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: A method and corresponding apparatus for providing spacecraft attitude, position and orbit data without the need for externally supplied data. Using onboard observations of the earth, sun and moon, the system determines spacecraft attitude, instantaneous position, and the orbit based on multiple position estimates. Position and orbit data are derived by multiple deterministic solutions, including some that employ star sensors and gyros, and the multiple solutions are accumulated in a Kalman filter, to provide continuous estimates of position and orbit for use when the sun or moon is not visible. The best estimate of position is selected from the multiple deterministic solutions and the Kalman filter solution, and can be used to control the spacecraft in various ways, without having to rely on ground-based equipment or other spacecraft for the determination of position and orbit.

Abstract: A three-axis attitude control system for a spacecraft in an equatorial orbit includes at least a horizon sensor for producing pitch-representative signals. A star sensor has a boresight parallel to the spacecraft pitch axis, and produces signals representative of the two-axis location of a selected pole star relative to the boresight. The star sensor signals are processed in conjunction with the pitch-representative signals and with orbit-location information, to produce roll and yaw information. The pitch, roll and yaw information controls spacecraft torquing. In a particular embodiment of the invention, the horizon sensor arrangement also produces roll signals, which are processed with the roll signals from the star sensor.

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 and apparatus for mitigating space debris having special applicability for use with anti-satellite weapons testing. To mitigate debris in a controlled and predictable fashion a first body is tethered to a lower altitude body. A suitable length of tether is chosen to correlate with the orbital characteristics of the higher altitude body such that the lower altitude body has a relatively low velocity for its orbital altitude. Upon release of the target from the tether or upon debris production by the lower altitude target, the reduced velocity of the body or debris produced for its altitude results in rapid deorbiting of the target or debris. To offset any momentum increase created when the debris is the result of an impact of an object with the lower altitude body, the orbital characteristics are controlled to ensure the proper deorbiting velocity is achieved. A counter rotation of the tether system can also be used to negate an increase in velocity due to impact.

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: A method for determining the attitude of a spacecraft without employing conventional attitude sensors, such as horizon sensors. A spin-stabilized spacecraft has an antenna mounted offset from the spin axis. The spacecraft is tracked over a substantial portion of its orbit from a ground (or tracking) station. A signal is transmitted from the spacecraft and received at the ground station. The spinning motion of the transmitting antenna causes the received frequency to oscillate sinusoidally due to Doppler shift acting on the transmitter-receiver range rate. The amplitude of the sinusoidal component of Doppler frequency shifts of the signals at several points along the orbit are used to determine the orientation of the spin axis. The attitude is adjusted as necessary, and thrusters are fired to move the spacecraft from the transfer orbit to the final orbit. The weight, cost and complexity of conventionl attitude sensors are eliminated from the spacecraft.

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

Abstract: The attitude of a satellite carrying terrestrial and solar sensors and spin-stabilized about an axis V--V is determined and controlled prior to its transfer from an elliptical transfer orbit to a circular geostationary orbit by means of an apogee motor firing. The attitude is defined relative to an inertial X-Y-Z frame of reference the Z axis of which is directed towards geographical North, by a declination .alpha. between the axis V--V and the X-Y plane and by a right ascension .beta. between the projection of the axis V--V onto the X-Y plane and the X axis. After injection of the satellite by a launch vehicle into the transfer orbit at its perigee, with an initial attitude approximating the predetermined final attitude for the apogee motor firing, the right ascension .beta. of the satellite is modified to confer on it an intermediate attitude such that the earth is in the field of view of the terrestrial sensors for a position of the satellite in the transfer orbit offset at least 10.degree.

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

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

Abstract: An attitude control system for a stabilized geostationary satellite comprising at least one terrestrial detector (T.sub.1) and a stellar detector oriented towards the North further comprises:a processing subsystem for each of at least the normal mode (I) and the stationkeeping mode (II), adapted to be connected to the detectors and to actuators (11, 12, 13) and comprising a pre-processing module (14-16) adapted to determine for each axis i an angular offset .epsilon..sub.i between a measured angle determined from the detectors and a reference angle together with a control and correction module (18, 20) adapted to determine for each axis a corrective torque of the form:c=k.sub.i..epsilon..sub.i +h.sub.i..epsilon..sub.iwhere k.sub.i and h.sub.i are parameters specific to each axis and to each processing line,a speed .epsilon..sub.i measuring unit (24,25),and a mode selection unit (26) connecting the detectors to a selected processing line and that processing line to at least one actuator.

