Abstract: A method and system of orienting a payload of an orbiting spacecraft (14) to maintain a desired pointing profile in the presence of orbit inclination. A cone (12) is determined which is traced in inertial space by a pitch axis of the payload to maintain the desired pointing profile throughout an orbit. A bias momentum vector of the spacecraft (14) is oriented at an attitude which lies along the cone (12). The attitude has a nonzero angle with respect to a plane spanned by an orbit normal vector and an equatorial normal vector. The payload is rotated about a single body-fixed axis perpendicular to the pitch axis to align the pitch axis along the cone (12). As a result, the desired pointing profile is maintained throughout the inclined orbit.

Abstract: A spacecraft includes a three-axis attitude control system. When velocity change thrusters are fired, their plumes impinge on a solar array, at angles which vary with the solar array position. This causes disturbance torques which vary with the solar array position. Disturbance torque information signals or torque bias signals which depend upon the solar array angle are summed with the torque demand signals which control the attitude control system during firing of the velocity change thrusters, to modify the attitude correction torques. The bias torque signals are generated by a Fourier processor based upon stored Fourier coefficients together with signals from a solar array angular position sensor.

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: A three-axis stabilized spacecraft includes a plurality of primary attitude control thrusters, the torque vectors of which lie in, or parallel to a primary plane. It also includes at least two more secondary attitude control thrusters, the torque vectors of which lie in a secondary plane which is not parallel to the primary plane. The control system produces attitude error signals, which are processed with a PID characteristic to produce impulse demand signals, all in known fashion. The impulse demand signals are transformed into an auxiliary coordinate system, in which two of the three auxiliary axes lie in the primary plane, and the third is orthogonal thereto. One of the secondary thrusters is selected, which has, along the third auxiliary axis, the largest torque magnitude and the same sign as the transformed impulse demand.

Abstract: An earth sensor for use in orbiting satellites over an altitude range between low earth orbit (LEO) and geosynchronous orbit (GEO) and beyond employs a combination of both visible wavelength senor technology and infrared (IR) sensor technology to result in high accuracy capability. During the vast majority of an orbit of the satellite, the sensor is operating in the visible region using reflected sunlight from the earth horizon. During the relatively small time period of the orbit when the sun is behind the earth and it is not possible to operate the sensor in the visible region, precise attitude information is generated by earth radiance balance operating in the IR region. The sensor therefore exhibits the known advantages and accuracy of operation in the visible region for a major portion of an orbit of a satellite, but also maintains attitude control of the satellite during a relatively small portion of the orbit when operation in the visible region is not possible.

Abstract: A method is provided for greatly increasing the ratio of direct sunlight to earthshine in a sun sensor used on a spacecraft which is orbiting the earth and is used for the attitude determination of spacecraft with respect to the sun illuminated earth. The sun sensor is provided with a detector in which radiation applied thereto is restricted to a spectral region where the earth's atmosphere is extremely absorbent. Therefore, the detector only receives radiation in that spectral region directly from the sun, while any reflected light from the earth in the selected spectral region is absorbed by the earth's atmosphere, thereby never reaching the detector. One suitable spectral region is the ultraviolet absorption band from 200-290 nm and a UV enhanced silicon detector would be used in the sun sensor.

Abstract: A spacecraft is controlled by a composite attitude signal including two components with different passbands. In one embodiment, a spacecraft attitude control system includes an attitude sensor and a controller which provides a time derivative function. High frequency noise components of the sensed attitude signal are enhanced by the derivative, and tend to cause attitude jitter or excess power consumption. The jitter is reduced by low pass filtration of the sensor signal, but this undesirably reduces the high frequency response of the attitude sensor. The high frequency response is restored by high pass filter coupled to a reaction or momentum wheel tachometer, which produces a signal representative of the high frequency components of the body rate. A summing circuit couples together the filtered attitude sensor signals with the high frequency components of the wheel speed signal to produce a relatively noise free broadband body rate signal.

Abstract: A spacecraft orientation procedure, in accordance with a first embodiment of the invention, is practiced with a sun sensor to bring the x (roll) axis of the spacecraft parallel to a ray of the sun, and with a gyro sensor and an earth sensor of the spacecraft in conjunction with one instruction provided either autonomously or by a ground tracking station regarding an orientation of a spacecraft reference plane to enable locating the earth by the earth sensor. Furthermore, in accordance with a second embodiment of the invention, the orientation is established without aid from the ground tracking station by use of at least one telemetry and command antenna having a continuous field of view, as measured in one plane, which is greater than a semicircle. In the second embodiment, the orientation procedure provides for rotation of the spacecraft about the x axis for a scanning of the antenna to intercept command signals broadcast from the earth, thereby to locate the earth in a first reference plane.

