Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: In order to control an attitude of a movable object having a flexible member (50) through an attitude maneuver, first, based on vibration of the flexible member at the time of the attitude maneuver, for example, a sampling function including no frequency components equal to or higher than a particular frequency is obtained. With the use of the sampling function, a control target value is created as a previously-frequency-shaping-type feedforward control law. Based on the control target value, attitude control data is created. The attitude control data can be used for the attitude maneuver with respect to the movable object.

Abstract: Methods and systems for controlling multi-body vehicles with fuel slosh are provided. In a method for stabilizing a vehicle with fuel slosh, a controller for a multi-body assembly is formed by selecting vehicle configuration and velocity as states of a rigid body. Equations of motion are written for the multi-body assembly, and velocity states are determined that are unactuated. A decoupling transformation is defined from the unactuated velocity states. The equations of motion are transformed using the decoupling transformation such that unactuated states are decoupled from actuated states. A slosh state estimator is coupled to the controller such that it is in operative communication with an input to the controller and an output from the controller.

Abstract: Methods and apparatus are provided for reorienting control moment gyros (CMGs) to compensate for CMG failure or change in spacecraft (S/C) mass properties or mission. An improved CMG comprises a drive means for rotating the CMG around an axis not parallel to the CMG gimbal axis. Releasable clamps lock the CMG to the spacecraft except during CMG array reorientation. CMGs arrays are combined with attitude sensors, a command module, memory for storing data and programs, CMG drivers and sensors (preferably for each CMG axis), and a controller coupling these elements. The method comprises determining whether a CMG has failed or the S/C properties or mission changed, identifying the working CMGs of the array, determining a new array reorientation for improved spacecraft control, unlocking, reorienting and relocking the CMGs in the array and updating the S/C control parameters for the new array orientation.

Abstract: A system for, and method of recovering a solar-powered spacecraft from an anomaly that renders the attitude of the spacecraft unknown includes maintaining a power-safe attitude by switching between two orthogonal axes using solar panel current sensors. The system and method may also include simultaneously determining spacecraft attitude using a star sensor. The system is applicable to spacecraft operating in a solar wing-stowed configuration.

Abstract: A spacecraft with a yaw steering system performing yaw steering. The yaw steering includes steering the spacecraft to have a yaw angle (?) for all sun elevation angles (?) in accordance with a yaw steering guidance law that provides a yaw steering motion about a yaw axis which is smooth for all sun elevation angles (?). At least part of the guidance law comprises a smoothing function (f), which is a function of an orbital position parameter (?) of the spacecraft, the smoothing function (f) smoothing the yaw steering motion for values of the orbital position parameter (?) where high rotational rates {dot over (?)} or high rotational accelerations {umlaut over (?)} would occur without the smoothing function (f).

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: In one aspect, a space vehicle includes a structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the structure. The structure includes at least one of an antenna mesh, a light-shielding mesh, an optical reflector mesh and a net. In another aspect, a space vehicle includes an antenna structure configured to expand from a first configuration to a second configuration and at least two equipment compartments attached to the periphery of the antenna structure. At least one of the at least two equipment components include a solar panel, a propulsion system and an antenna feed. In a further aspect, a space vehicle includes a payload element and at least two spacecraft-support structures attached to the periphery of the payload element. At least one of the at least two spacecraft-support structures includes a propulsion system and a tracking, telemetry and control system.

Abstract: A method for controlling an actuator of a vehicle comprises providing a dynamic condition sensor generating a vehicle movement signal and a position sensor for generating a reported position. A processor is coupled to the inertial sensor and the position sensor and comprises an estimator, a position measurement predictor having a filter, a comparator and a control shaping block, said estimator generating a vehicle position based upon the dynamic condition sensor, said position measurement predictor generating an estimated position measurement in response to the reported vehicle position and a matched frequency response to the movement signal, said control shaping block generating an actuator control signal in response to a comparison of the estimated position measurement and the reported vehicle position.

