Litton Systems, Inc.

Northrop Grumman Guidance and Electronics Company, Inc.

Northrop Grumman Systems Corporation

Northrop Grumman Corporation

Robert John Buchler

Gang Kevin Liu

Daniel A Tazartes

Allan J Brockstein

John G Mark

David Y Hsu

Joseph A Ratkovic

Peter Kyriacou

Charles P Bruner

Shaw-Wen Fann

Larry R Moeller

Kai-Nan Chueh

Kevin Liu

A method for estimating a probability of failure of a least-squares ambiguity decorrelation adjustment (LAMBDA) method is provided. The LAMBDA method is used for estimation of double difference carrier phase integer ambiguity. A plurality of condition sets are selected. Each condition set comprises a probability of failure (P) for a boot-strap method of estimation of the double difference carrier phase integer ambiguity, a number of space vehicles (N), and a ratio test tolerance for the LAMBDA method. A plurality of Monte Carlo simulations are run on the plurality of condition sets to obtain a plurality of result sets. Each result set comprises a probability of lambda fail (P) and a probability of lambda reject (P) for one condition set of the plurality of condition sets. A lookup table is created with the plurality of result sets. A value of Pfor given values of P, P, and Nis estimated through employment of the lookup table.

Estimation of probability of lambda failure through employment of lookup table

An apparatus comprising a position system configured to determine a global positioning (GPS) change vector, and an inertial navigation system (INS) change vector. The position system is configured to determine a relative position between a first platform and a second platform based on the GPS change vector and the INS change vector, and the position system is configured to correct a GPS position calculation error based on the GPS change vector and the INS change vector.

Ins based GPS carrier phase cycle slip detection and repairing

A first navigation unit of an apparatus in one example comprises a hybrid global positioning system (GPS)/inertial navigation system (INS) component. The hybrid GPS/INS component provides first GPS information and first INS information for the first navigation unit. The first navigation unit is configured to receive second GPS information and second INS information from a second navigation unit. The second navigation unit comprises a GPS component configured to determine a double-differenced GPS carrier phase measurement through employment of the first GPS information and the second GPS information. The first navigation unit comprises a relative Kalman filter configured to update an INS error estimate for an estimate of a relative position between the first and second navigation units based on the double-differenced GPS carrier phase measurement and the first and second INS information.

Estimate of relative position between navigation units

A method for attitude alignment of a slave inertial measurement system connected to a rotationally mobile platform supported by a vehicle that is stationary relative to a reference navigation frame comprises the steps of mounting a master reference inertial measurement system on the rotationally mobile platform and determining a master reference system attitude using measurements of acceleration and angular rates of the master reference inertial measurement system relative to the reference navigation frame. A slave system attitude is determined using measurements of acceleration and angular rates of the slave inertial measurement system relative to a slave system navigation reference frame and comparing the slave system attitude to the master reference system attitude to determine an attitude difference. The attitude difference is processed to obtain a correction to the slave system attitude.

Attitude alignment of a slave inertial measurement system

A phase error and/or amplitude error of a global positioning system carrier signal is estimated through employment of an optimal minimum variance of the global positioning system carrier signal.

Carrier signal phase error estimation

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 outputting 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, which actuates a divert propulsion system for guiding the projectile to a predetermined location.

Projectile guidance with accelerometers and a GPS receiver

The invention is a method continuing over a plurality of time periods for compensating for the output error in each of one or more navigation instruments in a system comprising a plurality of navigation instruments after the system is introduced into its operating environment. The practice of the method begins with determining the values of one or more of a set of coordinates that specify the position, velocity, and orientation of the system in space together with the error in a compensated output for each of the one or more navigation instruments. The method continues with determining a compensation model for each of the one or more navigation instruments. A compensation model specifies for a current time period an adjustment in amplitude of the output of a navigation instrument as a function of time and temperature. A compensation model is expressed as the sum of a first compensation model and a present second compensation model where the present second compensation model is a future second compensation model determined during the prior time period. The final steps of the method consist of measuring the time and the temperature of the navigation instruments and then obtaining a compensated output for each of the one or more navigation instruments by adjusting the output of a navigation instrument in accordance with its compensation model.