Abstract: A spacecraft (10) includes and attitude sensor (16) for generating sensed attitude signals. A controllable drive arrangement (238) is coupled to a reaction wheel (46), for driving it in response to speed error signals. The drive results in changes in reaction wheel speed. The spacecraft has a digital speed sensor (248) coupled to the reaction wheel, for generating quantized speed signals. A reaction wheel speed error signal generator (234) produces speed error signals in response to the difference between the attitude command and the sensed attitude signals, corrected by signals responsive to the speed of the reaction wheel, whereby at low reaction wheel speeds, the quantization causes attitude errors.

Abstract: Method and apparatus for calibration of a set of rate gyros on a three-axis stabilized satellite in orbit, the satellite including an attitude sensing system including the set of rate gyros, a set of actuators and control logic adapted to apply to the set of actuators control signals U derived from measurement signals M supplied by the attitude sensing system. From a time t.sub.0 satellite attitude errors (E1) relative to a reference frame of reference and constant drifts (E2) of the set of rate gyros about three axes (XYZ) of the satellite are estimated conjointly. The estimated values of the attitude errors and the difference between the measured speeds (.omega.) of the set of rate gyros and the estimated values of the constant drifts are applied in real time to the control logic and these estimated values of the constant drifts are stored when they become constant to within a first predetermined tolerance and the estimated values of the attitude errors become constant to within a second tolerance.

Abstract: A measurement arrangement is useful for controlling the attitude of a three-axis stabilized satellite equipped with sun-sensors for determining the orientation of the sun (sun vector) with respect to a satellite-fixed coordinate system, as well as with speed gyroscopes for detecting components of the satellite speed of rotation vector .omega.. It is necessary that the measurement range of the sun-sensors cover the round angle in a preselectable plane (for example XY plane) and perpendicularly thereto a limited angular range of maximum .+-..alpha..sub.2max on both sides of the plane. In addition, only an integrating speed gyroscope carrying out measurements in a single measurement axis that encloses with the plane an angle of at least (.pi./2)-.alpha..sub.2max should be provided.

Abstract: A method and system for determining the general three-axis attitude measurement for a spacecraft is disclosed. The system provides a cost-effective attitude measurement system for geosynchronous momentum bias spacecraft which need a continuously updated, moderately accurate, yaw measurement in addition to the usual roll and pitch measurements. The continuous yaw measurement enables control of the spacecraft yaw axis in the presence of disturbance torques. A sun sensor generates a signal which may be used to calibrate a continuously updated yaw measurement derived from the combination of signals generated by an Earth sensor and a space-to-ground link.

Abstract: An attitude sensing system utilizing simplified techniques and apparatus includes a Kalman filter which receives signals from an inertial measurement unit, a GPS receiver, and an integrated optical assembly. The output vector of the filter includes estimates of attitude misalignments and estimates of gyro drifts corresponding to the axes of the inertial measurement unit. The optical assembly includes a sensor array providing signals to the filter representing detection of the Earth's horizon or the center of the Sun. More particularly, a local vertical vector, computed from fore and aft detections of the Earth's horizon, is used in combination with GPS received signals to initially determine attitude by means of gyrocompassing. This attitude information is thereafter maintained by the inertial measurement unit and azimuth error resulting from drift of the inertial measurement unit and the initial gyrocompassing error is corrected by detections of the Earth's horizon and the Sun.

Abstract: A three-axis stabilized, earth-oriented satellite has an attitude control system with regulators, actuators, an earth sensor that carries out measurements along two axes and a sun-sensor arrangement that also carries out measurements along two axes. The field of view of the sun-sensor arrangement covers the round angle on a plane of the satellite-fixed coordinate system. Only the sun-sensor arrangement and the earth sensor act as measurement transducers. A sun and earth acquisition process for such a satellite has the following steps: seeking the sun; setting the sun vector in a first direction of reference; setting the speed of rotation of the satellite around the sun vector at a constant value; setting the sun vector in a second direction of reference, so that by rotating the satellite around the latter the optical axis of the earth sensor sweeps over the earth; and picking up the earth. Special regulating rules are disclosed.

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: The system includes a reflector assembly mounted on a first one of a pair of spacecraft. The reflector assembly includes a mirror around which are positioned a mirror and retroreflectors. A light source mounted on the second spacecraft illuminates the assembly the reflections from which are directed into a radiometer which is also mounted on the second spacecraft. The mirror has a known curvature, and the known curvature in conjunction with the retroreflectors allow calculation of the position and motion parameters of the first spacecraft relative to the second.