Abstract: A spacecraft attitude controller balances external torques, including those resulting from gravity gradient and those resulting from other orbital disturbances, to achieve a comparatively stable, neutral attitude or orientation. Torque is balanced by selecting spacecraft attitude Euler angles and angular rates such that orbital disturbances and cross-coupling inertial effects are cancelled by the external forces, based on Euler's equation. A spacecraft attitude torque-balancing controller and related method compares spacecraft attitude angles and angular rates with an orbit reference frame, and provides instructions to conventional momentum management and propulsion controls to responsively adjust the spacecraft attitude and angular rates. This feedback loop drives to zero (or an acceptably small quantity) the rate of change of the difference between spacecraft and reference attitude and angular rates, thus minimizing the net accelerations on the vehicle.

Abstract: A system for providing precision thrust and sun tracking attitude control is provided. The system determines a proximity region and alternately engages either an ideal operational mode or a predictive operational mode based on whether a thrust trajectory vector is in the proximity region in order to provide attitude control. The proximity region is determined based on an angle between the thrust trajectory vector and a sun vector. For example, the angle is about 20-30 degrees. The system engages the predictive operational mode when the thrust trajectory vector enters the proximity region. When in the predictive operational mode, the system periodically re-calculates the thrust trajectory vector and determines where the thrust trajectory vector will exit the proximity region.

Abstract: A ship includes a star tracker mounted on a platform stabilized in ENU by the inertial navigation system (INS). The line-of-sight (LOS) of the star tracker is directed toward two separated stars, and the LOS difference angles are noted. The angles are processed to generate vector triads representing geodetic (ephemeris) and navigation system attitude. The triads are processed to generate a coordinate transformation matrix. The transformation matrix is separated into systematic error and reference attitude error. The reference attitude error is summed with the inertial navigation system attitude to generate corrected ENU attitude. The corrected attitude is used as a reference for shipboard sensors, to reduce errors when the sensor data is linked to other platforms.

Abstract: A spacecraft system that includes a primary space vehicle and a secondary space vehicle, both of which are designed to optimize payload capacity and launch weight of the primary space vehicle. The primary and secondary space vehicles combine to form an on-orbit space vehicle capable of performing functions and maneuvers that exceed the physical capabilities of the primary space vehicle at the time of its launch. The spacecraft system is designed to minimize propellant containment-related disturbances while maintaining a standard level of functionality. The primary space vehicle is designed to be incapable of independently performing a propellant-intensive orbit change maneuver. Instead the primary space vehicle is designed to couple to a secondary space vehicle having propellant and thrust capability sufficient to perform an orbit change maneuver when the primary and secondary space vehicles are coupled.

Abstract: A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite's orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite's orbit.

Abstract: A structureless space telescope is disclosed. The structureless primary mirror includes a plurality of mirror elements for focusing and reflecting electromagnetic radiation, and the telescope further includes at least one secondary mirror and possibly a tertiary mirror for receiving the focused and reflected electromagnetic radiation from the structureless primary mirror and reflecting the electromagnetic radiation, and at least one focal plane array for receiving the focused and reflected electromagnetic radiation from the secondary mirror or tertiary mirror, wherein the plurality of mirror elements, the secondary mirror and the focal plane array are coordinated and controlled without a physical structure connecting the plurality of mirror elements.

Abstract: Apparatus, Systems, and Methods are provided for controlling motion of a spacecraft. One apparatus includes a non-contacting actuator and a passive mechanical system coupled in parallel with one another. A system includes a payload, a bus, and a hybrid actuator including a non-contacting actuator and a passive mechanical system coupled in parallel, and coupled between the bus and the payload. The system also includes an inertial actuator configured to maneuver the bus to maintain a relative position and/or attitude of the bus with respect to the payload. One method includes receiving a signal instructing a first controller to change the position and/or attitude of a payload and utilizing a hybrid system to change the position and/or attitude of the payload. The method also includes receiving the signal at a second controller and utilizing a system to change a position and/or attitude of the bus independent of the payload.