Temperature compensation method for strapdown inertial navigation systems

The present invention is a method and apparatus for detecting torpedo guidance system failures which are likely to cause the torpedo to make a circular run. The method consists of measuring the yaw-axis angular velocity of the torpedo and comparing this measurement with a computed estimate of the yaw-axis angular velocity based on measurements of torpedo dynamics other than the yaw-axis angular velocity. If the difference between the measured and computed values exceeds a threshold value, the presumption is that the torpedo is in a circular run and should be destroyed. The invention envisions various levels of precision in computing the estimate of the yaw-axis angular velocity. The various levels of precision involve the measurement of one or more of the group of dynamics parameters consisting of the three components of acceleration and the two components of angular velocity along the pitch and roll axes. Measured values of these dynamics parameters in various combinations together with other data such as gravity, earth angular velocity, torpedo elevator and rudder angles, propeller angular velocity, and the coefficients in the equations that define the dynamics of a torpedo permit the computation of the torpedo yaw-axis angular-velocity with various levels of precision.

Method and apparatus for detecting circular torpedo runs

The invention is a method for obtaining observables for input to a Kalman filter process which determines the attitude (roll, pitch, and heading) of a platform. The invention utilizes an inertial measurement unit (IMU) attached to the platform and an associated processor, a plurality of signal receiving antennas attached to the platform, and a plurality of satellite transmitters. The heading of the platform as determined by the IMU and its associated processor by themselves can be significantly in error. A comparison of the values of an attitude-sensitive function of the ranges from the platform antennas to different groupings of satellite transmitters obtained first by using IMU data and second by using the measured phases of the satellite-transmitter signals received at the platform antennas, a very accurate value for the range function is obtained. This accurate value of the range function is used in a Kalman filter process to obtain very accurate values for platform attitude.

Attitude determination utilizing an inertial measurement unit and a

The invention is a method for determining the relative position and velocity of a second vehicle with respect to a first vehicle. The method is practiced on board the second vehicle and utilizes a plurality of satellite transmitters and an inertial measurement unit on the second vehicle. The first vehicle has available a plurality of data items consisting of the first vehicle's measured position, measured velocity, and specific-force acceleration. The second vehicle has available data items consisting of the second vehicle's measured position, measured velocity, specific-force acceleration, autonomous position, autonomous velocity, autonomous position correction, and autonomous velocity correction. The measured position and measured velocity are determined utilizing the radio signals from a plurality of satellite transmitters and the specific-force acceleration is obtained from an inertial measurement unit. The autonomous position, autonomous velocity, autonomous position correction, and autonomous velocity correction are outputs of an autonomous Kalman filter process, the inputs being derived from the second vehicle's measured position, measured velocity, and specific-force acceleration. The method comprises the steps of (a) obtaining a plurality of the first vehicle's data items from the first vehicle and (b) determining the estimated relative position and the estimated relative velocity of the second vehicle with respect to the first vehicle utilizing the plurality of data items obtained from the first vehicle and one or more of the second vehicle's data items including either the specific-force acceleration or the group of data items consisting of autonomous position, autonomous velocity, autonomous position correction, and autonomous velocity correction.

Relative navigation utilizing inertial measurement units and a plurality

The navigation apparatus with improved attitude determination is intended for use on a mobile platform. It combines data from a platform inertial navigation unit and carrier phase data for a plurality of GPS satellite signals received at a plurality of receiving points on the platform for the purpose of obtaining estimates of the navigation states of the platform. The navigation apparatus comprises a processor which computes estimates of the states from inputs comprising (1) one or more measured phase double-differences calculated from the measured satellite signal carrier phases, (2) the estimated position of the platform, (3) the estimated positions of the receiving points, and (4) the positions of the satellites, a phase double-difference being defined as the difference in phase differences for signals received from two satellites and a phase difference being defined as the difference in carrier phase of a satellite signal received at two receiving points. The navigation and attitude determination solution is obtained by means of a Kalman filter process wherein the observables comprise measured phase double-differences and each predicted phase double-difference is obtained by operating on the Kalman state vector with an observation matrix defined by linearizing the equations relating the phase double-difference to the navigation states.

Growing community of inventors

Calabasas, CA, United States of America

Robert John Buchler