Abstract: A gyro platform assembly for determining the coning rate of a spinning and oning vehicle. A rotatable gimbal is attached to the vehicle such that the axis of rotation of the gimbal is in line with a spin axis of the vehicle. A rotatable platform, that supports three gyros, is attached to the gimbal. An axis of rotation of the platform is perpendicular to the axis of rotation of the gimbal. Each gyro has a rotation sensing axis, the sensing axes being mutually orthogonal, one such sensing axis being placed in line with the spin axis of the vehicle. A sensing axis, that is orthogonal to the spin axis of the vehicle, senses a coning rate of the vehicle. The output of the sensing axis is processed by a computer. The computer outputs a value equal to a sine of the coning rate of the vehicle.

Abstract: The attitude of a satellite is determined by using an array of IR detectors receiving an image of space and the Earth. The detectors are positioned so that at least one detector monitors the radiance from the Earth's horizon and a second detector monitors a region of the Earth adjacent to said horizon. The output of the first detector is compensated for variations in the radiance of the Earth by using the output of the second detector.

Abstract: A spacecraft for geodetic applications designed to travel along a trajectory in an orbital plane around a planet, embodying at least one cube corner retroreflector designed to face the planet at least temporarily and having an apex, a normal, and three faces which are substantially orthogonal with respect to one another, to within an arcminute, forming three dihedral angles at substantially equal angular distances from this normal, the effective diameter of this cube corner being at least 3 centimeters, the apex being located, relative to the center of mass of the spacecraft, at a distance whose projection on an imaginary line joining the center of mass to the center of the planet remains at all times at a known value to within variations of less than 5 centimeters.

Abstract: A method of controlling the attitude of a satellite according to which the direction of a predetermined celestial object is defined in a frame of reference related to the satellite, the instantaneous angular velocity vector of the satellite is detected and, by means of an actuating assembly, torques are applied to the satellite which are defined by a control law so as to rotate the satellite about the direction while orienting an aiming axis related to the satellite in the same direction, involves defining the direction of the predetermined celestial object in a frame of reference related to the satellite by a first quantity representing a first angle measured between an axis of sight and the projection of the direction onto a first reference plane containing the axis of sight and by a second quantity representing a second angle defined by the axis of sight and the projection of the direction onto a second reference plane containing the axis of sight, the second angle being calculated from the first angle and

Abstract: The attitude of a satellite is determined by using an array of IR detectors receiving an image of space and the Earth. The detectors are spaced so that at least in a central zone, the sun will not enter the optical field of more than one detector. The signals from each detector are monitored and if a signal is very high, indicating that the sun's image is falling on the detector, that signal is ignored.

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 of spacecraft attitude determination is proposed which calculates the bias between the usual two-cell and four-cell earth sensor measurements. During period of sun or moon interference, when the two-cell mode is selected, this bias is added to or subtracted from the two-cell measurements. This removes the step change in the calculated spacecraft roll and pitch angles when the switch is made between two and four cell modes and allows a substantially more accurate measurement to be obtained from a given sensor for the period of time for which interference would otherwise reduce sensor performance.

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

Abstract: A 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 star tracker 10 of an attitude control system 8 on board a spacecraft 9 is positioned so that the tracker's line-of-sight vector 12 intersects the origin of the reference axis but is askew from any reference axis of the spacecraft. The attitude control system transforms the dam from a star tracker reference coordinate system to a spacecraft reference coordinate system so as to allow control of the attitude of the spacecraft without modification of existing equipment.

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: At geosynchronous and other high altitudes a dual cone scanner has been a convenient way of determining both pitch and roll utilizing a single horizon sensor positioned on an orbiting spacecraft whose attitude is to be determined. However, for large roll angles, one of the scans may leave the earth or reference body defeating the purpose of the sensor. In accordance with the present invention, a third cone angle is scanned between the original two, thereby providing two good scans preserving the pitch and roll attitude information, even if one of the scans has rolled off of the reference body.

Abstract: A satellite attitude control system is usable in the absence of any inertial yaw attitude reference, such as a gyroscope, and in the absence of a pitch bias momentum. Both the roll-yaw rigid body dynamics and the roll-yaw orbit kinematics are modelled. Pitch and roll attitude control are conventional. The model receives inputs from a roll sensor, and roll and yaw torques from reaction wheel monitors. The model produces estimated yaw which controls the spacecraft yaw attitude. The model further produces estimates of the constant component of the disturbance torques for compensation thereof.

