Abstract: A vibration control assembly for an aircraft includes a housing operatively coupled to the aircraft. Also included is a cage disposed within an interior region of the housing, the cage rotatable within the housing about a first axis. Further included is a gyroscope wheel disposed within the cage and rotatable about a second axis other than the first axis, wherein a controllable moment is imposed on the aircraft upon rotation of the gyroscope wheel to counter vibratory moments produced by the vehicle. Yet further included is a control assembly at least partially surrounding the gyroscope wheel for controlling the controllable moment. The control assembly includes a structure having an inner surface, a track disposed along the inner surface, and an arm operatively coupled to the gyroscope wheel, the arm having an end disposed within the track, the gyroscope wheel angularly displaceable upon translation of the arm along the track.

Abstract: A motorized weapon gyroscopic stabilizer which creates a stabilizing effect for single shot, semi-automatic, and fully automatic weapons. The rotating mass that generates the gyroscopic stabilizing effect can be the rotor of the motor. The motor is designed to allow the mass to rotate around the open core of the motorized weapon gyroscopic stabilizer. Because of its open core design the motorized weapon gyroscopic stabilizer allows the fired projectile to pass through it, or be mounted in line with the sighting mechanism allowing the target alignment-line of sight to pass through the motorized weapon gyroscopic stabilizer, or both.

Abstract: According to some aspects of the subject disclosure, a spacecraft comprises first and second pluralities of thrusters. The pluralities of thrusters are attached to a spacecraft body by booms configured to move the first plurality of thrusters between stowed and deployed positions. The deployed position of the first plurality of thrusters is farther north than is the stowed position of the first plurality of thrusters. The deployed position of the second plurality of thrusters is farther south than is the stowed position of the second plurality of thrusters. The first plurality of thrusters comprises a first thruster and a second thruster separated from each other in an east-west direction. The second plurality of thrusters comprises a third thruster and a fourth thruster separated from each other in the east-west direction.

Abstract: A control moment gyroscope (CMG) is provided, selectively having a first spatial configuration and a second spatial configuration at least during operation of the CMG. In the first spatial configuration the CMG occupies a smaller volume than in the second spatial configuration. For example, in the first spatial configuration no part of the CMG projects beyond a predetermined geometrical boundary, while in the second spatial configuration, a portion of the CMG projects beyond the geometrical boundary.

Abstract: In a movement direction control apparatus and a non-transitory computer readable medium, a reaction torque accompanying rotation of a rotor is utilized to control a rotation angle of a body in a yaw direction. A specification of a yaw angle of a wheel as a target of the movement direction is received, and information of a friction torque is acquired. A rotation angular acceleration in the yaw direction is calculated, based on the yaw angle, the specification of which has been received, and the reaction torque is calculated, based on the calculated rotation angular acceleration in the yaw direction. An operation instruction to a motor for yaw is generated, based on the calculated reaction torque and the acquired information of the friction torque.

Abstract: A method of de-saturating a control moment gyroscope that leverages a torque on an aircraft that is generated by airflow over the aircraft. As an aircraft navigates through an airspace, the aircraft may destabilize and reorient to form a sideslip angle that forms the airflow torque on the aircraft. The control moment gyroscope may be de-saturated into a neutral position that in turn exerts a torque on the aircraft that counters the airflow torque. A scissor pair of first and second control moment gyroscopes can be used for generating a torque in a single plane.

Abstract: A method of de-saturating a control moment gyroscope that leverages a torque on an aircraft that is generated by airflow over the aircraft. As an aircraft navigates through an airspace, the aircraft may destabilize and reorient to form a sideslip angle that forms the airflow torque on the aircraft. The control moment gyroscope may be de-saturated into a neutral position that in turn exerts a torque on the aircraft that counters the airflow torque. A scissor pair of first and second control moment gyroscopes can be used for generating a torque in a single plane.

Abstract: A solar powered air vehicle that can stay aloft for indefinite periods of time. The vehicle employs photovoltaic solar cells for primary power and high speed counter-rotating flywheels for energy storage and steering of the vehicle. The flywheels are placed in the wing to reduce airfoil drag. A control law provides three-axis stabilized control of the vehicle by controlling propeller pitch to vary the speeds of the flywheels.

Abstract: A vertical takeoff and landing aircraft according to the invention includes multiple thrust producing devices that produce thrusts applied in the substantially vertically upward direction; a target attitude setting portion that sets a target attitude used in attitude control of the aircraft; an inertia moment deriving portion that derives an inertia moment applied around a predetermined rotational axis of the aircraft; and a thrust adjustment portion that adjusts, during the attitude control of the aircraft, the thrusts to be produced by the respective thrust producing devices based on the target attitude set by the target attitude setting portion and the inertia moment during the attitude control, which is derived by the inertia moment deriving portion.