Abstract: A device is disclosed for regulating a set value and/or for stabilizing elastic objects subject to spin and free to move around their axis of rotation, in particular aircraft and spacecraft. For that purpose, the regulating signals required to regulate the set value or to stabilize the freely moving objects can be obtained by means of a regulator having a transfer function with a denominator degree about three orders higher than the numerator.

Abstract: An Earth orbiting spacecraft (1) is returned to an Earth pointing attitude by means of a star sensor (6), Earth sensor (2), Sun sensor (4) and an angular momentum storage device. With the Sun sensor (4) locked on the Sun (5) the spacecraft is rotated until the star sensor (6) senses a predetermined brighter than magnitude 3 star such as Canopus (7). The angular momentum storage device is operated to preserve an attitude reference and the spacecraft (1) is further rotated by an angle .alpha. such that subsequent rotation around the angular momentum storage axis causes the Earth (3) to enter the view of the Earth sensor (2).

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

Abstract: An improved method for satellite station keeping is disclosed. The method provides for station keeping of a three axis stabilized satellite 10 using only two thrusters 20 and 22. The thrusters 20 and 22 can be ion or liquid propulsion thrusters and are mounted on the anti-nadir face 18 of the satellite 10. The north thruster 20 is canted at an angle .THETA. from the north-south axis 16 of the satellite 10 in a northern direction and the south thruster 22 is canted at the angle .THETA. from the north-south axis 16 in a southern direction. Both thrusters 20 and 22 are also translated to the east or west along an east-west axis 14 of the satellite 10 and swiveled at variable angles .alpha..sub.1 and .alpha..sub.2 respectively. A specific teaching of the invention discloses techniques for determining the angles .alpha..sub.1 and .alpha..sub.2 and the firing positions for the thrusters 20 and 22 to maintain the satellite 10 in a stationary orbit.

Abstract: A method for making a transition between a spin-stabilized spacecraft operating mode and a 3-axis stabilized mode includes the step of determining the orientation of the spin axis. A reference object detector such as a sun sensor produces a pulse as it crosses the reference, thereby completely specifying the attitude of the spinning spacecraft at that instant. Angular rate information is used to update the attitude information while a feedback system controls torques for slowing the spin and for assuming the desired attitude.

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

Abstract: One-dimensional sun sensing and angle measurement assemblies are deployed with substantially totally overlapping fields of view on a plane of a spacecraft body to protect sensors within a protected angular area of the plane, each of the analog sun sensors generating a signal indicative of incident sun along one axis and the presence of sunlight along an orthogonal axis thereto in the common plane. The sensors together generate an indication of the orientation of the sun relative to a target field of view which is a protected area. In an alternative embodiment, a digital sun sensor assembly is provided consisting of two or more independent digital channels and employing a slit aperture cooperating with a plurality of photocells laid out in a gray coded pattern. The digital device produces digital signals indicating relative position relative to axes in a plane.

Abstract: A stabilization system for a spacecraft (10) which enables the spacecraft to be stabilized by rotation about any principal moment of inertia axis. The system includes two pairs of control thrusters (14, 16, 18, 20) for producing positive and negative control moments about two orthogonal axes mutually perpendicular to the axis of rotation of the spacecraft. The control system (40) includes rate gyros (42, 46) which provide error signals which are sent to duty cycle modulators (54). The duty cycle modulators provide a pulse width modulated control signal to the control thrusters to approximate proportional control of spacecraft attitude.

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

Abstract: A method for achieving satellite (24) attitude stabilization against the undesired influences of solar pressure or other attitude disturbances to be used with earth-orbiting satellites of the type employing two solar collection panels (30a), (30b) extending from the satellite body in generally opposite directions (36), (38). Synchronized simultaneous movement of both solar panels about their normal sun-tracking position causes desirable torques to be generated which have a cancelling influence on the undesired satellite movement caused by solar pressure or other attitude disturbances.

