Abstract: A method and apparatus for estimating gyro scale factor during normal spacecraft operations, using small attitude motions that are compliant with mission pointing accuracy and stability requirements and a signal processing method that specifically detects the intentionally induced motions. This process increases operational availability by avoiding the need to take the spacecraft offline for large calibration maneuvers.

Abstract: The method for unloading the inertia wheels of a spacecraft comprising three references axes X, Y, Z, the axis Z corresponding to a pointing direction, consists in inverting the direction of accumulation of the angular momentum in the wheels by automatic rotational flipping of the spacecraft about the axis Z, the pointing direction remaining fixed. The method has application to the field of satellites or of interplanetary probes.

Abstract: A system and method for controlling and stabilizing a satellite or other vehicle about any axis is disclosed. Embodiments achieve this three-axis control and stabilization with a spherical motor system or reaction sphere capable of storing momentum in a rotor. The spherical motor system comprises a spherical rotor having permanent magnets arranged in evenly-spaced antipodal pairs. Each of the permanent magnets are oriented with the same magnetic pole facing outward from a center of the rotor. The spherical motor system also comprises a stator which has magnetic sensors surrounding electromagnets. The spherical motor system further comprises a control system that controls timing and duration of energization of the electromagnets based on the detection of one of the permanent magnets by the magnetic sensors. Such a spherical motor system provides three-axis stabilization and control of a satellite (or other vehicle).

Abstract: The method for unloading the inertia wheels of a spacecraft comprising three references axes X, Y, Z, the axis Z corresponding to a pointing direction, consists in inverting the direction of accumulation of the angular momentum in the wheels by automatic rotational flipping of the spacecraft about the axis Z, the pointing direction remaining fixed. The method has application to the field of satellites or of interplanetary probes.

Abstract: Spacecraft momentum management techniques are coordinated with station-keeping maneuvers or other delta-V maneuvers. A body stabilized spacecraft attitude is controlled, the spacecraft including at least one momentum/reaction wheel, and a set of thrusters. A first momentum storage deadband limit is adjusted, the adjustment being related to a first delta-V maneuver window. A momentum management strategy is executed with the adjusted first momentum storage deadband limit such that a first thruster firing that performs desaturation of the momentum/reaction wheel also provides velocity change beneficial to the first delta-V maneuver.

Abstract: The present invention relates to a method for maneuvering an imaging satellite, and more particularly a method for commanding control moment gyroscopes on an imaging satellite to change the attitude of the satellite. In one embodiment, a method for generating a gimbal rate trajectory for maneuvering a satellite to point to a target includes providing a final attitude for pointing to the target and a final satellite rate for imaging the target, providing an initial attitude and an initial set of gimbal angles, and determining a gimbal rate trajectory from the initial set of gimbal angles to a final set of gimbal angles.

Abstract: A system and method for deploying a satellite having an integrated bus and an aperture, wherein the integrated bus extends along its longitudinal axis and wherein the aperture intersects the longitudinal axis. The satellite is deployed into an orbit, wherein deploying includes orienting the satellite so that the aperture points at a desired location. The satellite is spun so that the logitudinal axis and the aperture remain pointing at the desired location and data regarding the desired location is captured via the aperture.

Abstract: A system for damping nutation and removing wobble of a spacecraft spinning about a given axis is provided. The system includes a sensor configured to determine three dimensional attitude measurements of the spacecraft, a processor operatively coupled to the sensor and configured to execute a process that facilitates aligning the spin axis with a spacecraft momentum vector. The processor, when executing the process, is programmed to receive spacecraft attitude data from the sensor, determine a torque command using the received attitude data, and control a momentum storage actuator on the spacecraft using the determined torque command such that an angular deviation about the given axis is reduced.

Abstract: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: The invention relates to a vertical take-off and landing gyropendular craft or drone device (FIG.

