major axis spin after spin transition. A control system monitors rotational rates about the principal axes to detect a separatrix crossing of a polhode path therein. Controlled thruster firings resulting from spin rate information successively decrease and increase a characteristic parameter and capture the spacecraft during a spin transition to a desired major bias orientation. It is possible to monitor only .omega..sub.1 and .omega..sub.3 in an alternate embodiment.

Abstract: A method of sun search and acquisition is disclosed for a spacecraft (18) which is stabilized in three axes. The method utilizes three axes gyro rate and integrated rate sensing together with three simple slit type sensors (10, 20, 30, 40). The method starts from an arbitrary attitude and from body rates up to the limits of the gyro rate sensors and is comprised of a rate nulling step followed by two consecutive simple search procedures. One search procedure is about the pitch axis until alignment of the sun with the roll/pitch plane and the other search procedure is about the yaw axis until alignment of the sun with the roll axis. The sun acquisition culminates in pointing the roll axis of the spacecraft (18) toward the sun. By using slit type sun sensors (10, 20, 30, 40) versus a wide field of view sensor it is easier to find a mounting location on a spacecraft (18) which is free from spurious reflections off other parts of the spacecraft.

Abstract: A system identification apparatus determines a transfer function of an object to be identified for input and output signals. Predetermined frequency bands of the input and output signals are sampled through band-pass filters, and the transfer function is calculated according to a method of least squares. A characteristic part of the transfer function is found from the spectrum of the differential of the output signal, and passbands of the band-pass filters are determined to include the characteristic part.

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 space launch system and a space transfer vehicle usable within such system. The space transfer vehicle includes a primary propulsion engine and attitude control system. A guidance system for both the atmospheric launch vehicle and the space transfer vehicle is integral with the space transfer vehicle.

Abstract: A spacecraft thermal disturbance control system for controlling the torque exerted on the main portion of a spacecraft as a result of a thermal gradient being applied to a portion of the spacecraft that projects from the spacecraft's main portion. The spacecraft thermal disturbance control system includes a network of distributed temperature sensors located on the surfaces of the projecting portion of the spacecraft, a microcomputer operatively connected to the temperature sensors for receiving temperature information from the temperature sensors and a reaction wheel assembly operatively connected to the microcomputer. The microcomputer provides command signals to the reaction wheel assembly based upon temperature information received from the temperature sensors to cause the reaction wheel assembly to generate a torque that counteracts the torque exerted on the main portion of the spacecraft as a result of the thermal gradient applied to the projecting portion of the spacecraft.

Abstract: Apparatus and process for successive (stepwise) position control of a spacecraft undergoing precession, in preparation for release of at least one payload therefrom, are provided.In order to bring the longitudinal axis of the craft from an initial position to a predetermined position to be reached, the positions are referenced with respect to an inertial frame of reference, during each control step. The angular velocity of the craft is dependent on the variation between the position to be reached and the position of the craft at the end of the preceding control step and dependent on the velocity of the craft during the preceding control step.

Abstract: A surveying satellite apparatus having an on-board microprocessor to process sensor-provided data from planetary and/or celestial reference scene. The sensor data is compared with the on-board spacecraft database to determine if any misorientation or translation error is present. The spacecraft attitude and ephemeris solutions are autonomously updated to reflect the realtime alignment.

Abstract: The purpose integrated aerospike rocket engine and space vehicle aerobrake is to provide for optimum installation and performance of these systems in a space vehicle. By integrating the aerospike engine into the middle of the aerobrake, the engine can be located with its center of gravity toward the face of the aerobrake. This location permits the maximization of the displacement of the aerodynamic center of pressure of the aerobrake relative to the center of gravity of the space vehicle, thereby maximizing the stability of the vehicle upon aerodynamic capture at the earth or another planet. By substituting an aerospike rocket engine for a conventional bell nozzle engine, the doors required to close the aperture in the aerobrake to accommodate the conventional engines can be eliminated. In addition to the unique combination of the aero-spike engine and the aerobrake, an included unique feature of the invention is the extendable nozzle plug.

Abstract: The effect of orbit plane precession is used to place a plurality of satellites into one or more desired orbit planes. The satellites are distributed within each desired orbit plane in a selected configuration. The satellites are transported into orbit on one or more frame structures referred to as "pallets". When more than one pallet is used, they are placed on top of each other in a "stack". After the stack of the pallets has been launched into an initial, elliptical orbit, the pallets are separated sequentially from the stack at selected time intervals. Thrust is applied to transfer a first pallet from the initial orbit to a first, circular orbit, wherein the initial and first orbits are in planes that process at different predetermined initial and first rates, respectively.

Abstract: In a rendezvous maneuver for approaching a rendezvous spacecraft to a target spacecraft through a reference trajectory for approach which is different in altitude and period from a trajectory of the target spacecraft, when the rendezvous spacecraft is deviated from the reference trajectory for approach as a result from a failure in the maneuver, it is first maneuvered in the direction of altitude after 0.2 revolutions of the rendezvous spacecraft, and then maneuvered in the direction of phase after another 0.2 revolutions. The rendezvous spacecraft is thereby restored to the reference trajectory with no fear of entering a collision course, so that high reliable retry/recovery may be performed promptly and safely.

