Abstract: An attitude control system for a three-axis controlled spacecraft (1) in which the location of thrusters causes significant cross-coupling torques. A thruster command conditioning electronics module (TCCEM) (3) is positioned between the conventional roll, pitch, and yaw loop controllers (41-43) and the thrusters (61-63). The TCCEM (3) converts spacecraft-axis-based torque requirement signals (TDX, TDY, TDZ) emanating from the loop controllers (41-43) into scalar quantities (UL, UM, UN) that are input to thruster modulators (51-53) that turn on one or more thrusters of the thruster pairs (L,M,N). The TCCEM (3) causes the spacecraft (1) attitude to be adjusted as commanded by the torque requirement signals (TDX, TDY TDZ) despite the presence of the cross-coupling. Two embodiments are illustrated: a working embodiment in which cross-coupling torques are produced about two axes (Y,Z), and a more general embodiment in which cross-coupling torques are present about all three axes (X,Y,Z).

Abstract: A system located on board a geosynchronous satellite for detecting spin axis precession and developing thruster control signals for restoring the spin axis to the desired attitude. Fuel economy is enhanced by performing the major attitude correction at a time when the fuel expended contributes to that required for orbit inclination control.

Abstract: An axial electrostatic ion thruster for use on a dual-spin stabilized spacecraft is disclosed. The thruster is mounted on the spun portion of the spacecraft and continuously fired for an integral number of spin periods to provide north-south velocity and attitude control. To achieve zero attitude precession with a statically imbalanced despun platform, or an intentional attitude precession generally, the ion thrust is spin-synchronously varied between two non-zero levels to achieve a prescribed Fourier component amplitude and phase at the spin frequency. Modulation of the thrust is achieved with constant ion beam current by complementary modulation of the beam voltage, V.sub.B, and accel voltage, V.sub.A by varying the ion extraction system grid voltages. This varies the thrust while maintaining a constant total accelerating voltage V.sub.TOT =V.sub.B +.vertline.V.sub.A .vertline..

Abstract: A thruster for use in a spacecraft uses vaporized propellant in its combustion chamber (110). Normally two propellants are used of the kind used for main thrusters so that two vaporization chambers are needed. Pressure valves (108, 109) are controlled by electronic means (220) to maintain current pressure.

Abstract: A thruster system has a main bipropellant propulsion unit including a propulsion engine (50) and bipropellant supply having a tank (3) for an oxidizing agent and a tank (4) for a fuel. An auxiliary cold propellant thruster (21) uses propellant supplied by bleeding off vapor from the contents of the oxidizing agent tank (3).

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: An energy efficient solid propellant attitude control system for use on post boost vehicles equipped with post boost control systems, post boost propulsion systems, and velocity control systems utilizes gas from a gas generator to drive a gas turbine which drives an alternator to produce electrical power. The turbine effluent gas is selectively directed to highly vectorable hinged nozzles to provide attitude control for the post boost vehicle. Power from the alternator is used to drive electromechanical actuators which orient the highly vectorable hinged nozzles. The turbine effluent gas is also selectively directed to fixed thrusters and twin axial thrust motors.

Abstract: An automotive vehicle is subject to attitude control input impulses effective only with respect to a roll axis and a selected one of a pitch and yaw axis. A selected and controlled sequencing of roll attitude control allows attitude control also of both the selected one axis and the other of the pitch and yaw axis for the vehicle.

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: 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 and system is described for simultaneously precessing and nutation damping a spinning spacecraft that includes a computer, thrusters for precession and nutation damping and gyros aligned with or parallel to each of three orthogonal axes with the x-axis being the spin axis. The gyro information from three orthogonal axes is used to compute the momentum vector and target vector. The error vector is computed from the computed momentum vector and target vector. When the error vector's component along a given axis orthogonal to the x-axis changes sign, the thrusters are fired to apply corrective torques according to a computed control error.

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 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: Disclosed is a single control thruster having the expanded capability of operating at various thrust levels and more than one propellant fluid so that the number of control thrusters for a maneuvering vehicle can be reduced to one third from that now required for appropriate maneuverability in space, resulting in large cost savings and improved reliability. Also disclosed is a "plug in" package with control thruster which can be removed from the vehicle and replaced to make the vehicle easily maintainable.