Abstract: In an optical navigation sensor used in a navigation system on a spinning spacecraft, a single CCD is positioned on an optical axis of an optical lens system at an image forming plane on which images of stars are formed by the lens system. The exposed surface of the CCD is divided into a first and a second region, the first region for dark images of stars, the second region for a bright image of a candidate planet. The second region of the CCD is implemented with a light shading film with a slit so that the second region is exposed by the bright image but with a light exposure reduced to an adequate level for the CCD. Alternatively, the second region may be provided with a light attenuation filter film. The CCD is driven by a single driver with a charge shift clock signal synchronous with the spinning rate of the spacecraft. The charges induced by the dark and bright images are time delayed and integrated to form an electronic signal of the images.

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: 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: This invention discloses a method for maintaining a desirable orientation of the solar wings of an orbiting satellite (10) relative to the sun. A dual array of sun sensors (26,28) is positioned on a body (12) of the satellite (10) in order to get a measurement of the position of the sun relative to the body (12) once per every orbit of the satellite (10). In addition, an estimate of the position of the solar wings (16,17) relative to the body (12) of the satellite (10) is attained. The sun-to-body angle is then subtracted from the body-to-wing angle to drive an error signal which is applied to a wing driver mechanism in order to maintain the solar wings (16,17) of the satellite (10) in a proper orientation relative to the sun.

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 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 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: 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 satellite attitude control system is usable in the absence of any inertial yaw attitude reference, such as a gyroscope, and in the absence of a pitch bias momentum. Both the roll-yaw rigid body dynamics and the roll-yaw orbit kinematics are modelled. Pitch and roll attitude control are conventional. The model receives inputs from a roll sensor, and roll and yaw torques from reaction wheel monitors. The model produces estimated yaw which controls the spacecraft yaw attitude.

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: 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: A horizon sensor for spaced-based satellites consisting of a high critical temperature superconductor which changes temperature based upon its exposure to space-based radiation. The horizon sensor may be flexibly positioned along the outer surface of the space- based satellite. As the orientation in space of the satellite varies, certain portions of the satellite body will be alternately exposed to radiation while other portions of the satellite body will be shadowed from it. As the sensor is exposed to radiation due to the change in orientation of the satellite body, the temperature of the superconductor changes due to radiation absorption. This change in temperature causes the conductivity of the superconductor within the sensor to vary, and this causes a change in voltage within the sensor. This voltage may be appropriately processed via land based or satellite based control systems to accurately measure and/or change the orientation of the satellite in space.

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 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 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.

Abstract: An attitude and nutation control system (30) for a momentum-biased vehicle (10) having roll, pitch, and yaw axes which employs a normal mode estimator (32) which predicts steady-state values for roll attitude, roll rate, and yaw rate. The normal mode estimator (32) receives instantaneous roll attitude information from an earth sensor (34) and optionally receives roll and/or yaw rate information from roll gyro (35a) and yaw gyro (35b). A logic circuit (36) coupled to the normal mode estimator (32) generates correction signals when the steady-state values for roll attitude, roll rate, and yaw rate are outside predetermined limits. A plurality of thrusters (14a-d) produce torque for bringing the steady-state values for roll attitude, roll rate, and yaw rate within predetermined limits.

Abstract: The attitude of a space vehicle is controlled to direct an instrument in a particular direction. For certain maneuvers, the reorientation may temporarily result in an undesirable orientation of the spacecraft. For example, the instrument may be a telescope which, during slewing of the spacecaft from one direction in space to another, may pass through an attitude in which it is directed toward the Sun. A control system is provided which performs the attitude control with an avoidance constraint. The control system generates a first signal representative of the current and desired final orientations, and slews the attitude in response thereto. A second signal is also generated which represents the difference between the current orientation of a vector associated with the spacecraft and a vector in space which it is to avoid. A threshold signal is generated when the difference between the avoidance vectors is a particular value.

Abstract: A method for acquiring Earth-pointing attitude of a three-axis, body-stabilized spacecraft orbiting the Earth in a prescribed orbit plane, e.g. a geosynchronous communications satellite, including the steps of aligning the roll axis of the spacecraft with the sun line (which is the vector directed from the spacecraft to the Sun); then, orienting the spacecraft such that the angle formed between the yaw axis and the sun line is equal to the Earth-Sun angle (which is the angle formed between the sun line and a vector directed from the origin of the spacecraft internal coordinate system to the Earth); then, orienting the spacecraft such that the yaw axis is aligned with the center of the Earth; and finally, rotating the spacecraft about its yaw axis until its pitch axis is oriented at a desired attitude relative to the orbit plane, e.g., normal to the orbit plane, to thereby complete acquisition of the Earth-pointing attitude.

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: 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.