Abstract: The invention is a system and method for propellantless, ultrahigh precision satellite formation flying based on ultrahigh precision intracavity laser thrusters and tethers with an intersatellite distance accuracy of nanometers at maximum estimated distances of tens of kilometers. The repelling force of the intracavity laser thruster and the attracting force of tether tension between satellites form the basic forces to stabilize matrix structures of satellites. Users of the present invention can also use the laser thruster for ultrahigh precision laser interferometric metrology, resulting in simplification and payload weight reduction in integrating the thruster system and the metrology system.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: A method of determining the attitude of a spinning spacecraft is provided. The method includes stabilizing the spacecraft, initializing the attitude of the spacecraft using star tracker data, and estimating the attitude of the spacecraft.

Abstract: An integrated power and attitude control system and method for a vehicle efficiently supplies electrical power to both low voltage and high voltage loads, and does not rely on relatively heavy batteries to supply power during the vehicle initialization process. The system includes an energy storage flywheel, and a solar array that is movable between a stowed position and a deployed position. The energy storage flywheel is spun up, using electrical power supplied from a low voltage power source, to a rotational speed sufficient to provide attitude control. Then, after the solar array is moved to its deployed position, the energy storage flywheel is spun up, using electrical power supplied from a second power source, to a rotational speed sufficient to provide both attitude control and energy storage.

Abstract: A method and apparatus for maneuvering a satellite in orbit to alternately optimize the collection of solar energy and to take sensor data of terrestrial objects is disclosed The longitudinal axis of a large payload package is oriented perpendicular to the orbital plane to minimize the disturbance torque due to gravity gradient, and to allow simple rotation about the axis for attitude change between optimal Sun and optimal ground coverage.

Abstract: Momentum control is maintained in a geosynchronous orbiting spacecraft that uses a plurality of reaction wheel assemblies and a plurality of magnetic torquers to control the spacecraft momentum, each orbit of the spacecraft being comprised of a set of time steps, by determining a current momentum error for a current time step of a current orbit by adding a system momentum change determined for an immediately preceding orbit to an average system momentum determined for the immediately preceding orbit, and then subtracting a magnetic control torque momentum change determined for the immediately preceding orbit, determining a current duty cycle for each of the magnetic torquers based on the current momentum error and on a torque value applied by each magnetic torquer at each time step of the immediately preceding orbit, and commanding each magnetic torquer to operate at the current time step in accordance with its respective determined current duty cycle, wherein the magnetic torquers apply a magnetic momentum c

Abstract: Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through operation of a “Satellite Inspection Recovery and Extension” (“SIRE”) spacecraft which can be operated in the following modes (teleoperated, automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and other vehicles and objects in space, collectively defined as “target satellites”. The three basic types of SIRE proximity missions are defined as “Lifetime Extension”, “Recovery”, and “Utility”. A remote cockpit system is provided to permit human control of the SIRE spacecraft during proximity operations.

Abstract: A method for performing east-west station keeping for a satellite in an inclined synchronous orbit is described. The method includes averaging a value of a right ascension of the ascending node for an inclination vector associated with the satellite over a period of the control cycle, and managing corrections for the satellite such that an eccentricity vector, directed at perigee, is substantially collinear with the inclination vector.

Abstract: A system for providing attitude and antenna steering for a spacecraft is disclosed. The spacecraft has a number of reaction wheels and a number of antennas. The system includes control logic configured to: determine a beta angle, the beta angle being the angle between a sun vector and an orbit plane of the spacecraft, and alternately engage either a first mode or a second mode to provide attitude and antenna steering based on the beta angle.

Abstract: The disclosure relates to an APS based integrated sun sensor comprising: a diaphragm unit, a detection unit, a processing electronics unit and an interface unit. The diaphragm unit is operatively connected with the detection unit for forming a sunspots image. The detection unit is configured for outputting a gray value of each pixel. The processing electronics unit is operatively connected with the detector unit and the interface unit respectively, for evaluating an attitude angle on the basis of the gray value and coordinate of each pixel. The interface unit is operatively connected with a host computer, for transferring the attitude angle to the host computer. This disclosure has such merits as high accuracy, wide FOV (Field of View), low power consumption, low weight, small size and high update rate.