Abstract: A disturbance-cancellation system and method is provided that facilitates the adaptive cancellation of periodic disturbances. The disturbance-cancellation system and method includes a pair of control moment gyroscopes (CMGs) arranged together as a scissored pair. The scissored pair of CMGs is used to create a periodic torque with a controllable amplitude and frequency. The periodic torque created by the CMGs is adaptively controlled to cancel out periodic disturbances in the system. The system and method creates a periodic torque with a controllable amplitude and frequency by rotating the inner-gimbal assemblies of the CMGs in opposite directions at substantially equal phase and angular velocity. Amplitude matching can be achieved by adjusting the rotor spin rate. The steady-state constant motion results in a periodic torque along the output axis of the scissored pair, while components of torque along axes orthogonal to this one cancel out.

Abstract: In a GPS-supported inertial attitude and heading reference system equipped with a Kalman correction filter and a multiple axis fiber optic gyroscope, the invention provides for only that scale factor error which is determined for the measurement axis (e.g. the vertical axis) with relatively fast motion dynamics to be used as the Kalman filter correction value for the scale factor error correction for all the measurement axes of the FOG to determine and compensate for the scale factor error caused by changes in the wavelength of a common light source. The scale factor error correction is used only with a long time constant.

Abstract: A method and system for processing pulse signals within an inertial device is provided. The inertial device may have inertial sensors, such as accelerometers and gyroscopes. The inertial sensors may output signals representative of a moving body's motion. The signals may require correction due to imperfections and other errors of the inertial sensors. The inertial device may receive signals from the inertial sensors and process the signals on a signal-by-signal basis so that when processing the signals, the inertial device at least recognizes which sensor output a signal and when the signal was output. The inertial device may then correlate signals that were output from the inertial sensors at selected times in order to transform the signals into a desired navigational frame of reference.

Abstract: A system for directing an object is disclosed. The system includes a first gyroscope with a first rotating rotor and a second gyroscope with a second rotating rotor. Each gyroscope has a motor that rotates a first gimbal coupled to the rotor and a brake that decreases the rotation of a second gimbal coupled to the first gimbal. The first gyroscope and the second gyroscope generate a torque.

Abstract: Apparatus, methods, and systems are disclosed for removing stress on gyroscopes during acrobatic flight and for causing gyroscopes to fail in a controlled manner during testing and training. A valve for protecting a pneumatic-gyroscopic aircraft instrument comprises a body and a seat having a selective interrupter and an arm. The selective interrupter is positioned inside the body and can rotate inside the body. A method for protecting a pneumatic-gyroscopic aircraft instrument comprises allowing a pneumatic flow to the instrument during flight, and then, redirecting, selectively, without interrupting the pneumatic flow to the instrument. A system for protecting a pneumatic-gyroscopic aircraft instrument comprises a means for allowing a pneumatic flow to the instrument during flight, and a means for selectively redirecting, without interrupting, the pneumatic flow to the instrument without undue stress on the driver source.

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: In accordance with this invention, an apparatus and method for aircraft navigation are provided that utilize a blended architecture consisting of a global positioning system (GPS) and micro-electromechanical sensors (MEMS) for the primary navigation system and a laser gyroscope system for the secondary navigation system. The blended architecture of the present invention provides a navigation system that is at least as accurate, redundant and fault-tolerant as conventional navigation systems. In addition, the navigation system components may be distributed throughout the aircraft and may share computing resources with other avionics systems to process signals and provide data to the avionics systems. Overall, the navigation system of the present invention is significantly less expensive and easier to maintain, but equally or more precise and redundant, relative to conventional navigation systems.

Abstract: A calculation system for calculating the control signals which are designed to perform a frequency-dependent weighting of the angular velocity signals and the angular position signals. The angular velocity indicators are preferably used for stabilization in a high frequency band while the angular position indicators are suitable for stabilization in a low frequency band. A further calculation system which is designed to perform the frequency-dependent weighting such that the control signals for frequencies below a certain frequency are substantially determined by the angular position signals and the control signals for higher frequencies are substantially determined by the angular velocity signals.

Abstract: A system for directing an object is disclosed. The system includes a first gyroscope with a first rotating rotor and a second gyroscope with a second rotating rotor. Each gyroscope has a motor that rotates a first gimbal coupled to the rotor and a brake that decreases the rotation of a second gimbal coupled to the first gimbal. The first gyroscope and the second gyroscope generate a torque.

Abstract: A calculation system for calculating the control signals which are designed to perform a frequency-dependent weighting of the angular velocity signals and the angular position signals. The angular velocity indicators are preferably used for stabilization in a high frequency band while the angular position indicators are suitable for stabilization in a low frequency band. A further calculation system which is designed to perform the frequency-dependent weighting such that the control signals for frequencies below a certain frequency are substantially determined by the angular position signals and the control signals for higher frequencies are substantially determined by the angular velocity signals.

