Abstract: A spacecraft system including a spacecraft assembly or “stack” having an upper stage of a rocket-powered launch vehicle providing a final boost phase during launch. The stack also includes a payload structure rotatably interconnected with the upper stage. The upper stage and the payload structure together define a central axis that is generally coincident with the thrust axis during launch. The stack has an axis of maximum moment of inertia that is not parallel to the central axis. The stack has internal damping such that unstable nutation occurs if the upper stage and the payload structure rotate together about the central axis at the same rotational rate and in the same direction. The system includes a controller that rotates the payload structure relative to the upper stage during the final boost phase to alleviate coning motion of the stack.

Abstract: A prognostic health and fault management system and method that accurately determines the health and rate of degradation of one or more flywheel systems, by comparing actual and simulated responses of flywheel systems to supplied command signals. Based on the determined health status and degradation rate of the flywheel systems, one or more unhealthy, active flywheel systems are deactivated, and one or more healthy, deactivated flywheel systems are activated. The system can be used in terrestrial applications, or in vehicles such as satellites, spacecraft, or ships.

Abstract: A momentum control system and method is provided that provides attitude control for a vehicle while minimizing the negative effects of the momentum control system. The momentum control system and method include at least one more reaction wheel than the degrees of freedom under control. For example, in a vehicle designed to rotate in all three direction, at least four reaction wheels would be provided. The additional reaction wheel provide an additional control parameter that can be used to minimize the cost of the momentum control system's performance. The cost of the momentum control system that can be minimized includes the effects of vibration, power consumption, and undesirable changes in rotational direction, among others.

Abstract: A momentum control system and method is provided that provides attitude control for a vehicle while minimizing the negative effects of the momentum control system. The momentum control system and method include at least one more reaction wheel than the degrees of freedom under control. For example, in a vehicle designed to rotate in all three directions, at least four reaction wheels would be provided. The additional reaction wheel provides an additional control parameter that can be used to minimize the cost of the momentum control system's performance. The cost of the momentum control system that can be minimized includes the effects of vibration, power consumption, and undesirable changes in rotational direction, among others.

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 momentum wheel configuration that may be used on a spacecraft having a momentum bias that prevents zero wheel speeds. The momentum wheel configuration comprises three or more momentum wheels. The momentum wheels are oriented such that their wheel axes are not parallel to each other. The momentum wheels are oriented such that their axial momentum vectors are not coplanar. None of the momentum wheels have their wheel axis parallel to the momentum bias direction of the spacecraft. No two of the momentum wheels are oriented such that their axial momentum vectors are coplanar with the momentum bias direction of the spacecraft.

Abstract: An inertial control and measurement system is provided for applying inertial control to a vehicle and measuring angular rate of the vehicle. The system includes a control moment gyroscope having a spinning rotor, a gimbal, a gimbal support assembly attached to a spacecraft for allowing rotation of the gimbal about a gimbal axis, and a gimbal motor for rotating the gimbal about the gimbal axis to induce torque. A controller controls the gimbal motor to generate a control torque. The controller further determines an angular rate of the vehicle as a function of a determined torque and a gimbal angle acceleration.

Abstract: A satellite energy and attitude control system includes a first energy wheel, a second energy wheel and a third energy wheel. The first energy wheel is mounted in a first double gimbal and the second energy wheel is mounted in a second double gimbal. The third energy wheel is mounted in a third double gimbal and the first, second and third energy wheels, which have at least two degrees of motion orthogonal to their respective spin axes, are located within a single plane and positioned with respect to each other such that the total angular momentum provided by the first, second and third energy wheels sums to zero approximate an operational range center of the gimbals.

Abstract: Methods and structures are provided that enhance attitude control during gyroscope substitutions by insuring that a spacecraft's attitude control system does not drive its absolute-attitude sensors out of their capture ranges. In a method embodiment, an operational process-noise covariance Q of a Kalman filter is temporarily replaced with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. The error effect of the substituted gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range. In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.

Abstract: A reaction wheel system is provided that includes at least two rotors. The first rotor is the primary rotor that provides the large output torques to the vehicle. The second rotor is a vernier control rotor. The primary rotor and vernier control rotor each rotate about a common axis. The vernier control rotor has an inertial mass that is less than the inertial mass of the primary rotor, and rotates independently of the primary rotor. Because the vernier control rotor can be rotated independently from the primary rotor, it can be used to significantly improve the performance of the reaction wheel system. Specifically, the vernier control rotor is used to provide relatively small output torques. These relatively small output torques can be used to reduce the disturbances created by motor ripple, provide precise torque output control and/or reduce the disturbances created by static friction.