Abstract: A system (18) includes: a) A vehicle (12) includes an attitude or angular velocity control system (38), a plurality of star trackers or star sensors (22) each having a field of view (28); b) a memory (30) having a star catalog (32), a star pair catalog (58) and a reference table (56) stored therein; and c) a processor (24) coupled to the attitude or angular velocity control system (38), the star trackers or star sensors (22), and the memory (30). The processor (24) determines the vehicle inertial attitude or angular velocity or sensor alignment, based, in part, on the star pair catalog (58) and reference table (56). The design of the star pair catalog (58) and reference table (56) is suitable for rapid determination of attitude or angular velocity or sensor alignment, and an efficient use of memory.

Abstract: A sensing device is described that is configured to be launched from an air vehicle for deployment on the ground. The sensing device includes at least one sensor, an inertial measurement unit (IMU), at least one flight control surface, a flight control unit configured to control a position of the flight control surfaces, and a processing unit. The processing unit is coupled to the IMU and is configured to receive a desired trajectory from an external source and, upon launch of the sensing device, is further configured to determine an error between the desired trajectory and a current position as determined by the IMU. The processing unit is also configured to cause the flight control unit to adjust a position of the flight control surfaces to minimize the error.

Abstract: Pulsed detonation engines (PDEs) are adapted for use in reaction control systems (RCS), such as thrusters for orbital correction and control (e.g., for earth-orbiting satellites), divert thrust generation and control for space-based interceptor devices, and for missile trajectory correction and motion control. According to one aspect of the invention, PDEs are adapted for motion control of so-called “kill vehicles,” which are small devices, typically launched from satellites, for strategic missile defense.

Abstract: A method and apparatus for reducing centroiding error of a star sensor having a plurality of pixels is disclosed. The method comprises the steps of computing a star sensor angular slew rate of ? pixels per star sensor integration period ?, collecting star sensor data while slewing the star sensor according to the selected star sensor angular slew rate ?, and filtering the collected star sensor data according to a frequency determined by the selected star sensor angular slew rate.

Abstract: Method of and system for on-board substantially autonomous control for transferring a spacecraft from an initial orbit to a final geosynchronous orbit, by a trajectory that minimizes remaining transfer time and orbit transfer fuel. The spacecraft determines its orbit using a GPS-based system to determine the spacecraft orbital elements. Based on the measured orbit error, corrected co-state parameters are calculated and used to generate an updated thrust trajectory. The corrections are calculated using an innovative numerical procedure, carried out repetitively at a fixed interval until the target geosynchronous orbit is achieved.

Abstract: A spacecraft including control vanes for controlling the pointing direction of the spacecraft by providing a variable absorptive, reflective, emissive and/or transmissive properties. The control vanes provide a torque on the spacecraft to maintain the spacecraft pointing towards the sun. The control vanes can include at least one xb-axis control vane and at least one yb-axis control vane where the xb axis control vane provides xb-axis control torque on the spacecraft rotated about the xb-axis, and the yb-axis control vane provides yb-axis control torque on the spacecraft when rotated about the yb-axis.

Abstract: A vehicle (12) including a control system (18) is used for controlling vehicle attitude or angular velocity (38). The processor (24) is coupled to a star sensor or tracker (22) and a memory (30) that may include a star catalog (32), and an exclusion list (36). The exclusion list (36), a list of stars to be temporarily excluded from consideration when determining attitude or angular velocity or relative alignment of star sensors or trackers, is calculated on-board. Such a calculation prevents the necessity for a costly, periodic, ground calculation and upload of such data. By manipulating the star catalog, or sub-catalogs derived from said catalog, based upon the exclusion list (36), measurements of such excluded stars are prevented from corrupting the attitude or angular velocity or alignment estimates formulated on board.

Abstract: An integrated inertial stellar attitude sensor for an aerospace vehicle includes a star camera system, a gyroscope system, a controller system for synchronously integrating an output of said star camera system and an output of said gyroscope system into a stream of data, and a flight computer responsive to said stream of data for determining from the star camera system output and the gyroscope system output the attitude of the aerospace vehicle.