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 novel propellant resupply system for, and method of recharging the propellant tanks of, the attitude control system (ACS) of a spacecraft or the like are described, comprising first and second propellant conduits respectively interconnecting the outlets of the fuel and oxidizer pump inducers of the primary propulsion system with the fuel and oxidizer tanks for the ACS thruster engines, and controller valves within each propellant conduit for controllably diverting fuel and oxidizer under pressure from the fuel and oxidizer pumps to the thruster fuel and oxidizer tanks during operation of the primary propulsion system engines.

Abstract: A digital system for controlling the attitude of a spacecraft (14), e.g., a satellite, with respect to three orthogonal axes. The system can control satellites (14) in parking orbit, transfer orbit, or final operational orbit, whether or not the satellite (14) employs one or more momentum wheels. A processor (2) converts weighted data from spacecraft sensors (12), representing angular orientation and angular velocities about each of the three axes, into thruster pulsetrain signatures to command the firing of each of six thrusters positioned about spacecraft (14), a positive and a negative thruster being positioned to impart both directions of angular momentum about each of the three axes. Each thruster pulsetrain signature is created once every processed error cycle period T1, and contains a varying number of pulses of varying widths. A negative feedback PWPF loop (2, 16, 20, 22, 24) is actuated for each of the three axes, preferably several times each T1 in order to achieve better thruster on/off resolution.

Abstract: An active nutation damping system for a spacecraft may comprise a nutation sensor controlling an actuator, a thruster say, via a control system so as to produce nutation opposing torques. Following the realization that performance limits on known systems may be due to a simplistic dynamic analysis which assumes the spacecraft to be a rigid body when it may in fact comprise one or more oscillatory elements, for example a fuel tank in which fuel is able to `slosh` about freely, the damping system herein is operable to identify two types of nutation component, for one of which appropriate opposing torque components are generated while, for the other type, torque components are generated that, at first sight, would appear to increase the nutation components but which, in fact, improve the overall nutation damping performance.

Abstract: A propulsion system and method is disclosed for controlling the attitude and drag of a space vehicle (10). A helium dewar (12) contains liquid helium (11) which cools an experiment package (22). The helium is heated or vented to keep the temperature between 1.5 and 1.7 degrees K to maintain helium boil-off gas adequate as a propellant without adversely affecting the experiment package which is contained in the helium dewar for protection from solar heating. The apparatus includes auxiliary heater (32) and temperature sensor (30) for controlling the temperature of the helium. The boil-off gas propellant is delivered to thruster modules (A) to control vehicle attitude and compensate for drag.

Abstract: A dual pressure solid propellant control system capable of operating at temperatures of approximately 3000.degree. F. and at multiple pressures. A solid propellant gas generator is connected to a plurality of valve clusters by a manifold, the valves and manifold being of a high temperature material and each valve being independently operable. A pressure feedback loop maintains the system pressure at a commanded value by effectively increasing or decreasing the gas exit area by varying the pulse duration modulation of the valves.

Abstract: Transients and steady state error induced in maneuvering a satellite due to a disturbance torque caused by thrust mismatch or differential in the alignment of thrusters with respect to the center of mass are minimized by introducing a torque balancing bias at the input of a thrust modulator of the thrusters prior to sensing position or attitude error. The bias may instantly off-modulate the thruster control signal to cancel the effects of attitude transients before errors develop. Other axes thrusters may be on-modulated instantaneously to compensate for cross-axis torque. The bias may be introduced into the satellite control scheme by manual ground control or in automatic on-board compensation based on stored parameters obtained for example from calibration measurements. Specifically, the torque balancing bias may be developed by reference to thrust mismatch detected and stored during previous maneuvers, thus anticipating expected attitude error without actual detection thereof.

Abstract: A space vehicle 10 is steered from outer space toward the earth's atmosphere main rocket nozzle end first. An inflatable braking member 28 is deployed around the vehicle 10 by introducing gas into it. The main rocket motor 16 is ignited to produce a substantially throttled stream of gases. This stream provides a cooling layer forwardly of the braking member 28 to protect the member 28 from heating by friction with the atmosphere. The reduction of the velocity of the vehicle 10 is controlled by directly varying the drag of the vehicle 10 to compensate for variations in the density of the atmosphere. The drag is varied by varying the thrust of the motor 16 and/or varying the shape of the member 28. A backup method is to deflate the member 28 in the atmosphere. The member 28 may be jettisoned after the vehicle 10 leaves the atmosphere to enter a low earth orbit.