Abstract: A method for controlling an attitude control system (30) for a space vehicle (10), the attitude of the space vehicle being controlled during at least one preparation phase followed by an observation phase during which an image capture is performed. In an attitude control system that includes a maneuvering subsystem (300) which includes at least one reaction wheel, the method includes, during the at least one preparation phase, a preparation step (50), during which commands are issued to the maneuvering subsystem in order to control the attitude of the space vehicle, followed by a step (55) of stopping the at least one reaction wheel prior to the observation phase. In addition, during the observation phase, a fine control subsystem (310) having a lower vibration signature than that of the maneuvering subsystem, is sent commands in order to control the space vehicle's attitude. An attitude control system for a space vehicle is also described.

Abstract: An exoatmospheric device comprises an inertial measurement unit comprising a spin axis gyroscope adapted to obtain data regarding angular rotation of the exoatmospheric device about a spin axis, and at least one off-spin axis gyroscope adapted to obtain data regarding angular oscillation of the spin axis. The exoatmospheric device further comprises a processing unit coupled to the inertial measurement unit and adapted to detect attitude error by analyzing correlations between the measured rotation about the spin axis and the measured angular oscillation of the spin axis.

Abstract: The invention is for spacecraft reaction wheel with an axial flux dual Halbach rotor with an air coil stator. This wheel is more efficient and has fewer disturbances than a conventional reaction wheel with a brushless DC motor.

Abstract: Provided is a method and system for maneuvering a first spacecraft relative to a nearby second spacecraft that occupies a finite volume of space including obstacles in the vicinity of the second from a measured present relative position to a desired ending relative position. A trajectory is computed for the first spacecraft from the present relative position to the ending relative position using the A* algorithm. One or more thrusters of the first spacecraft are then fired. The first spacecraft's relative position and velocity are then measured and compared to the computed trajectory.

Abstract: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: Provided are an attitude control system and method of a spacecraft of an artificial satellite that may enhance a maneuverability and a controllability by simultaneously applying a reaction wheel and a thruster among drive units used to maneuver an attitude of the spacecraft of the artificial satellite. The attitude control system may include: a thruster-based attitude controller which control firing time of thrusters mounted on the spacecraft; and a reaction wheel-based attitude controller controlling driving of a reaction wheel mounted on the spacecraft. The spacecraft may include a plurality of reaction wheels. When a defect occurs in the spacecraft due to a partial malfunction of the reaction wheels, an attitude maneuverability of the spacecraft may be corrected by simultaneously applying the thruster-based attitude controller and the reaction wheel-based attitude controller.

Abstract: Methods and systems are provided for controlling attitude of a vehicle using a reaction wheel onboard the vehicle. One exemplary method involves receiving a torque command for adjusting the attitude of the vehicle using the reaction wheel, determining a phase error of the reaction wheel based at least in part on the torque command, and determining a motor torque command for the reaction wheel based on the phase error. The motor torque command is provided to an electric motor of the reaction wheel to apply a corresponding torque to the rotor of the reaction wheel. The relationship between the magnitude of the motor torque command and the magnitude of the phase error is nonlinear. In exemplary embodiments, the magnitude of the motor torque command exceeds the stiction torque, at least instantaneously, when the reaction wheel has fallen behind an expected position by more than a threshold amount.

Abstract: A system for controlling an aerospace vehicle using on-line inertia estimation may include an attitude sensor to measure an attitude of the aerospace vehicle. The system may also include a processor on board the aerospace vehicle. An inertia estimator operable on the processor may generate an on-line inertia estimate of the aerospace vehicle. A rate and attitude estimator operable on the processor may determine an angular position and angular velocity of the aerospace vehicle using the attitude measurement of the aerospace vehicle and the on-line inertia estimate for controlling movement and orientation of the aerospace vehicle without any rates of rotation of the aerospace vehicle being required.

Abstract: A system and method are disclosed for an innovative spacecraft (S/C) safing methodology. The system and method use wheel momentum to perform a rotisserie rotation so that the spacecraft can slew at a fast rate in a controlled fashion in order to keep the spacecraft power safe and maintain telemetry and command (T&C) coverage. The system and method involve a reaction wheel system and a spacecraft control processor (SCP). The SCP is in communication with the reaction wheel system. The SCP is operable to determine a rotisserie axis, and take the cross product of the rotisserie axis and a momentum vector to result in a unit vector. The SCP is further operable to slowly slew the spacecraft about the unit vector through an angle until the rotisserie axis aligns with the momentum vector, and rotate the spacecraft about the rotisserie axis to keep the spacecraft power safe.