Abstract: An aircraft and associated method of manufacture and operation. The aircraft can include a fuselage having a first portion and a second portion projecting upwardly from the first portion, with the first portion housing a passenger deck and the second portion being positioned above the passenger deck. A first wing can extend outwardly from the first portion of the fuselage and the second wing can extend outwardly from the second portion of the fuselage, with the second wing being positioned above and forward of the first wing. Accordingly, the fuselage can include a plurality of passenger doors simultaneously accessible to ground-based passenger load/unload equipment with at least one of the passenger doors positioned beneath the second wing.

Abstract: A method and system for calculating a control function for a structural system (10) that can be used to determine an appropriate control force to apply to an active member (18) within a stationary member (12) on the structural system (10). An active member (18) and a stationary member (12) are defined as a two-mass system in which the active member (18) and the stationary member (12) move in opposite directions. The stationary member (12) is mounted to an isolation subsystem (14) that is composed of six isolators (28) at multiple degrees of freedom. The isolation subsystem (14) is softer than the stationary member (12), active member (18) and a spacecraft surface (16) due to a damping element (32) of the isolation subsystem (16). The isolation subsystem (16) is mounted to the spacecraft (16) and decouples the spacecraft (16) from the stationary member (12) and thus the active member (18). An accurate control force for the active member (18) can be determined based upon the above structure (10).

Abstract: An attitude control system (ACS) for a spacecraft includes an attitude control assembly (ACA) interface, a control moment gyro (CMG) array and a reaction wheel assembly (RWA) control unit. The ACA interface converts torque commands received from the RWA control unit into CMG gimbal rates for the CMG array. The ACA interface receives CMG gimbal angles from the CMGs and converts the CMG gimbals angles into RWA speeds, which are then provided to the RWA control unit.

Abstract: A method for operating a space vehicle equipped with at least one device cooperating with a celestial body, wherein rotation of the space vehicle takes place around at least one axis of the space vehicle. The rotation of the space vehicle is slowed when the device cooperating with the celestial body is oriented substantially toward the celestial body. The device can be a solar panel intended to face the sun and produce electrical energy.

Abstract: A satellite energy and attitude control system includes a first energy wheel, a second energy wheel and a third energy wheel. The first energy wheel is mounted in a first double gimbal and the second energy wheel is mounted in a second double gimbal. The third energy wheel is mounted in a third double gimbal and the first, second and third energy wheels, which have at least two degrees of motion orthogonal to their respective spin axes, are located within a single plane and positioned with respect to each other such that the total angular momentum provided by the first, second and third energy wheels sums to zero approximate an operational range center of the gimbals.

Abstract: The present invention provides a maneuver device for an artificial satellite, which causes small attitude error during maneuver and which requires a shorter period of setting time for obtaining a target attitude. The maneuver device is provided with: a feed forward torque instruction signal generator 8 which outputs a feed forward torque instruction signal 11 based on a maneuver plan; a thruster 10 which outputs control torque based on the feed forward torque instruction signal 11; and an attitude control signal calculator 6 to which an attitude angle and an angular velocity of the artificial satellite as well as a target attitude angle and a target angular velocity are input and which outputs an attitude control signal 13. The maneuver device is further provided with a disturbance compensating signal calculator to which the feed forward torque instruction signal 11 and a detected angular velocity signal 16 are input, and which generates and outputs a disturbance compensating signal 12.

Abstract: A dynamic unbalance compensation system is provided that compensates for dynamic unbalance of a rotating assembly on a vehicle, such as a spacecraft, to compensate for the presence of a dynamic unbalance moment. The system includes a vehicle, such as a spacecraft, a rotational assembly mounted on the vehicle and rotatable about an axis of rotation relative to the vehicle, and one or more momentum devices mounted on the rotational assembly and generating a momentum vector component perpendicular to the axis of rotation. The one or more momentum devices generate a compensation torque during spinning of the rotational assembly so as to compensate for dynamic unbalance of the rotational assembly.