Abstract: The present invention provides a method for the attitude control of satellites in elliptic orbits or satellites initially placed in circular orbits perturbed to elliptic orbits due to environmental disturbances. The method relies on the application of solar radiation pressure to provide the desired torque for the satellite attitude control. The satellite is equipped with two-oppositely placed light-weight solar panels extending away from the satellite along a predetermined direction (satellite body fixed Y-axis). By rotating one of these solar panels or both of them through desired angles about their axis using the respective driver motors as per the simple open-loop control law, the torque about the satellite axis is developed to achieve the desired attitude performance.

Abstract: In order to control the attitude of a satellite having at least four gyroscopic actuators with respective spinners mounted on gimbals steerable about axes parallel to one or the other of only two different directions that are fixed relative to the satellite, the attitude of the satellite is measured using sensors on board the satellite, the control torque required to perform an attitude-changing maneuver is calculated, local linearization calculation is performed based on pseudo-inversion of the Jacobean matrix of the function associating the orientations of the actuator gimbals with the total angular momentum of the cluster in order to determine a new gimbal orientation, and precession speeds of at least one of the gimbals of the actuators are controlled to deliver the control torque for reaching the desired configuration.

Abstract: A method and apparatus for estimating spacecraft momentum is disclosed. The method comprises the steps of generating a plurality of spacecraft momentum measurements, fitting the plurality of spacecraft momentum measurements to a parametric model of a spacecraft momentum profile having a time period of tp, determining the momentum of the spacecraft from the parametric model; and generating an estimate of the momentum to be removed from the spacecraft at least in part from the determined momentum of the spacecraft.

Abstract: A method for yaw steering a spacecraft including performing yaw steering of the spacecraft to have a yaw angle (?) for all sun elevation angles (?). The method further including smoothing a yaw steering motion for orbital position parameters (?) where high rotational rates {dot over (?)} and high rotational accelerations {umlaut over (?)} would occur. The instant abstract is neither intended to define the invention disclosed in this specification nor intended to limit the scope of the invention in any way.

Abstract: A closed-loop LRF pointing technology to measure the range of a target satellite from a chaser satellite for rendezvous is provided that includes: LOS angle measurements of the target, a relative navigation Kalman filter, attitude determination of the visible sensor with gyros, star trackers and a Kalman filter, pointing and rate commands for tracking the target, and an attitude controller. An analytical, steady-state, three-axis, six-state Kalman filter is provided for attitude determination. The system provides improved functionality and precision for relative navigation, attitude determination, pointing, and tracking for rendezvous. Kalman filters are designed for the closed-loop system to allow for pointing the laser rangefinder to a target even if a visible sensor, a laser rangefinder, gyros and a star tracker are misaligned and the LOS angle measurements from the visible sensor are interrupted.

Abstract: A system (18) includes: a) A vehicle (12) includes an attitude or angular velocity control system (38), a plurality of star trackers or star sensors (22) each having a field of view (28); b) a memory (30) having a star catalog (32), an allocated area for a star pair catalog (58) and a reference table (56) stored therein; and c) a processor (24) coupled to the attitude or angular velocity control system (38), the star trackers or star sensors (22), and the memory (30). The processor (24) populates the star pair catalog (58), using the method described herein. The processor (24) then periodically determines the vehicle inertial attitude or angular velocity or sensor alignment, based, in part, on the star pair catalog (58) and reference table (56). The novel ability of the software to autonomously populate the star pair catalog (58) allows users to avoid uploading a large amount of data, and the problems associated with such an upload.

Abstract: A method and apparatus for estimating a slave payload attitude is disclosed. The method includes accepting a plurality of slave payload attitude measurements, deriving a model of the relative attitude of the slave payload and a master payload attitude at least in part from the plurality of slave attitude measurements, predicting the relative attitude between the slave payload attitude and the master payload attitude using the derived model, and estimating the relative attitude between the slave payload attitude and the master payload attitude at least in part from the predicted relative attitude between the slave payload attitude and the master payload attitude. Furthermore, the absolute attitude of the slave payload is computed using the relative attitude and the master payload attitude.