Abstract: An actuator device for varying the attitude of a spacecraft includes a reversible conversion chain of electrical energy to mechanical rotation energy of a flywheel. The electrical energy is stored in a capacitive element that may be a supercapacitor. The actuator device also includes an electrical power converter, connected, on one hand, to the capacitive element and intended to be connected, on the other hand, to the spacecraft power bus. The converter makes it possible to compensate for losses to keep a total energy of the actuator device constant.

Abstract: Various embodiments of the present invention include assemblies and methods for minimizing the amplitude of attitude jitter. In one embodiment, a split flywheel assembly includes a plurality of independent concentric flywheels axially aligned and in operable engagement with one another such that each flywheel is configured to be independently controlled in order to manipulate the phase difference therebetween.

Abstract: Various embodiments of the present invention include an attitude control system for use with small satellites. According to various embodiments, the system allows rapid retargeting (e.g., high slew rates) and full three-axis attitude control of small satellites using a compact actuation system. In certain embodiments, the compact actuation system includes a plurality of single-gimbaled control moment gyroscopes (SGCMG) arranged in a pyramidal configuration that are disposed within a small satellite.

Abstract: A method includes providing two or more momentum wheels arranged for rotation on a spacecraft in a momentum-canceling set. The method further includes causing the two or more momentum wheels in the momentum-canceling set to rotate at momentum-canceling speeds. Additionally, the method includes reducing rotational speed of a momentum wheel in the momentum-canceling set to initiate a slew of the spacecraft.

Abstract: Spacecraft vehicles and methods for providing reorbiters capable of moving target objects are provided. More particularly, a reorbiter utilizing electrostatic or Coulomb force for acting on target objects and for moving the target objects into new orbits or altitudes are provided. The reorbiter may establish an electrostatic force by controlling the electrical potential of the reorbiter through active charge emission. The target object may acquire an electrical potential due to its interaction with the space plasma and photoelectrons or the reorbiter may impart additional charge to the target object. After establishing the electrostatic force, a propulsion system of the reorbiter is operated to provide thrust. Thrust is directed to change the orbit and/or position of the target object that is being moved by the electrostatic force between the reorbiter and the target object.

Abstract: Spacecraft control modules can steer a spacecraft during a slew maneuver. The spacecraft control modules can receive a target attitude for the spacecraft slew maneuver. An angle between a current attitude and the target attitude may be determined. A smooth attitude offset profile with respect to a steering frame may be established from the current attitude to the target attitude. A smooth attitude command may be generated from the smooth attitude offset profile. The smooth attitude command may be applied to an attitude control module of the spacecraft.

Abstract: Methods and systems are provided for orienting an agile vehicle using a control moment gyroscope array. A method comprises obtaining initial vehicle parameters for the vehicle and obtaining target vehicle parameters for the vehicle. The method further comprises determining command parameters based on a difference between the target vehicle parameters and the initial vehicle parameters, and simulating operation of the control moment gyroscope array using the command parameters and a torque value being at least equal to a maximum achievable torque for the control moment gyroscope array. When the simulated vehicle parameters are substantially equal to the target vehicle parameters, the method further comprises determining a torque profile for the control moment gyroscope array based on the simulated operation and operating the control moment gyroscope array using the torque profile.

Abstract: The effects of IMU gyro and accelerometer bias errors are significantly reduced in accordance with the present teachings by a system or method for commanding an IMU or vehicle through a series of preprogrammed maneuvers. The maneuvers can be designed to minimize the effects of other gyro errors including scale factor errors, nonlinearities, cross coupling/misalignment, and scale factor asymmetries. A sample maneuver is provided which demonstrates performance based on a sequence of roll and yaw maneuvers resulting in zero build up of error at the end of a maneuver cycle period as a result of these errors. Modification of the system involves the addition of control logic to determine the maneuver period, maneuver rate, and vehicle orientation. No additional hardware beyond possible fuel required to perform the maneuver is required.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies results in minimal HCT gimbal stepping.