Abstract: A hybrid stabilization system for isolating a pointing vector of a gimbal from the motion of a vehicle base is provided. The hybrid stabilization control system includes a rate feedback loop generating a rate feedback compensation value in response to a measured rate difference between a pointing vector rate of motion and a vehicle base rate of motion, a rate feedforward loop generating a rate feedforward compensation value in response to a measured inertial vehicle base rate of motion, a position feedback loop generating a position feedback compensation value in response to a measured position difference between a pointing vector angular position and a vehicle base angular position, a position feedforward loop generating a position feedforward compensation value in response to a measured inertial vehicle base angular position.

Abstract: The present invention provides methods of and apparatus for determining the inertial attitude of an aerospace vehicle. In one embodiment, the invention provides a rotational astronomical object-sighting concept to determine the inertial attitude of an axis of the aerospace vehicle without the star identification or dragback. In another embodiment, the invention provides an attitude measurement apparatus comprising a high sensitivity optical sensor and a low power inertial sensor.

Abstract: A momentum wheel configuration that may be used on a spacecraft having a momentum bias that prevents zero wheel speeds. The momentum wheel configuration comprises three or more momentum wheels. The momentum wheels are oriented such that their wheel axes are not parallel to each other. The momentum wheels are oriented such that their axial momentum vectors are not coplanar. None of the momentum wheels have their wheel axis parallel to the momentum bias direction of the spacecraft. No two of the momentum wheels are oriented such that their axial momentum vectors are coplanar with the momentum bias direction of the spacecraft.

Abstract: A gyrostabilizer without a physical shaft or axle is constructed. The stabilizer has dual counter-revolving concentric rings filled with weights such as spherical balls that are propelled in orbital fashion by fluid pressure or electromagnet propulsion. The diameter of the concentric ring can vary from a few inches to more than ten (10) feet. Without an axle or shaft the weight of the gyrostabilizer is shifted to the perimeter where most of the momentum is generated at a fraction of the weight of gyrostabilizers that spin on an axle. Also without an axle or shaft the space between the center of revolution and the revolving balls is usable or void. When properly mounted in a vehicle or structure unsteadiness such as tremors, vibrations, sway, pitch, roll, yaw can be dampened.

Abstract: A spacecraft with a reaction wheel system can be autonomously safed by setting the solar wings to continuous tracking, determining a slew rate vector based on the total angular momentum, and slewing the spacecraft using the slew rate vector until commanded to stop autonomous safing. When the total angular momentum of the spacecraft to large to be handled by rotisserie, then the spacecraft is reoriented to align a suitable rotation vector with the system momentum. In a typical application, the spacecraft has a reaction wheel assembly with four wheels arranged to form a right regular pyramid. Two reaction wheels on opposite edges of the pyramid form a first pair and the two remaining reaction wheels forming a second pair.

Abstract: A method of controlling the attitude of a vehicle in space relative to three reference axes, the method consisting: initially in adjusting the attitude of the vehicle in conventional manner to desired angles on the basis of a two-axis detector which points towards a reference heavenly body; and subsequently in adjusting the attitude of the vehicle about the third reference axis by using the speeds of rotation of the reaction wheels which reflect conservation of the total angular momentum of the vehicle to perform an angle measurement about the third axis.

Abstract: The present invention provides an attitude control actuation apparatus and method to perform reaction wheel desaturation. Two or more reaction wheels are used for each of the three axis. The two reaction wheels are used to perform a simple cost effective electric powered desaturation or slowing of the reaction wheels by exchanging momentum with each other.

Abstract: A spacecraft attitude control system that enables friction and stiction compensation for attitude control of a spacecraft using at least one reaction wheels that can spin through zero rpm without degradation in pointing. The system removes nuisance disturbances relating to friction or stiction that detract from the ability of the system to maintain or change spacecraft attitude. It can remove these effects before they are sensed by the systems attitude control sensors. A friction or stiction compensation buffer is arranged between the spacecraft attitude control subsystem and the torque actuators for the at least one reaction wheel. Since the compensation is localized to each wheel, it is possible to optionally: control wheel operating speeds to safe levels, provide wheel failure flags, provide accurate wheel tracking at slow or zero wheel speeds and provide full access to the wheel torques.

Abstract: Reaction wheels are placed around an enclosure so that the reaction wheel rotors provide a radiation shield for electronic components inside the enclosure.

Abstract: A satellite attitude control system includes a programmed processor system which includes a gyroscopic actuator first control stage for changing the attitude of the satellite and a reaction wheel second control stage for assuring that pointing of the satellite is accurate and stable. The method is intended to be used for a satellite including the two control stages indicated above, which it uses selectively for the operations indicated above.

