Abstract: A flight instrument and associated method where vertical speed information is combined with gyro information to produce gyro enhanced vertical speed information that is displayed to a pilot. The gyro information may include a pitch rate gyro and may include an azimuth rate gyro. The pitch rate gyro and azimuth rate gyro may be combined to yield a corrected pitch rate in turns. In one embodiment, the display is an airplane symbol that moves relative to a horizon line responsive to the gyro enhanced vertical speed information. The horizon line may also tilt in response to the azimuth rate gyro.

Abstract: Measurement of a spatial velocity vector in order to make determination of a geographical position possible. A spatial velocity meter includes an inertial measurement unit, a direction-sensing module and a velocity processor. The inertial measurement unit registers acceleration parameters and angular velocity parameters in three dimensions. The direction-sensing module registers a natural reference signal. The velocity processor receives the acceleration parameters, the angular velocity parameters and the natural reference signal, and based thereon, generates the spatial velocity vector.

Abstract: A process to control the trajectory of a spinning projectile. The projectile's accelerations are measured by means of an inertial unit in three axes within a reference linked to the projectile so as to prepare the piloting commands for the projectile's trajectory. The velocity vector V components (VX, VY, VZ) of the projectile are measured using a global positioning system (GPS) along three axes and within a land reference (GX, GY, GZ). Based on these measurements, the components (?X,?Y,?Z) of the acceleration vector ? are evaluated in the same land reference (GX, GY, GZ). At each of these measurements and periodical evaluations by the GPS, the Euler angles of the projectile are recalculated by combining the resultants of the velocity (V) and acceleration (?) measurements given by the GPS and those of the acceleration given by the inertial unit.

Abstract: An attitude sensing apparatus for determining the attitude of a mobile unit is provided that can reliably estimate an alignment angle between a GPS antenna coordinate system and an IMU coordinate system with good accuracy regardless of the magnitude of the alignment angle. Based on observation of the difference between a GPS angular velocity and an IMU angular velocity, an alignment angle estimating section estimates an alignment angle and sensor errors. An alignment angle adder and a sensor error adder cumulatively add and update the estimated alignment angle and sensor errors, respectively. The estimated alignment angle is fed back to an inertia data converter while the estimated sensor errors are fed back to an inertia data correcting section. The apparatus repeatedly performs estimation until the estimated alignment angle gradually approaches a true alignment angle by successively feeding back estimated values to a flow of alignment angle estimation process.

Abstract: A method for testing the performance of a Missile Time-Space-Position Information Unit. The method is utilized for Functional, Acceptance and Qualification test of the Missile Time-Space-Position Information Unit. The method has the capability of testing the Missile Time-Space-Position Information Unit to verify and document functionality of the Missile Time-Space-Position Information Unit prior to use of the Missile Time-Space-Position Information Unit in an operational environment.

Abstract: An accelerometer includes an inertial platform maintaining an attitude in response to a platform stabilizing controller signal and defining a spin axis and a reference plane. An accelerometer, coupled to the inertial platform a distance from the spin axis, defines a flex axis. The accelerometer generates an accelerometer signal in response to acceleration of the accelerometer. A second accelerometer defines a second flex axis, and is also coupled to the inertial platform a distance from the spin axis. The second accelerometer generates a second accelerometer signal in response to acceleration of the second accelerometer. A controller receives the first accelerometer signal and the second accelerometer signal and generates a linear acceleration signal in response to a sum of the first accelerometer signal and the second accelerometer signal and generates an angular acceleration signal from the difference.

Abstract: A dual capacitance accelerometer system includes two flexure plates coupled to the housing and defining a respective parallel flex axes. A fixed plate is adjacent to and in substantially parallel relation to the two flexure plates and is also coupled to the housing. The fixed plate and one of the flexure plates define a first distance and the fixed plate and the other flexure plate define a second distance. The first and second distances vary in response to acceleration forces acting upon the flexure plates.

Abstract: An accelerometer includes an inertial platform maintaining an attitude in response to a platform stabilizing controller signal and defining a spin axis and a reference plane. An accelerometer, coupled to the inertial platform a distance from the spin axis, defines a flex axis, which is perpendicular to the stability axis. The accelerometer generates an accelerometer signal in response to acceleration of the accelerometer. A second accelerometer defines a second flex axis also perpendicular to the stability axis, and is also coupled to the inertial platform a distance from the spin axis. The second accelerometer generates a second accelerometer signal in response to acceleration of the second accelerometer. A controller, including an angular acceleration signal generator, receives the first accelerometer signal and the second accelerometer signal and generates an angular acceleration signal from a difference in amplitudes between the first accelerometer signal and the second accelerometer signal.

Abstract: The present invention provides a digital signal processing method and system thereof for producing precision platform orientation measurements and local Earth's magnetic measurements by measuring threes axes gravity acceleration digital signals by an acceleration producer, detecting Earth's magnetic field vector measurement by an Earth's magnetic field detector to achieve digital three-axes Earth's magnetic field vector signals, and producing pitch, roll, and heading angles using said three-axes gravity acceleration digital signals and said digital three-axes Earth magnetic field vector signals by a Digital Signal Processor (DSP) chipset.