Abstract: A method of controlling inertial attitude of an artificial satellite in order to perform a navigation function and to maximize terrestrial coverage of the Earth by the satellite. The method includes deploying the artificial satellite in an orbit about the poles of the Earth; applying gyroscopic precession to the artificial satellite spin axis to precess and maintain the satellite near the ecliptic pole; deploying the artificial satellite so that the spin axis is initially perpendicular to or substantially perpendicular to sun lines; and applying gyroscopic precession to the artificial satellite spin axis to precess the spin axis away from an initial deployed attitude at a selectively-variable precession rate and to maintain the spin axis perpendicular to or substantially perpendicular to a the sun lines.

Abstract: Disclosed herein are two separate processes that do not require a propellant and do not produce an equal and opposite reaction against any external form of matter in the Local Inertial Reference Frame and do not violate Newton's Laws in the Universal Reference Frame. The first process produces horizontal motion, relies on the earth's gravitational field as an external force, and has been successfully tested. The second process produces vertical motion and relies only on the aether. It has been successfully tested considering the effect of the earth's gravity. Due to the law of conservation of angular momentum, the first process is considered by some to not be possible, but with the proper use of an external field (for example, gravity) and the phenomenon of precession, it is clearly possible. A clear distinction is made between a simple rotor and a gyroscope which is a far more complex device.

Abstract: A method of controlling attitude of a spacecraft during a transfer orbit operation is provided. The method includes providing a slow spin rate, determining the attitude of the spacecraft using a unified sensor set, and controlling the attitude of the spacecraft using a unified control law. The use of a unified set of sensors and a unified control law reduces spacecraft complexity, cost, and weight.

Abstract: The gyrodyne includes an inertial wheel mounted, via a wheel support, on the moving part or rotor of a cardan assembly. The cardan assembly is provided with a stator and the rotor is rotatable with respect to the stator about a first axis of rotation, it being possible for the spinner of the inertial wheel to be set in rotation about a second axis of rotation not aligned with the first axis of rotation. The stator of the cardan is mounted on a block and fixed to this block via an arrangement of vibration attenuators or insulators. The mechanism for setting the rotor in rotation is at least partially housed in the interior volume of the block.

Abstract: Disclosed is an attitude change control system that is designed to efficiently output a torque for attitude change of a space craft using CMGs and realizes a real time CMG driving rule. A CMG gimbal steering law 15 generates target profiles for setting angle and angular velocity for each gimbal by applying an anisotropic weighted gradient method based upon the necessary torque calculated by a feed back controller 13 and a feed forward controller 14 from the angle and angular velocity in the target direction from the attitude navigator 12 and the current angle and angular velocity of the space craft estimated by the attitude estimator 11 as well as the current condition of each gimbal from the CMG 40, thereby controlling the CMG 40 for changing the attitude of the space craft dynamics 50 to the target direction.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies results in minimal HCT gimbal stepping.

Abstract: A control moment gyroscope system for delivering a target torque to a spacecraft including a rotor assembly having a rotor and a motor to spin the rotor about a rotor axis. A gimbal assembly has a gimbal for supporting the rotor assembly and a gimbal torque motor to rotate the gimbal about a gimbal axis, which is normal to the rotor axis, to generate an output torque. A control system has a sensor for determining the output torque and a processor in communication with the rotor assembly, the gimbal assembly and the sensor. The processor requests the target torque and establishes a feedback control loop to generate a torque error signal based on the output torque for bringing the output torque within a predetermined range of the target torque.

Abstract: Methods and systems are provided for reorienting an agile vehicle, such as a satellite or spacecraft, using a control moment gyroscope (CMG) array. The CMG array comprises a plurality of CMGs onboard the agile vehicle. A method comprises obtaining an input torque command for reorienting the vehicle using the CMG array and, when the angular momentum of the CMG array violates or is approaching a momentum boundary criterion, decreasing the input torque command in the kinetic momentum direction, resulting in a modified torque command, and operating the CMG array using the modified torque command.