Abstract: A device for measuring a velocity and/or position of a body includes a magnetometer and a measurement unit. The magnetometer is configured to sense a magnetic flux to which the body is exposed. The measurement unit includes a sensor sensing one of an acceleration and a velocity of the body. The measurement unit also includes a processor coupled to the magnetometer and to the measurement unit and configured to determine one on a velocity value and a position value for the body based on the sensed magnetic flux and based or one of the sensed acceleration and the sensed velocity of the body.

Abstract: A method for operating a space vehicle equipped with at least one device cooperating with a celestial body, wherein rotation of the space vehicle takes place around at least one axis of the space vehicle. The rotation of the space vehicle is slowed when the device cooperating with the celestial body is oriented substantially toward the celestial body. The device can be a solar panel intended to face the sun and produce electrical. energy.

Abstract: An integrated system for providing 3-axis attitude-control, energy-storage, and electrical power in which two counter-rotating inertia-wheels are tandem mounted with rotational axis oriented on and parallel to the longitudinal-axis of an airframe. A horizontal- and vertical-translating support is centrally-located between the wheels and houses a common wheel-bearing mount communicating with and supporting both inward-pointing inertia-wheel shafts. The preferred embodiment further consists of two outwardly-mounted tilt wheel bearing and suspension means, each communicating with and supporting an inertia-wheel outward-pointing shaft. Taken together, these supports and their associated two-axis actuator enable both wheels to be tilted equally in opposite directions to generate a combination of pitch and yaw reaction torque. Further, an electric generator integrally mounted to each inertia-wheel, and an electric load controller connected to each electric generator enable the creation of reaction roll torque.

Abstract: Systems and methods that limit the effects of actuator saturation to certain body axes of a spacecraft. A plurality of actuators, or momentum wheels, disposed on the spacecraft are oriented non-parallel to x, y, and z axes of the spacecraft. The plurality of actuators each produce torque that is applied to the spacecraft. Outputs of the plurality of actuators or momentum wheels are processed to control the respective actuators or momentum wheels to manage actuator saturation so that saturation does not produce attitude errors about one or more predetermined axes of the spacecraft.

Abstract: An antenna deployment system for use in storing and deploying three antennas that are located on the same side of a spacecraft or other fixed body. The three antennas are nested and are stacked in a stowed condition and are individually and sequentially deployed to their respective deployed positions. One or more feed horns are attached to the spacecraft or fixed body that illuminate the respective antennas. A dual axis deployment mechanism is used to deploy each antenna. The dual axis deployment mechanism is also used to steer the beam produced by the antenna. The dual axis deployment mechanism comprises a dual-axis rotatable hinge structure affixed to the spacecraft or fixed body that is coupled to the antenna by way of a substantially rigid reflector support structure. The dual axis deployment mechanism is actuated and controlled to deploy the antenna and steer the antenna beam.

Abstract: An integrated system for providing 3-axis attitude-control, energy-storage, and electrical power in which two counter-rotating inertia-wheels are tandem mounted with rotational axis oriented on and parallel to the longitudinal-axis of an airframe. A horizontal- and vertical-translating support is centrally-located between the wheels and houses a common wheel-bearing mount communicating with and supporting both inward-pointing inertia-wheel shafts. The preferred embodiment further consists of two outwardly-mounted tilt wheel bearing and suspension means, each communicating with and supporting an inertia-wheel outward-pointing shaft. Taken together, these supports and their associated two-axis actuator enable both wheels to be tilted equally in opposite directions to generate a combination of pitch and yaw reaction torque. Further, an electric generator integrally mounted to each inertia-wheel, and an electric load controller connected to each electric generator enable the creation of reaction roll torque.

Abstract: A method of controlling the attitude of a vehicle in space relative to three reference axes, the method consisting:

Abstract: The invention provides a momentum wheel energy storage device for a spacecraft or satellite. The device has six wheels whose axes of rotation are parallel to edges of a tetrahedron. It thus provides redundancy which makes it possible simultaneously to store energy and to generate an arbitrary net angular momentum, even in the event of one of the wheels failing. The failure of one wheel can be accommodated without degrading operation, with energy storage being optimal regardless of the angular momentum to be generated. The failure of two wheels can be accommodated but operation may then be in degraded mode. The invention makes it possible to limit the number of momentum wheels used in a device that can accommodate failures.