Abstract: A space launch system and a space transfer vehicle useable within such system. The space transfer vehicle includes a primary propulsion engine and attitude control system powered by the same storable bipropellant fuel and a guidance system for both the atmospheric launch vehicle and the space transfer vehicle. The space transfer vehicle is fuelable in space and has a primary propulsion engine which is configured for variable thrust burns.

Abstract: A method for controlling reorientation of a spacecraft's spin from a minor axis spin bias to a desired major axis spin after spin transition. A control system monitors rotational rates about the principal axes to detect a separatrix crossing of a polhode path therein. Controlled thruster firings resulting from spin rate information successively decrease and increase a characteristic parameter and capture the spacecraft during a spin transition to a desired major axis bias orientation. It is possible to monitor only .omega..sub.1 and .omega..sub.3 in an alternate embodiment.

Abstract: Selective reconfiguration of existing spacecraft on-board control and propulsion equipment is provided. The modification combines add-on switch means such as solid state multiplexers or relay banks and rate sensing gyros with existing accelerometers, propulsion control electronics and thrusters. The switch means reconfigure the input circuits and the output circuits of the propulsion control electronics. The switch means select the rate sensing gyro and the liquid apogee motors for control of spacecraft spin axis coning during perigee solid rocket motor firing, and reselects the accelerometer and reaction control thrusters for on-orbit nutation control.

Abstract: A method and apparatus for changing the orbit of an artificial satellite. The apparatus is caused to approach a target satellite and to be coupled thereto in space so as to constitute a dumbbell-like coupled system. The apparatus has a propulsion unit. The unit generates a thrust, whereby the velocity of the center of gravity of the apparatus increases, and the apparatus rotates around the center of gravity. As a result, the coupling system is placed in transition orbit.In the transition orbit, the coupled system is released at a timing when the direction of elongation of the coupled system becomes perpendicular to the orbital velocity vector. Subsequently, the target satellite is placed in a final target orbit, and the separated orbit changing apparatus is placed in an orbit different from the target orbit.

Abstract: A method to control the attitude in roll (X) and in yaw (Z) of a satellite including two solar generator panels adapted to be oriented independently of each other about a pitch axis. In a preliminary stage: two geometrical axes x and z are selected in the plane of the roll and yaw axes, there being associated with the z axis a tolerable command torque error much lower than for the x axis, and a correlation law is established between satellite panel depointing angles .gamma..sub.N and .gamma..sub.S and possible command torques due to solar radiation pressure.

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 attitude control system for a spacecraft includes sensors for generating attitude control signals and logic for producing torque demand signals in the form of a T.sub.d matrix. The thrusters are oriented in complementary pairs about the center of mass for, when energized, producing mutually opposite torques. Information about the location of thrusters and their thrusts may be represented by a pulse-width-to-torque transformation matrix C. A weighted pseudo inverse of C is precalculated asQ=WC'(CWC').sup.- (13)and is stored in memory associated with the flight computer, thereby avoiding the need for the flight computer to perform the intensive calculations of Q. Thruster pulsewidth for attitude control is easily calculated by the flight computer as.DELTA..sub.p =.tau..sub.p QT.sub.d (14)where .tau..sub.p is the control cycle time.

Abstract: A method and arrangement for aligning a space vehicle with respect to a reference object is disclosed. A measuring direction is obtained by determining the angular orientation direction of the space vehicle with respect to the reference object and subsequently the space vehicle is aligned by means of actuators such that the measuring direction corresponds to a reference direction on the space vehicle. The actuators are driven by control signals which are derived from the measuring direction to cause control torques on the space vehicle. The orientation of the space vehicle and reference object is determined only with respect to a single measuring direction along a main axis of a direction sensor.

Abstract: A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method.

Abstract: A method for East/West stationkeeping of a geostationary satellite maintains the osculating value of longitude from exceeding a specified deadband for a specified drift period between maneuvers. In the planning, the mean longitude, the mean drift rate and the mean eccentricity vector are calculated and then maneuvers are executed to maintain the values below a magnitude such that the osculating longitude will be within the specified deadband over a specified drift period. The target conditions are achieved through a plurality of maneuvers which initiate a period of free drift which lasts for a specified number of days. During the free-drift period, longitude remains within its deadband, and no additional maneuvers are needed or are performed. The method has the advantage that it can take into account limitations on thruster on-time by allowing for a generalized number of stationkeeping maneuvers.

Abstract: A stationkeeping method for a satellite in geostationary orbit comprises the steps of:determining at the same time the angle .alpha..sub.1 between the satellite-Sun direction and the satellite-Earth direction and the angle .alpha..sub.2 between the satellite-Pole Star direction and the satellite-Earth direction,deducing therefrom a state vector E consisting in orbital parameters by the formuls:Z=H.E+C.Bwhere:Z is a measurement vector the components of which are deduced from the angles .alpha..sub.1 and .alpha..sub.2,H is a measuring matrix,C is a bias sensitivity matrix,B is a bias vector determined beforehand by comparison of the measured vector Z and measurements made on the ground,determining stationkeeping manoeuvres and applying same by using thrusters.