Abstract: This invention uses existing technology to provide a means to initiate and control the rotation of a space station of the type proposed by Werhner von Braun in the 1950s, without the need for an external torque. This is accomplished by creating an angular momentum vector within the hub of the station which is precisely equal and opposite to that of the rest of the station.

Abstract: A system for providing attitude control with respect to a spacecraft is provided. The system includes a reaction wheel control module configured to control a number of reaction wheel assemblies associated with the spacecraft in order to control attitude, and a maneuver control module configured to use a number of gimbaled Hall Current thrusters (HCTs) to control the total momentum associated with the spacecraft during an orbit transfer. The total momentum includes the momentum associated with the reaction wheel assemblies and the angular momentum of the spacecraft. Using the gimbaled HCTs to control the momentum associated with the reaction wheel assemblies during the orbit transfer results in minimal HCT gimbal stepping.

Abstract: A system and method are disclosed for an innovative spacecraft (S/C) safing methodology. The system and method use wheel momentum to perform a rotisserie rotation so that the spacecraft can slew at a fast rate in a controlled fashion in order to keep the spacecraft power safe and maintain telemetry and command (T&C) coverage. The system and method involve a reaction wheel system and a spacecraft control processor (SCP). The SCP is in communication with the reaction wheel system. The SCP is operable to determine a rotisserie axis, and take the cross product of the rotisserie axis and a momentum vector to result in a unit vector. The SCP is further operable to slowly slew the spacecraft about the unit vector through an angle until the rotisserie axis aligns with the momentum vector, and rotate the spacecraft about the rotisserie axis to keep the spacecraft power safe.

Abstract: Techniques for providing singularity escape and avoidance are disclosed. In one embodiment, a method for providing control moment gyroscope (CMG) attitude control singularity escape includes calculating a Jacobian A of a set of control equations, calculating a measure of closeness to a singularity, and comparing the calculated closeness to a threshold value, when the calculated closeness is less than or equal to the threshold value, recalculating the Jacobian A. Recalculating may include determining a new direction of virtual misalignment of ? and ?, recalculating the Jacobian inputting the new direction of the virtual misalignment, recalculating the measure of closeness to a singularity, and comparing the measure of closeness to the threshold value. Further, the method may include calculating a gimbal rate command if the of closeness is greater than the threshold value and generating a torque from the gimbal rate command to control the attitude of a satellite.

Abstract: Methods and systems are provided for reorienting an agile vehicle, such as a satellite or spacecraft, using a control moment gyroscope (CMG) array. The CMG array comprises a plurality of CMGs onboard the agile vehicle. A method comprises obtaining an input torque command for reorienting the vehicle using the CMG array and, when the angular momentum of the CMG array violates or is approaching a momentum boundary criterion, decreasing the input torque command in the kinetic momentum direction, resulting in a modified torque command, and operating the CMG array using the modified torque command.

Abstract: Techniques for providing singularity escape and avoidance are disclosed. In one embodiment, a method for providing control moment gyroscope (CMG) attitude control singularity escape includes calculating a Jacobian A of a set of control equations, calculating a measure of closeness to a singularity, and comparing the calculated closeness to a threshold value, when the calculated closeness is less than or equal to the threshold value, recalculating the Jacobian A. Recalculating may include determining a new direction of virtual misalignment of ? and ?, recalculating the Jacobian inputting the new direction of the virtual misalignment, recalculating the measure of closeness to a singularity, and comparing the measure of closeness to the threshold value. Further, the method may include calculating a gimbal rate command if the of closeness is greater than the threshold value and generating a torque from the gimbal rate command to control the attitude of a satellite.

Abstract: A modular control moment gyroscope (CMG) system for a spacecraft attitude control system (ACS) is formed by a plurality of CMG modules, wherein each CMG module has a modular enclosure design that is identical to that of the other CMG modules, such that the plurality of CMG modules are mountable in a spacecraft array bus structure in any desired one of multiple array configurations.

Abstract: An actuator device for varying the attitude of a spacecraft includes a reversible conversion chain of electrical energy to mechanical rotation energy of a flywheel. The electrical energy is stored in a capacitive element that may be a supercapacitor. The actuator device also includes an electrical power converter, connected, on one hand, to the capacitive element and intended to be connected, on the other hand, to the spacecraft power bus. The converter makes it possible to compensate for losses to keep a total energy of the actuator device constant.