Abstract: A system for managing electrical energy and for controlling attitide of a satellite has an attitude control unit which controls a plurality of reaction wheels oriented about different axes for instance three wheels. An on-board electricity power supply network has a solar array, a payload and a storage unit connected to a power line. The storage unit comprises a buffer supercapacitor and one inertia wheel unit with a respective motor/generator and typically two contra-rotating wheels. Interchanges between the storage units and the power supply line are controlled by a regulator which attenuates power demand transients and thus torque transients on the inertia wheel or wheels.

Abstract: Methods and systems for stabilizing satellite spin about an intermediate inertia axis (Z) are disclosed. A set of gyros (22) sense the X component and the Y component of the angular velocity of the satellite body. A single degree of freedom momentum wheel (26) has a fixed transverse orientation with respect to the intermediate axis in order to store momentum. In one embodiment, the momentum wheel (26) is oriented to store momentum parallel to the Y axis. A tachometer (30) senses the rotation rate of the momentum wheel (26). A processor (24) forms a control signal representative of a control torque to be applied to the momentum wheel (26). The control torque is based upon the X component and the Y component of angular velocity of the satellite, and the angular velocity of the momentum wheel (26).

Abstract: A device for both sensing an angular rate of an object and applying an angular rotation torque to the object. The device includes a rate sensing mechanism, such as a fiber optic gyroscope, which senses the angular rate of the object. The device also includes a reaction/momentum wheel mechanism, including a flywheel, for applying the angular rotation torque to the object. The rate sensing mechanism is contained within the reaction/momentum wheel mechanism or attached thereto such that the input axis of the rate sensing mechanism is coaxial with the spin axis of the reaction/momentum wheel mechanism.

Abstract: A momentum management system for attitude control of a spacecraft includes a housing to be fixed to the spacecraft and a momentum wheel rotor in the housing for storing angular momentum. A gimbal assembly mounts the rotor in the housing. The rotor is driven by a drive with its output coupled to the rotor. A torque generation imparts torque to the rotor about axes orthogonal to the drive axis. The gimbal assembly includes a gimbal ring coupling the drive output to the rotor. The gimbal ring in turn includes flexure joints connecting the gimbal ring to the drive and the rotor. The flexure joints are configured to permit the rotor to tilt about two flexure axes orthogonal to the drive axis to incline the rotor axis through a range of angles from about 0 degrees to about 7 degrees with respect to the drive axis under the control of said torque generation device. The preferred flexure joint is formed from two resilient, crossing webs.

Abstract: An agile steering control arrangement is provided for a spacecraft having a skewed scissor pair control moment gyro (SSP CMG) attitude control arrangement gimbaled with respect to the spacecraft body. The control arrangement includes a steering control process which converts arbitrary spacecraft boresight commands into spin torque and gimbal rate commands for a pair of spinning CMG. The steering control process generates suitable CMG torque commands based on requiring that the boresight nominally follow an arbitrary boresight trajectory described by required angular accelerations for the boresight together with one of various requirements on yaw motion. In this manner, a SSP CMG actuated spacecraft can be steered having tight boresight control while allowing a certain amount of error in yaw about the desired boresight. If spin motor capacities are exceed, new gimbal rates are determined based on saturation wheel spin torque.

Abstract: A method for estimating motion of a dual-spin spacecraft having a gimballed momentum wheel. The method disengages the gimbal from its drive train in anticipation of a short disturbance. Then, the gimbal slip resulting from the disturbance is measured. The method may be used to stabilize any dual-spin spacecraft that uses a gimballed momentum wheel. To stabilize the dual-spin spacecraft, the gimbal angle and gimbal rate are measured during and after the disturbance to provide an indication of the inertial spacecraft motion along gimbal axes. The magnitude and direction of the disturbance are determined by comparing motion of the gimbal before and after the disturbance. Then, torques are applied to the gimbal to counteract the spacecraft motion resulting from the disturbance.

Abstract: A moment control device for positioning a spacecraft which employs a plurality of spinning bodies operable to impart a desired torque to a space craft, the bodies being constructed in a unitary combination and the unitary combination being mounted to the spacecraft to be positioned.