Abstract: A spacecraft or transfer vehicle for transportation of a payload from a space shuttle parking orbit to a higher operation orbit, wherein the transfer vehicle utilizes housekeeping subsystems integrated into the payload, to avoid unnecessary duplication of these subsystems, for such functions as power supply, guidance, control, and communications. The transfer vehicle disclosed includes a propulsion system with at least one low-thrust bipropellant liquid rocket engine, to provide a "soft" or low-acceleration ascent, and employs two tanking stages, each having four propellant tanks, arranged symmetrically in a circular configuration. The first stage of four tanks is jettisoned during ascent, and the second stage, including the remaining four tanks, the rocket engine, and related supporting structure, is separated from the payload when the operational orbit is reached.

Abstract: A method for changing the spin rate of a spinning spacecraft having a dynamic balancing mechanism which utilizes the communications channel between the spacecraft and the receiving terminal, the master index pulse generator on the spacecraft, and the spacecraft axial thrusters. To change the spin rate, a dynamic imbalance of a known wobble angle and phase is introduced by the dynamic balancing mechanism. An axial thruster is then fired for a predetermined time interval thereby coupling a portion of the axial thrust force into the rotational energy of the spinning body.

Abstract: An attitude control system for spacecraft having thrusters utilizes the exhaust plume of the thrusters for attitude control. At least one of the thrusters is canted so that its reaction line of thrust, when misaligned with one of the spacecraft's primary axes; causes a control torque to develop about said axis which aids the plume torque about said axis resulting from the plume of the one thruster impinging on a surface such as that of a solar panel on the spacecraft.

Abstract: In a method of and apparatus for damping nutation of a spinning spacecraft (10), spin axis attitude disturbances are substantially reduced by controlling at least one nutation damping gas thruster (18) to fire with non-uniform gas pulses. During the beginning of a nutation control sequence, the duration of successive gas pulses is gradually increased (up pulsed) from zero to a predetermined maximum duration. The duration of successive pulses is then maintained constant for a time period. Finally, at the end of the nutation control sequence, the duration of successive gas pulses is gradually reduced to zero (down pulsed). Up pulsing of the gas thruster (18) is initiated in response to a predetermined maximum nutation angle measured by an accelerometer (20). Down pulsing of the thruster (18) is initiated in response to a predetermined minimum nutation angle. The number of variable and constant duration gas pulses delivered by the thruster (18) depends upon the mass distribution of the spacecraft (10).

Abstract: An active nutation control system for a space vehicle spinning about a geometric axis has thrusters for creating an external torque about an axis transverse to the first. The amplitude of the nutation and its phase in relation with the spinning movement are detected. A control arrangement receives signals from the nutation detectors, computes the changes in the angular direction of the angular momentum of the vehicle caused by thruster activation and energizes the thrusters when contemporaneously the amplitude of the nutation movement is greater than a predetermined threshold, the phase corresponds to maximum efficiency of said thrusters in reducing nutation momentum of the vehicle or at least does not increase it.

Abstract: A thrust unit for a space vehicle has a container in which an electrolyte is electrolyzed to produce combustible gases (for example, water producing hydrogen and oxygen). At a predetermined pressure the gases are ignited and combustion products are exhausted from the container through a thrust nozzle. The pulsed nature of the thrust unit operation, and the duration of each pulse, are such that overheating of the unit is avoided.

Abstract: A satellite, especially a geostationary satellite which may be stabilized in its three major axes and which has a spin wheel or momentum wheel supporting a position control system and which also has an engine or propelling system, is controlled by actuating the spin or momentum wheel at the latest when the satellite is in a phase of being transferred from one orbit to another orbit. Thereafter, the satellite is rotated about its axis which extends in parallel to the axis of the spin wheel or momentum wheel whereby the rotational impulse of the satellite is smaller than that of the momentum wheel. The stabilizing of the satellite relative to the three major axes is then made at the point in time when the satellite is to change from one orbit to another orbit whereupon the engine system for reaching the new orbit is switched on. The thrust of the engine system is so dimensioned that the position control system intended for a subsequent mission retains its orientation even during the phase of changing orbits.

Abstract: A controller for the attitude stabilization of a satellite. The controller odifies the action of a first control circuit by responding to the signals of attitude .phi. and angular velocity of attitude .phi. developed by the first control circuit. A second control circuit includes a comparator, two Schmitt trigger circuits, two AND elements and two OR elements. Signals from the AND elements and uncorrected thrust nozzle control signals from the first control circuit are supplied to the OR elements. The output of the OR elements control the thrust nozzles of the satellite.