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: A gyroscope suitable for a remote-controlled helicopter is provided that can improve the control characteristics to a disturbance factor such as wind as well as the response characteristics under yaw-axis control in a correct pitch operation. The angular velocity detection signal from the yaw-axis angular velocity sensor 10 is added to the yaw-axis control signal and the angular velocity reference value from the controlling gear. The PID controller 2 receives the resultant addition signal. The control signal from the PID controller 2 is added to the correct pitch control signal and the yaw-axis control signal sent via the correct pitch to yaw-axis control mixing unit 3. The actuator 6 receives the resultant addition signal and controls the yaw-axis driving unit 7. The yaw-axis angular velocity detection sensor 10 detects the angular velocity around the yaw-axis of the airframe 9.

Abstract: A missile guidance system employs a strap down inertial measurement unit with a separate accelerometer and star sensor that are moved for preflight accelerometer initialization. During flight, light is directed on the accelerometer and inertial measurement unit to produce on the star sensor images of the light source, the accelerometer and the inertial measurement unit. The location of these images, which manifest the acceleration and position of the inertial measurement unit relative to the vehicle, are used to improve the accuracy of the inertial measurement unit data for vehicle guidance.

Abstract: In a helicopter having a device for stabilizing yaw motion through the use of a tail rotor with a plurality of rotor blades extending radially from a hollow rotor shaft which is mounted for rotation about a transverse rotor axis. The tail rotor is provided with a device for varying the thrust by a pilot input to the blades of the tail rotor and through the use of a gyro rotor for automatic control of the tail rotor. The invention further comprises a three point mixing linkage that permits the control of the tail rotor through both the pilot inputs and the gyro rotor inputs.

Abstract: A dual flywheel assembly for use in an airborne vehicle for storing mechanical energy therein prior to flight and for permitting the extraction of electrical energy therefrom during flight, which assembly includes two flywheels which are linked by a suitable linkage structure such that, if roll motion of the vehicle starts to occur during flight, the flywheels tilt in equal and opposite directions out of their normal planes of rotation, which tilting motions act in a passive manner to stabilize the roll motion of the vehicle.

Abstract: An in-field method for correcting the thermal bias error calibration of the gyros of a strapdown inertial navigation system. The method is begun after initial alignment while the aircraft remains parked with the inertial navigation system switched to navigation mode. Measurements are made of navigation system outputs and of gyro temperatures during this data collection period. A Kalman filter processes the navigation system outputs during this time to generate estimates of gyro bias error that are associated with the corresponding gyro temperature measurements. Heading error correcting is performed after the extended alignment data collection period as the aircraft taxis prior to takeoff. The gyro bias error-versus-temperature data acquired, along with the heading error corrections, are employed to recalibrate the existing thermal model of gyro bias error by means of an interpolation process that employs variance estimates as weighting factors.

Abstract: A method and apparatus are provided for addressing the effect of centripetal acceleration upon estimates of cross-track velocity, for determination of east gyro bias error, generated with a taxiing aircraft. After initial estimates of crab angle, ratio of crab angle to centripetal acceleration and lever arm are provided, velocity, heading angle and heading angle rate are observed as the aircraft taxis. An estimated value of centripetal acceleration is taken as the product of heading angle rate and heading velocity. Cross-track velocity is computed from cross-heading velocity and this is integrated to generate cross-track position. A Kalman filter generates various gains, including one associated with the ratio of crab angle to centripetal acceleration, for error allocation.

Abstract: In a helicopter having a tail rotor with a plurality of rotor blades extending radially from a hollow rotor shaft which is mounted for rotation about a transverse rotor axis, and having a push-pull rod extending through the hollow shaft and operably connected to the blades to manually vary the collective pitch of the blades, a device for automatically stabilizing the yaw motion of the helicopter includes a gyroscopic assembly having a gyro rotor mounted to rotate with the tail rotor, to pivot about a substantially longitudinal pivot axis by and at the outboard end of the push-pull rod and to automatically vary the collective pitch of the blades in response to yaw motion.

Abstract: An aircraft banking angle indicator which includes a banking angle sensing arrangement for generating an actual banking angle signal, a maximum banking angle input device for manually entering a selected maximum banking angle, a signal processing arrangement having respective inputs connected to the banking angle sensing arrangement and to the maximum banking angle input devices for generating an excessive banking output signal indicative of a banking angle exceeding the maximum banking angle, and an indicating arrangement having a signal input connected to the banking angle output signal for indicating the excessive banking angle.

Abstract: A terrain surveillance, powered aircraft is disclosed to have an on-board video camera that is not gimballed relative to the aircraft, the aircraft typically comprising multiple components; the method of the invention including: assembling the components into aircraft configuration at or near a launch site; launching the aircraft at that site with no human pilot on board same; remotely controlling the flight path of the aircraft by radio or cable link, to fly to a location for terrain surveillance; operating the on-board video camera to survey the terrain while varying yaw of the aircraft to provide sweep viewing of the terrain; and recovering the aircraft by controlling its flight to steeply descend to a landing zone.