Abstract: A self-contained momentum control system (MCS) for a spacecraft is provided for small satellites. The MCS features a miniaturized gyroscopic rotor with a rotational speed in excess of 20,000 RPM. The MCS includes at least three control moment gyroscopic mechanical assemblies (CMAs) rigidly mounted within a single enclosure, where each CMA mounted in an orientation whereby the longitudinal axis of each CMA is either orthogonal to every other CMA or is parallel to another CMA but in the opposite orientation. In order to further reduce the size of the MCS, an electronics package that is configured to interface command and control signals with and to provide power to the CMAs is included within the MCS enclosure. A plurality of shock isolation devices are used to secure each of the CMAs to the enclosure in order to reduce the launch load upon the CMAs thereby allowing the use of smaller rotor spin bearings. The MCS enclosure surrounding the CMAs and support structure is hermetically sealed.

Abstract: An attitude control actuator of a spacecraft includes a first momentum wheel driven in rotation by a motor and a second momentum wheel. Both momentum wheels are movable in rotation about an axis and are mechanically coupled by coupling means imposing between a rotation speed ?1 of the first momentum wheel and a rotation speed ?2 of the second momentum wheel a ratio R=?2/?1, negative on the one hand and continuously modifiable by the coupling means in response to a command on the other hand so as to modify the total angular momentum of the actuator while the total kinetic energy of the wheels are kept constant. In one embodiment, both momentum wheels have the same inertia and are coupled through a toroidal variator, where ratio R varies between ?3 and ?1/3. The motor drives the momentum wheels and operates as a generator for braking the momentum wheels.

Abstract: A control system for adjusting the attitude of a spacecraft comprises a set of control moment gyroscopes (CMGs) configured to allow null space maneuvering. The control system further comprises a momentum actuator control processor coupled to the set of CMGs and configured to determine a mandatory null space maneuver to avoid singularities and determine an optional null space maneuver to increase available torque. The mandatory null space maneuver can be calculated based upon certain gimbal angles, and can be implemented by augmenting the inverse-Jacobian control matrix.

Abstract: The invention relates to a drive system comprising a control unit (1), which is connected to drive units (3a, 3b, 3c, 3d, 3e, 3f, 3g) via a data bus (2) for the exchange of data. According to the invention, one drive unit (3a, 3b) is connected to the drive motor (7a, 7b) in order to control the latter (7a, 7b) and an additional drive unit (3c, 3d, 3e, 3f, 3g) is connected to a magnetic bearing (11c, 11d, 11e, 11f, 11g) of a magnetic spindle bearing arrangement (23) in order to control said bearing (11c, 11d, 11e, 11f, 11g). The invention thus provides a drive system, in which a magnetic spindle bearing arrangement (23) is integrated.

Abstract: A system and a method for commanding a spacecraft to perform a three-axis maneuver purely based on “position” (i.e., attitude) measurements. Using an “inertial gimbal concept”, a set of formulae are derived that can map a set of “inertial” motion to the spacecraft body frame based on position information so that the spacecraft can perform/follow according to the desired inertial position maneuvers commands. Also, the system and method disclosed herein employ an intrusion steering law to protect the spacecraft from acquisition failure when a long sensor intrusion occurs.

Abstract: Provided are an attitude control system and method of a spacecraft of an artificial satellite that may enhance a maneuverability and a controllability by simultaneously applying a reaction wheel and a thruster among drive units used to maneuver an attitude of the spacecraft of the artificial satellite. The attitude control system may include: a thruster-based attitude controller which control firing time of thrusters mounted on the spacecraft; and a reaction wheel-based attitude controller controlling driving of a reaction wheel mounted on the spacecraft. The spacecraft may include a plurality of reaction wheels. When a defect occurs in the spacecraft due to a partial malfunction of the reaction wheels, an attitude maneuverability of the spacecraft may be corrected by simultaneously applying the thruster-based attitude controller and the reaction wheel-based attitude controller.