Honeywell International Inc.

Honeywell G.m.b.h.

Kevin D Vanderwerf

Mats Anders Brenner

Lawrence C Vallot

Randolph G Hartman

James E Lenz

Kent Stange

Mark A Ahlbrecht

Mark T Manfred

Subhabrata Ganguli

Karl Abraham Keyzer

Scot Griffith

Mark Hickey

Brenner Mats

Systems and methods for providing automatic detection of inertial sensor deployment environments are provided. In one embodiment, an environment detection system for a device having an inertial measurement unit that outputs a sequence of angular rate measurements comprises: an algorithm selector; and a plurality of environment detection paths each receiving the sequence of angular rate measurements, and each generating angular oscillation predictions using an environment model optimized for a specific operating environment. The environment model for each of the environment detection paths is optimized for a different operating environment. Each of the environment detection paths outputs a weighting factor that is a function of a probability that its environment model is a true model of a current operating environment given the sequence of angular rate measurements; and wherein the algorithm selector generates an output based on a function of the weighting factor from each of the environment detection paths.

Systems and methods for providing automatic detection of inertial sensor deployment environments

A navigation system to transition from a stationary alignment filter to an in-motion alignment filter is provided. The system comprises a processing unit configured to implement a stationary alignment Kalman filter (SAKF) in gyrocompass alignment mode to generate state estimates and provide corrections when the object is stationary, and to implement an algorithm to compute a covariance for the SAKF that accounts for uncertainty in the SAKF estimates; wherein the processing unit is further configured to implement a continuous alignment filter (CAF) that generates a secondary solution which remains unaffected by the SAKF corrections during a delay period accommodating a delay between the time of actual motion to the time of detected motion, and to implement an algorithm to compute a covariance for CAF that accounts for the uncertainty in CAF during delay period; and wherein outputs of the CAF and its covariance are communicated to an in-motion alignment filter.

Transitioning from stationary alignment mode to in-motion alignment mode

A method to calibrate at least one lever arm between at least one respective global positioning system (GPS) antenna/receiver and a communicatively coupled inertial navigation system is provided. The method includes receiving signals from the at least one GPS antenna/receiver at the inertial navigation system communicatively coupled to a Kalman filter; and estimating, in the Kalman filter, at least one fixed lever arm component while accounting for a bending motion of the lever arm based on the received signals.

Global positioning system (GPS) self-calibrating lever arm function

One embodiment is directed towards a method for providing integrity for a hybrid navigation system using a Kalman filter. The method includes determining a main navigation solution for one or more of roll, pitch, platform heading, or true heading for the vehicle using signals from a plurality of GNSS satellites and inertial measurements. Solution separation is used to determine a plurality of sub-solutions for the main navigation solution. The method also includes determining a separation between the main navigation solution and each of the sub-solutions, and a discriminator for each of the separations. The method also includes determining a separation variance between the main navigation solution and each of the sub-solutions, a threshold for detection of a satellite failure based on the separation variances; and a protection limit that bounds error in the main navigation solution based on the threshold.

Method and system for providing integrity for hybrid attitude and true heading

Systems and methods for takeoff assistance and analysis are provided. In one embodiment, a takeoff ground roll assist system for an aircraft comprises: a runway centerline estimator, wherein the runway centerline estimator generates a virtual runway centerline estimate for a runway based on coordinates for the runway from a runway database; an (Inertial Navigation System) INS Output Filter and Fault Detector wherein the filtering of navigation measurements from an on-board inertial navigation system, using a runway centerline start point as established by activation of a pilot operated start point sensor occurs, a centerline tracking estimator coupled to receive a filtered output of the on-board inertial navigation system from the INS output filter and fault detector and to receive the virtual centerline estimate from the runway centerline estimator, wherein the centerline tracking estimator generates a centerline tracking feedback signal that varies as a function of a difference between an aircraft look-ahead point and the virtual runway centerline; and a takeoff criteria evaluator coupled to the runway centerline estimator and the centerline tracking estimator, wherein the takeoff criteria evaluator estimates a lateral deviation between the aircraft and an actual runway centerline for the runway based on an estimated navigation error component, the runway database error component, and a centerline tracking error component.

Systems and methods for takeoff assistance and analysis

A navigation system for a vehicle having a receiver operable to receive a plurality of signals from a plurality of transmitters includes a processor and a memory device. The memory device has stored thereon machine-readable instructions that, when executed by the processor, enable the processor to determine a set of error estimates corresponding to pseudo-range measurements derived from the plurality of signals, determine an error covariance matrix for a main navigation solution using ionospheric-delay data, and, using a parity space technique, determine at least one protection level value based on the error covariance matrix.

Navigation system with apparatus for detecting accuracy failures

A navigation system for a vehicle having a receiver operable to receive a plurality of signals from a plurality of transmitters includes a processor and a memory device. The memory device has stored thereon machine-readable instructions that, when executed by the processor, enable the processor to determine a set of error estimates corresponding to pseudo-range measurements derived from the plurality of signals, determine an error covariance matrix for a main navigation solution using ionospheric-delay data, and, using a solution separation technique, determine at least one protection level value based on the error covariance matrix.

A method for characterizing distortions in the earth's magnetic field caused by a vehicle having a magnetometer affixed therein is described. The method includes repeatedly measuring the distorted magnetic field utilizing the magnetometer and obtaining a three-dimensional orientation of the vehicle axes with respect to the earth at a time of each magnetometer measurement. The method also includes receiving undistorted earth magnetic field data for the vicinity of the vehicle relative to the earth at the time of each magnetometer measurement and characterizing distortions caused by one or more of the vehicle and magnetometer errors utilizing the magnetic field measurements, the orientations of the vehicle, and the undistorted earth magnetic field data.

Methods and apparatus for automatic magnetic compensation

A global positioning system for use with a plurality of transmitting satellites is disclosed wherein a full position solution is obtained from a number N of satellite transmissions and a plurality of sub-solutions each based on a different number (N&minus;1) of satellite transmissions are obtained to determine if a satellite is faulty. A plurality of sub-sub-solutions is then obtained based on a different combination of N&minus;2 satellites with each combination containing the satellites of different sub-solutions and excluding one other satellite in addition. An examination of the sub-sub-solutions provides an indication of which one of the satellites is the failed satellite.

Fault detection and exclusion for global position systems

The novel navigation system, suited for commercial aircraft, includes a sensor for outputting a signal representing a first navigation parameter, a memory having condensed gravity compensation data, and a processor for deriving a second navigational parameter. The processor derives the second navigational parameter from the signal and a subset of the condensed compensation data based on aircraft position. The condensed memory, based on regional gravity deflections, stores sequential gravity compensation values and corresponding navigational coordinates. To conserve memory, adjacent compensation values differ by at least a preset gravity compensation increment. The compensation increment, based on a gravity deflection between 30 and 100 micro-radians, is adequate for commercial aircraft.

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Oro Valley, AZ, United States of America

Kevin D Vanderwerf