Abstract: A Stationkeeping Mode for a spacecraft having a gimballed momentum wheel for attitude control in which the stationkeeping is divided into short duration open-loop stationkeeping maneuvers to correct orbital errors, and interlaced with closed-loop momentum management operations to manage momentum stored in the momentum wheel. The sequence of open-loop pulses are spaced to deadbeat a given dominant flexible motion of the spacecraft. Between and during the thruster pulses, the gimbaled momentum wheel is maintained in the closed-loop feedback control using gyro sensor or other inertial reference measurements. Limitations on the range of the momentum storage device and its torque capability are minimized by combining the maneuver thruster pulses with pulses from other thrusters, or by reducing the fire time of one of the maneuver thrusters, to minimize momentum buildup.

Abstract: A spacecraft with a reaction wheel system can be autonomously safed by setting the solar wings to continuous tracking, determining a slew rate vector based on the total angular momentum, and slewing the spacecraft using the slew rate vector until commanded to stop autonomous safing. When the total angular momentum of the spacecraft to large to be handled by rotisserie, then the spacecraft is reoriented to align a suitable rotation vector with the system momentum. In a typical application, the spacecraft has a reaction wheel assembly with four wheels arranged to form a right regular pyramid. Two reaction wheels on opposite edges of the pyramid form a first pair and the two remaining reaction wheels forming a second pair.

Abstract: A control method and improved equatorial-normal body-stabilized spacecraft that operates in an inclined orbit. The control method controls pointing of payloads (communication antennas and instruments) and solar arrays disposed on the spacecraft. The method controls the spacecraft to provide Earth-pointing payloads and sun-pointing solar arrays so that the spacecraft body and solar arrays each rotate only around a single axis, a pitch axis of the spacecraft.

Abstract: A continuous three axes satellite pointing system for a solar powered satellite, the system comprising a body control module for producing three orthogonal axes body control torques, a momentum integration module, a momentum control module, and a program torque decoupling module. The program torque decoupling module de-couples the programmed motion parameters represented by the equation: IB&ohgr;′+B&ohgr;X(IB&ohgr;+BH).

Abstract: A method and apparatus for controlling a spacecraft having a first axis is disclosed. The apparatus comprises a momentum wheel unidirectionally rotatable about a spin axis substantially co-aligned with the first axis; a momentum wheel mount for tilting the momentum wheel about at least one of two axes substantially orthogonal to the spin axis; and a processor in communication with the momentum wheel and the momentum wheel mount. The processor commands either the momentum wheel speed or the momentum wheel tilting to control the attitude of the spacecraft in a first orientation and in a second orientation inverted from the first along the first axis, using the same control law.

Abstract: A method is disclosed for controlling the attitude of a satellite by controlling the speed of the gimbals of CMGs in a cluster of CMGs, the CMG having respective wheels mounted on gimbals that are mounted on a satellite platform to rotate about different orientation axes. From starting conditions and end conditions relating to attitude and angular speed and time, a cluster configuration is determined that is remote from any singular configuration such that exchanging angular momentum between the cluster of CMGs and the satellite during a given length of time will give rise to the desired attitude maneuver. The orientation of each gimbal is changed in simultaneous and independent manner into its reference orientation by using an angular position reference, applied in an open loop in the local servo-control of the angular positions of the gimbals.

Abstract: An apparatus for determining the pointing uncertainty of a satellite communications system comprises a communications terminal coupled to a pointing mechanism which is operative to move and position the communications mechanism. A data storage device is configured to store data from which an actual position of a preselected light source which radiates a reference light can be determined. An acquisition sensor is positioned to move with the communications terminal and is configured to focus incident light onto the acquisition sensor. A control mechanism is coupled to the pointing mechanism and is operative to move the pointing mechanism to a position which would focus the reference light source at a preselected position on the acquisition sensor if the pointing uncertainty were substantially zero. The reference light focusing at a measured position on the acquisition sensor.

Abstract: A system and method of estimating the attitude of a spacecraft is disclosed. A three-axis inertial-based estimate of spacecraft attitude is produced and is compared to a stellar-based estimate of spacecraft attitude. A Kalman filter having states associated with gyro scale factor and/or misalignment errors compares the stellar-based attitude estimate to the inertial-based estimate of attitude and apportions the total error into three time varying matrices. A first time varying matrix is associated with gyro scale factor and misalignment errors, a second time varying matrix is associated with gyro bias errors, and a third time varying matrix is associated with attitude errors. The time varying matrices are applied as corrective feedbacks to the inertial-based estimate of spacecraft attitude and are adaptively adjusted to minimize the error therein.