Abstract: A method for controlling transition from thruster control to momentum wheel control to minimize nutation angle and transition time in a three-axis stabilized spacecraft employing an internal body fixed-momentum wheel as an attitude stabilizer. The method employs direct, substantially real-time measurement of momentum components and angular rates and generates paired roll thrust impluses and paired yaw thrust impulses in synchronization with the nutation period of the spacecraft to drive the spacecraft to a state of substantially zero nutation. The controller employs a deadbeat principle in a closed-loop real-time feedback system. In a preferred embodiment, the impulsive thruster firing commands in roll and yaw are simultaneous, spaced one-half nutation period apart in time with the effect that the target momentum is achieved in one to three pairs of impulse firings.

Abstract: In order to shift the moment of inertia H.sub.i of a body in rotation initially in rotation approximately about its principal axis of inertia Z into any given orientation H.sub.F one applies transversely to the axis Z three torque impulses in a plane containing H.sub.i and H.sub.F, the first for a time T.sub.1, such that, at the end of one rotation of a satellite about itself, the axis Z shifts to Z.sub.1 in the plane bisecting H.sub.i and H.sub.F, whilst the kinetic moment moves to H.sub.1, the second impulse at the instant when the principal axis of inertia is at Z.sub.1 for a time T.sub.2 appropriate to shift the kinetic moment into a position H.sub.2 symmetrical with H.sub.1 with respect to this bisecting plane, and the third impulse at the instant when the principal axis of inertia is parallel to H.sub.F for a duration equal to T.sub.1.

Abstract: A high precision linear actuator for compensating short stroke, high frequency vibrations in a space deployed structure comprises a set of cylindrical magnetic pole pieces affixed to respective ones of a pair of graphite flexures, that are mounted to a thermally conductive cylindrical housing, which is to be affixed to the space deployed structure. Reaction against the "proof mass" pole pieces is achieved by a magnetic field produced by set of stationary magnetic coils that surround the pole pieces and are energized by a vibration compensation reaction current supplied from a sense/control processor. Coaxial with and mounted integrally with the cylindrical pole pieces of the proof mass is an internal accelerater, which generates an output signal representative of the force to which the proof mass is subjected.

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

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

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

Abstract: An aerobraking orbital transfer vehicle (AOTV) (80) has aerobrake (82) with a blunted raked-off circular-cone configuration. The other components of the AOTV, including command/control module (95), fuel tanks (86, 88, 89 and 91), rocket engines (94) and afterbody (84), are positioned substantially along resultant force axis (104) of the AOTV (80). The axis (104) coincides with the resultant (sum of lift and drag) force vector. Afterbody (84) is mounted behind the aerobrake (82) with its length extending rearwardly from the aerobrake. The base flow clearance angle .phi. of the aerobrake (80) is 25.degree., thus allowing the afterbody (84) to extend rearwardly from the aerobrake (82) to a much greater extent than possible with a raked-off elliptic-cone aerobraking shield configuration.

Abstract: A linear force actuator system for stabilizing a support structure employs a linear dc motor whose primary winding is driven by a pulse width modulation control signal representative of a force input, through which the secondary member of the motor which acts as an inertial mass, is to be controllably translated. For improved control accuracy, the actuator system of the invention employs a pair of feedback loops, one of which monitors the current in the motor's primary winding to maintain a constant force output to the secondary member and a secondary of which monitors long term deviations from the center of the secondary member and corrects for centering offsets. An opto-electronic position sensing arrangement monitors the movement of the secondary member whereby precise control of commutation of the coils of the primary winding and a smooth translation of the secondary member are obtained.

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: An orbital system comprising a main platform and a satellite tethered by a long cable which may be used to perform measurements in the upper layers of the atmosphere. Air drag on the satellite causes the orbit parameters to decay, and means of regenerating the orbit parameters by electro-dynamic propulsion are provided. The cable is electrically conductive and, in a propulsion configuration, a current is passed through the cable to provide electro-geomagnetic thrust. Winches enable the cable to be wound in and out, either constantly, or in synchronization with angular oscillation of the system to damp the oscillations.On an east-bound orbit, the satellite may be transferred from a measurement configuration, where it is below the platform, to a propulsion configuration, where it is above the platform. This is achieved by winding the cable in synchronism to excite the oscillations until it swings above the platform and then to damp the oscillations so that it remains above the platform.