Abstract: A method for installation alignment of an inertial reference unit (IRU) with vehicle axes when the IRU is installed within the vehicle is provided. The vehicle axes include roll, pitch, and yaw axes. The method includes recording vehicle angular position data including roll and pitch, using an angular position measurement device with the vehicle being in a starting position, recording IRU data including roll and pitch, receiving measured nose plunge data, computing initial roll and pitch misalignment corrections, applying initial roll and pitch misalignment corrections to measured nose plunge data. A nose plunge yaw misalignment is determined using corrected nose plunge data and utilized to adjust the assumed heading reference.

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: A load relief system for a launch vehicle including a wind sensing system responsive to wind speed and direction at selected locations for providing an output of sensed wind speed and wind direction at the selected locations, a plant model responsive to the sensed wind speed and direction within a finite horizon, a current state of the launch vehicle, and control commands of the launch vehicle to predict the trajectory of the launch vehicle, an error circuit responsive to the predicted trajectory and a reference trajectory to produce a trajectory error, and an optimizer responsive to the trajectory error and configured to provide control commands to compensate for wind load over the finite horizon of the launch vehicle by reducing the total angle of attack.

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: An inertial reference system for a spacecraft includes an attitude control assembly and a plurality of force sensors. The attitude control assembly is coupled to a spacecraft body at a plurality of attachment points and at least one of the plurality of force sensors is integrated at each of the attachment points. In one embodiment, the force sensors are piezoelectric transducers. In another embodiment, the control assembly is coupled to the spacecraft body at four attachment points. In yet another embodiment, each of the force sensors is preloaded by a bolt that attaches the attitude control assembly to the spacecraft body.

Abstract: A weapon system comprising a mobile platform (1) e.g. an aircraft incorporating a first three-axis attitude reference sub-system (13), and a guidable vehicle (3) launchable from the platform (1) and incorporating a guidance sub-system (15) incorporating gyros (17) wherein correction of gyro parameters, i.e. scale factor and zero offset, of the vehicle attitude reference sub-system (15) is effected prior to launch of the vehicle (3) on the basis of repetitive comparisons of attitude data as measured by the platform and vehicle sub-systems (13, 15). The vehicle (3) may comprise a dispenser of a number of munitions (5) each itself incorporating a guidance sub-system (25) incorporating gyros (27), in which case corrections of gyro parameters of the missile guidance sub-system (25) may be effected prior to missile launch on the basis of repetitive comparisons of attitude data as measured by the dispenser and munition sub-systems (15, 25).

Abstract: A method and system for pointing and stabilizing a device that needs to be pointed and stabilized with a desired direction, are disclosed, wherein current attitude measurement and attitude rate measurement of the device measured by an attitude producer, which includes an inertial measurement unit, and the desired direction information measured by a target coordinates producer are processed by a pointing controller to compute rotation commands to an actuator. An actuator rotates and stabilizes the device at the desired direction according to the rotation commands in the presence of disturbances and parametric uncertainties to account for the undesired vibration due to disturbances. A visual and voice device provide an operator with visualization and voice indication of the pointing and stabilization procedure of the device.

Abstract: A projectile guidance system without gyros in which the projectile has an orthogonal body coordinate system. The projectile has a triax of accelerometers providing x, y and z acceleration data measured along the x, y and z axes respectively. A GPS antenna and receiver means provides onboard GPS position and velocity data in earth referenced navigational coordinates. A computer and program means stores and accesses time indexed GPS position and GPS velocity data and transforms x, y and z axis acceleration data from body to navigation coordinates. The program means is responsive to corresponding time indexed acceleration data and to GPS velocity and position data for calculating and outputing an estimated projectile roll, pitch and yaw angle via optimal smoothing techniques with respect to local level for each time index iteration of present position to a flight control system for guiding the projectile to a predetermined location.

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 payload launching method measures the attitude of a booster during a launch with an inertial measurement unit on the payload to compare the actual booster trajectory with a desired trajectory to control the payload at the deployment location, the inertial measurement unit on a booster carrying the payload using said desired trajectory to reach the deployment location

Abstract: A system which is packaged within a projectile fuze body and obtains data relative to the projectile during a launch. Sensors are provided which obtain in-bore data as well as in-flight data. The in-bore data is recorded at a fast rate during in-bore travel of the projectile and is read out, continuously, at a slower rate during in-flight travel. Both in-bore data and in-flight data are encoded and transmitted to a ground station for analysis.

Abstract: A sole means global navigation apparatus adapted for use on an aircraft includes a GPS receiver to provide GPS measurement data and an inertial sensor system adapted to provide inertial translational and rotational data which during time periods is independent of the GPS position data. A navigation system coupled to both of the GPS receiver(s) and the inertial sensor system determines a navigation solution as a function of both the condition of the GPS satellite data and the uncertainty in the Inertial data. An augmentation system coupled to the navigation solution determining system is used to increase the accuracy and/or integrity of the GPS/inertial sensor, thereby achieving style means navigation requirements.

Abstract: A projectile to be fired from a weapon barrel includes a projectile body and a plurality of stabilizing wings circumferentially distributed about a rearward portion of the projectile body. A pin pivotally attaches each stabilizing wing to the projectile body for pivotal motion about a pivot axis to assume a folded state and an outwardly pivoted, deployed state. In the folded state of the stabilizing wings the center of gravity of the respective stabilizing wings is at a greater radial distance from the longitudinal axis of the projectile body than the pivot axis. A gliding body is releasably mounted on each stabilizing wing. Each gliding body has a wing-shaped extension for contacting an inner surface of the barrel in the folded state of the stabilizing wings as the projectile travels in the barrel upon firing.