Abstract: A sensor alignment system and method is provided that facilitates the precise alignment determination of satellite sensors. The system and method utilizes an Inertial Measurement Unit (IMU) to facilitate alignment determination of multiples sensors on a satellite. The system and method performs a roll of the satellite around a preliminary sensor active axis, using the IMU to measure the rotation rate in both magnitude and direction. The sensor input is monitored by the IMU during the roll around the preliminary sensor axis. The data from sensor output obtained during the roll can be processed to calculate the difference between the preliminary sensor axis and the actual sensor axis. From this calculation the actual alignment of the sensor axis can be determined. Furthermore, by performing this alignment determination operation for multiple sensors the relative alignment between sensors can be determined.

Abstract: An inventive autonomous active manoeuvring method and system 1 for spinning spacecraft is provided having a capability to enhance the AOCMS performance of passive spinning satellites and to fulfil the emerging autonomy requirements applicable to new generation satellites. In broad terms, the invention resides in (a) the overall concept of providing autonomous execution of spin axis re-orientation manoeuvring for spinning spacecraft designed and executed autonomously on-board the spacecraft by the AOCMS and (b) in the proposed strategy set in place to execute the re-orientation manoeuvres with respect to the handling of residual nutation. Advantageously, the provision of coupling nutation avoidance manoeuvres with active nutation damping 8 on board the spacecraft reduces/minimises the manoeuvre settling time required to return the spacecraft to the steady state pointing performance, while not imposing constraints upon the particular spacecraft inertia sensor properties.

Abstract: A method and apparatus for refining a spacecraft state estimate, such as an attitude estimate or an angular velocity estimate, is disclosed. The method computes a plurality equations using residuals describing the difference between observed star positions and predicted positions based on inertial measurements, and solves those equations to generate refined estimates of the spacecraft state estimates.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method uses solar tacking and similar techniques to produce differential solar torques that are used to control the momentum and attitude of the spacecraft during the appendage deployment.

Abstract: A method and an apparatus for controlling the attitude and momentum of a spacecraft while deploying an appendage from the spacecraft. The method comprises the steps of predicting an environmental torque the spacecraft will be subjected to during deployment of the appendage, computing a magnitude and a direction of momentum to add to the spacecraft to at least partially oppose the predicted environmental torque, and storing the computed magnitude and direction of momentum in at least one of the momentum wheels before deploying the appendage.

Abstract: A system and method for performing in-orbit alignment calibration using on-board attitude sensors to improve reflector alignment after deployment to improve spacecraft pointing.

Abstract: A vehicle, such as a missile, with a pilot valve system controls the vehicle's thrust valves despite a hostile propellant gas environment. The pilot valve system can have one or more pilot valves. Using refractory elements, the pilot valve ball reciprocates between a supply seat and a vent seat which is subject to the filtered inflow of propellant thrust gases. When open, the pilot valve allows the stray thrust gas to communicate to a control chamber which closes a poppet against a valve seat in the nozzle. When an associated solenoid closes the pilot valve by pushing the pilot valve ball against the supply seat, the control chamber is vented to ambient. The poppet may then travel into the cylinder bore and the nozzle is opened to exhaust propellant gases and exert lateral thrust on the vehicle. Certain nozzle thrust geometries provide useful vehicle guidance.

Abstract: Controlling the solar torque imposed on a spacecraft (10) in flight by providing a film (52) with variable absorptive, reflective, emissive and/or transmissive properties on the sun side of a thermal shield (50) of the spacecraft (10). As the orientation of the thermal shield (50) changes relative to the sun line, the absorptive, reflective, emissive and/or transmissive properties of the shield (50) change to cause the shield's (50) center of solar pressure to change, thus aligning it with the spacecraft (10) center of mass (24) as viewed from the direction of the sun line. In accordance with another embodiment of the invention, the spacecraft (100) is provided with a plurality of control vanes (110-116) that have a variable absorptive, transmissive, reflective and emissive property to maintain the spacecraft (100) stably pointed towards the sun.

Abstract: A method and system for minimizing the solar array sun tracking disturbance is disclosed. The method separates the North solar array wing stepping and South solar array wing stepping in time, and use the disturbance caused by one solar array wing to cancel the disturbance caused by the other solar array wing. It separates the step times of these two solar arrays by a half cycle of the array frequency excited by the solar wing drive stepping. Because of the symmetry between the North and South solar arrays, their disturbances cancel each other out.