Abstract: An aerobraking orbital transfer vehicle which includes an aerobraking device which also serves as a heat shield in the shape of a raked-off elliptic or circular cone with a circular or elliptical base, and with an ellipsoid or other blunt shape nose. The aerobraking device is fitted with a toroid-like skirt and is integral with the support structure of the propulsion system and other systems of the space vehicle. The vehicle is intended to be transported in components to a space station in lower earth orbit where it is assembled for use as a transportation system from low earth orbit to geosynchronous earth orbit and return. Conventional guidance means are included for autonomous flight.

Abstract: An attitude control system for satellites includes a sensor (10) for generating output signals in response to variations in the attitude of the satellite. These signals are applied to a Kalman filter (12) which models the dynamic state of the satellite and instructs an attitude controller (13) to generate control signals for controlling the actuators. The system is responsive to the actual physical performance of the actuators to provide feed back to the modelling circuit.

Abstract: A passive step trimmer for a maneuvering re-entry body (MRB) which uses existing inertial forces acting on the MRB to deploy a free flowing mass from the rear of the MRB to trim the MRB altitude. The mass deployment from the MRB causes a forward shift in the center of gravity (CG) which in turn creates a more stable body which flies at a reduced angle-of-attack (.alpha.).

Abstract: The invention is a method of orienting a satellite in a geosynchronous orbit so as to direct a beam from an antenna mounted on the satellite to cover a desired target site, and pointing the beam axis at a desired bore site target. One embodiment of the method according to the invention includes a step of orienting the spin axis of a geosynchronous satellite at a direction angle which equals the inclination angle of the current orbital plane plus an adjustment or correction angle (.beta.), when the inclination angle is greater than zero. The correction angle is determined using one or both of the following equations:.beta.=BS-tan.sup.-1 {[ sin (bs-I)]/[6.61-cos (bs-I)]},or.beta.=.vertline.-(BS-tan.sup.-1 {[ sin (bs+I)]/[6.61-cos (bs+I)]}).vertline.,where, BS=the elevation angle of the bore site target in spacecraft coordinates, bs=the latitude of the bore site target, I=the inclination angle of the orbital plane with respect to the equatorial plane, and .beta.

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

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

Abstract: A method for passively stabilizing the attitude of a spinning orbiting body subject to orbital precession to maintain the spin axis orientation of the body essentially fixed and stable relative to the orbital plane. The preferred embodiment comprises selecting a body mass geometry and orbital parameters, including an orbit inclination, an orbital rate and a rate of regression of orbit line of nodes, so that the precessional motion of the body and the orbit are equal and opposite in direction and locating the spin axis of the body in a plane containing the north axis of the body and the orbit normal and between the north axis and the orbit normal, so that the gravity gradient precession of the spin axis equals the regressional motion of the orbit normal to produce a planar equilibrium configuration.

Abstract: A pointing apparatus for a dual-spin spacecraft utilizing a first sensor for sensing the time of arrival of an inertial attitude reference as the spinning portion of the spacecraft rotates, and a second sensor for sensing the time of arrival of an index reference which relates the position of the despun portion with the spinning portion. A digital processor estimates the spin rate and phase of the spinning portion from the inertial attitude reference time of arrival, estimates the relative spin rate and phase between the spinning portion and the despun portion from the index reference time of arrival, and estimates the bearing friction bias torque on the motor means which controls the pointing direction of the despun portion of the spacecraft.

Abstract: A spin-stabilized spacecraft has a pair of three-axis accelerometers (A, C) mounted on the spacecraft rotor (10) to provide for self-aligning, enduring and sophisticated navigational capability. A second pair of accelerometers (B, D) is added for redundancy and added sensitivity. Each pair of accelerometers are disposed at diametrically opposed positions on the rotor and respective axes are co-aligned. The orientation of spacecraft angular momentum h is determined by star and planet sensors.Radial alignment of each accelerometer relative to the spacecraft angular momentum is attained by zeroing non-radial accelerometer readings during spacecraft quiescence. Axial alignment is attained during periods of substantially axial thrust by isolating axial translational acceleration by summing readings across an accelerometer pair to cancel rotational, spin-periodic contributions to the readings and time-averaging or otherwise filtering to cancel other non-axial periodic contributions to the readings.