Abstract: Apparatus for use in determining the orientation of a spinning spacecraft d for controlling the attitude motion thereof. In response to commands from an external processor, the apparatus enables a selected thruster at a commanded time delay after a solar-or earth-sensor reference pulse and fires the thruster a commanded number of times. The time difference between various sensor inputs is measured to determine spin rate and other data which is used by the external processor in calculating the firing sequence of the thrusters.

Abstract: Roll and yaw attitude errors introduced by orbit inclination deviations from the nominal orbit plane are minimized by sinusoidally varying the momentum produced by a transverse wheel mounted on the spacecraft. The wheel is mounted on the spacecraft such that its axis is parallel to the spacecraft's yaw axis. Sinusoidal variation of wheel momentum is obtained by sinusoidally varying wheel speed in response to a sine wave signal periodically updated from an earth station. In response to the sinusoidal variation of transverse wheel momentum, the spacecraft is rolled to minimize thereby the roll error introduced by the orbit deviation from the nominal orbit plane. Yaw error is minimized by providing sufficient transverse wheel momentum so as to maintain the total spacecraft momentum vector perpendicular to the nominal orbit plane.

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

Abstract: Spacecraft nutation caused by applying attitude and/or orbit control forces to a spacecraft along an axis which does not pass through the spacecraft's center of gravity is minimized by a signal responsive control system which first operates the attitude and/or orbit control forces for a predetermined time period and automatically after an appropriate non-operating time period, the attitude and/or orbit control forces are again operated for the same predetermined time period.

Abstract: An active attitude control apparatus for damping roll/yaw nutation motion within the control deadband of a satellite which is three-axis stabilized to local vertical by means of a pitch biased, body-fixed reaction wheel and offset roll-yaw thrusters. The apparatus is incorporated into a system in which large attitude motion outside the deadband is controlled by means of the offset roll/yaw control thrusters with an accompanying control mechanism for activating the thrusters in an appropriate direction when motion outside the deadband is sensed. The present apparatus includes logic circuitry apparatus to cause a second thruster activation, of the same sense, after each deadband actuated thruster firing, where such a firing in the same sense has not been triggered within a certain elapsed portion of the satellite nutation period, approximately 5/8ths of the period being preferred.

Abstract: An attitude control system for a space vehicle in which angular rate signals are generated by rate gyros mounted closely adjacent to gimbaled engines at the rear of a vehicle, and wherein error signals representative of a commanded change in vehicle angle or attitude are obtained from a precision inertial platform located in the nose region of the vehicle. The rate gyro derived signals dominate at high frequencies where dynamic effects become significant, and platform signals dominate at low frequencies where precision signals are required for a steady vehicle attitude. The blended signals are applied in a conventional manner to control the gimbaling of vehicle engines about control axes.

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

Abstract: Method of providing position, e.g. attitude control for space vehicles having control thrusters producing thrust for imparting momentum upon the vehicle, further having a control apparatus for the thrusters causing them to produce thrust bursts or pulses so that controlled, incremental momenta be imparted upon the vehicle, such thrust bursts being triggered respectively when the vehicle tends to move beyond selected descrete deviations and error position levels. The highest level, one in either direction, define a range of positional tolerance, and the thrusters are operated, so that a relatively small momentum is produced when the vehicle tends to move past either of said maximum levels; the thrusters are operated on the other levels under these conditions.

Abstract: A system for controlling the attitude of a vehicle along three-axes is provided with a simple momentum wheel of one degree of freedom and several jets so as to enable correction of attitude error and dampen nutation action of the vehicle. The jets are controlled by means of a two channel control circuit terminating in logic gates that select the particular jet required to make the error correction. One of the channels is responsive to the polarity or sign of a voltage provided as input thereto whereas the other channel is responsive to threshold voltage levels provided to it, both channels being responsive to inputs from a low pass filter. The threshold voltage levels are adjusted by feedback correction paths provided in that particular portion of the control circuit.

Abstract: Spacecraft nutation caused by applying attitude and/or orbit control forces to a spacecraft along an axis which does not pass through the spacecraft's center of gravity is minimized by a signal responsive control system which first operates the attitude and/or orbit control forces for a predetermined time period and automatically after an appropriate non-operating time period, the attitude and/or orbit control forces are again operated for the same predetermined time period.