Abstract: Apparatus for stabilization and control of an aircraft includes a pair of strapdown devices that are rigidly fixed in space relative to one another. Each device includes two two-axis gyroscopes, permitting the generation of triple redundancy. The measurement axes (M.sub.11, M.sub.12 and M.sub.21, M.sub.22) have different arrangements in the horizontal plane. The redundant axis of each strapdown device is arranged in such a way (preferably at 45 degrees) that an allocation to the two horizontal measurement axes is provided by computational transformation.

Abstract: Vacuum systems for use aboard aircraft and other vehicles, wherein, by simplified design, the systems of this new invention are virtually foolproof. This new invention utilizes the age old venturi due to its simplicity and lack of moving parts, but due to innovation, the limitations of the venturi have been overcome. This new invention installs the venturi within the exhaust tailpipe of engines and puts the extreme high velocity of exhaust gas movement to good use. These systems are compatible with just about any engine, and obviously venturi ice problems are eliminated. These systems may be used to provide redundant power for gyro systems and windshield wiper systems. By the addition of a suction relief valve and connective tubing the system will provide regulated suction pressure for the operation of a Gyro Horizon and or a Directional Gyro. By further addition of a restrictor needle valve and additional connective tubing, one can power either a Turn and Bank indicator or A Turn Co-ordinator.

Abstract: A set of two two-degrees-of-freedom rate gyroscopes and three linear accelerometers are assembled in a single module adapted to be mounted within a single aircraft electronics control unit, the unit comprising a strapped-down attitude and heading reference system. The module base provides a common keyed support for the two pre-calibrated gyros and three accelerometers in intimate mechanical and thermal association. The two gyros are oriented in the base and the base oriented in the aircraft so that the spin axis of one gyro is oriented parallel to the aircraft Z axis and that of the other gyro parallel to the aircraft Y axis while the gyro pickoffs and torquers (input and output axes, respectively) are rotated or skewed forty-five degrees about the spin axes to positions such that the input and output axes lie along the slant heights of a forty-five degree half angle right circular cone, the axis of which lies along the aircraft X axis.

Abstract: A circuit for applying rate feedback to the pitch and yaw canard commands a guided projectile or missile to stabilize the airframe. The circuit demodulates and differentiates the signal from the gyro cage coil to obtain a signal representative of the airframe movement at its natural frequency. Phase detectors provide a phase reference by comparing the cage signal carrier with the output of hall effect elements in the speed sensing circuitry. The rate amplitude from the differentiator is then multiplied by the phase references to resolve the rate into pitch and yaw command components. Summation circuits then scale and subtract the rate commands from the guidance commands to provide negative rate feedback.

Abstract: A four-axis gyro stabilized inertial reference system. The system comprises a nested gimbal configuration, which includes a platform and three successively larger and concentrically disposed inner, middle and outer gimbals and an outer casing. The platform and inner, middle and outer gimbals are each configured to permit rotation about an associated axis with that axis being perpendicular to the corresponding axis of the neighboring gimbal or platform. The system further includes a platform stabilizing means for maintaining the platform in a substantially fixed spatial orientation, independent of the spatial orientation of the outer casing and over a predetermined range of angular rates.

Abstract: The control system includes a sensing mechanism by which a true pure signal from flapping bending activity resulting from an external load force on a rotor blade in a rotor system is generated and fed back to a gyroscope which then precesses to return a correcting feathering motion to the rotor blade through a swashplate. Such generation is developed by the sensing mechanism which comprises a cantilevered beam or spring system secured at its one end to a fixed hub arm about which the blade feathers and having its other end operatively connected to the gyroscope. The gyroscope is positioned outside of a high force blade feathering loop and is independently sized from the rotor to which it is coupled. The system also provides for pilot command input to produce a control feathering motion to a pair of blades independently of the sensing means.

Abstract: Apparatus for stabilizing a platform characterized by a frame carrying a stabilizing gyroscope that includes a substantially horizontal rotor with its shaft fixed with respect to the frame so as to effect a precessing force 90.degree. in the direction of rotation responsive to a torque force on the shaft in any azimuthal direction in a horizontal plane; a shifting carriage for shifting a load carrying compartment to move its center of gravity with respect to the frame to a plurality of successive second locations to offset to a desired degree a torque force and the reactive precessing force; and a control device for detecting a slight change in attitude from a desired attitude and effecting via the shifting carriage a controlled shifting of the load carrying compartment to a new second location to reestablish the desired attitude. The invention is readily applicable to vertical take-off and landing craft or other platforms having a lifting device connected therewith.