Abstract: A controller for effecting transistion of a spacecraft from one orbit to another without reliance on ground-based control stations. The controller as disclosed includes at least two conical earth sensors which are normally employed for detection and control of spacecraft orientation, to provide signals indicative of altitude. When the spacecraft altitude is below a selected threshold, the controller activates a rocket motor to raise the high point of the orbit. A second comparison, with a low threshold value of sensed angular diameter of the earth, indicates when a desired orbital altitude is first reached. The controller, and a related method for its use, enable a spacecraft to effect a transistion between orbits without complex on-board computers and without significant control from the ground.

Abstract: A spacecraft is adapted for orbital flight in two alternative modes, referred to as the earth-oriented and quasi sun-oriented modes. The spacecraft comprises a spacecraft body, at least one solar array extendible outwardly from the spacecraft body, passive attitude control means utilizing gravity gradient stabilization means which is extendible and retractable relative to the spacecraft body, and active attitude control means. The gravity gradient stabilization means is effective when extended to stabilize the spacecraft in the earth-oriented mode, and the active attitude control means is effective to stabilize the spacecraft in the quasi sun-oriented mode.

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

Abstract: One or more spinning flywheels are adapted to despin or partially despin a rotating mass such as a satellite and to convert the rotational energy into heat or electrical energy. The flywheels have one degree of freedom, parallel to the satellite, spin axis. This one degree of freedom allows the flywheel to precess in the manner of a gyroscope. As the flywheel precesses it turns a shaft which drives an electrical generator which produces electrical power that can be taken out of the despin device to drive external electrical devices that will disperse the rotational energy in the form of heat, light or electrically driven chemical changes. The flywheels exchange rotational energy after the spin vector of the spinning flywheel becomes congruent with the spin vector of the spinning satellite, thereby allowing the energy extraction process to be repeated. Bearing loss makeup energy is restored to the spinning flywheel system by feedback from the extracted satellite rotational energy.

Abstract: A closed loop system reduces pointing errors in one or more spacecraft instruments. Associated with each instrument is a electronics package (3) for commanding motion in that instrument and a pointing control system (5) for imparting motion in that instrument in response to a command (4) from the commanding package (3). Spacecraft motion compensation logic (25) compensates for instrument pointing errors caused by instrument-motion-induced spacecraft motion. Any finite number of instruments can be so compensated, by providing each pointing control system (5) and each commanding package (3), for the instruments desired to be compensated, with a link to the spacecraft motion compensation logic (25). The spacecraft motion compensation logic (25) is an electronic manifestation of the algebraic negative of a model of the dynamics of motion of the spacecraft. An example of a suitable model, and computer-simulated results, are presented.

Abstract: A system for controlling the velocity and attitude of an exoatmospheric projectile (10) that spins about a spin axis (z). Several thrusters (e.g., 1-7) are disposed along the outer surface (8, 9, 11) of the projectile (10) for performing three prescribed functions. The number of thrusters (e.g., 1-7) is minimized to save weight. As few as four thrusters (1-4) can be used to perform the three prescribed functions, which are: (A) reorienting the spin axis (z) in inertial space, by firing the axial thrusters (1, 2, 6, 7); (B) adding velocity to the projectile (10) in any direction, without a concomitant change in the orientation of the spin axis (z), by firing continually a combination comprising at least one of the thrusters (1-7); and (C) changing the projectile's spin rate (W), by firing one of the radial thrusters (3, 4). Nutation dampers, such as ball-in-tube nutation dampers (20), can be used to decrease the cone angle E and thereby improve the pointing accuracy of the spin axis (z).

Abstract: A system for controlling an attitude of an artificial satellite to decide the present attitude of the satellite on a spatial coordinate and angular velocities of rotation about body axes of the satellite, by the use of inertial sensors and star sensors, to calculate an objective attitude of the satellite on the basis of the present attitude information, the angular velocities of rotation about the body axes and orbit information received from a ground station so that the attitude of the satellite and the changing rate thereof are controlled simultaneously.

Abstract: Apparatus for enhancing the stability of a spacecraft (2) about a sensing axis (4). The spacecraft (2) may be of the three axis stabilized or spin stabilized variety. An attitude sensor (3) determines the offset angle (A) formed between a face (8) of the spacecraft (2) and an astronomical body (6), such as the earth. The sensor (3) produces an attitude signal (14) which is processed by compensation electronics (21) and fed to a torquing means (25) to close the angle (A) to within a desired preselected deadband (24).