# US 9,435,661 B2 · Systems and methods for attitude fault detection based on air data and aircraft control settings > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 9,435,661 B2 - Document ID: US9435661B2 - Status: Granted - Kind: B2 - Application number: 14/564,344 - Filed: 2014-12-09 - Published: 2016-04-14 - Granted: 2016-09-06 - Earliest claimed: 2014-10-08 - Assignees: Honeywell International Inc. - Applicants: Honeywell International Inc. - Inventors: Mats Anders Brenner; John R. Morrison; Danny Thomas Kimmel; Jay Joseph Hansen - Web page: https://patentgrep.com/patent/US9435661B2 ## CPC - G01C 23/005: {Flight directors (indicating arrangements specially adapted for rotary gyroscopes G01C19/32)} - B64D 45/00: Aircraft indicators or protectors not otherwise provided for - G05D 1/0077: {using redundant signals or controls} - G05D 1/0808: {specially adapted for aircraft} - G07C 5/0808: {Diagnosing performance data (testing of vehicles G01M17/00; testing of electrical installation on vehicles G01R31/005)} - G01C: MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY (measuring liquid level G01F; radio navigation, determining distance or velocity by use of propagation effects, e.g. Doppler effects, propagation time, of radio waves, analogous arrangements using other waves G01S) - B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT - G05D: SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES - G07C: TIME OR ATTENDANCE REGISTERS; REGISTERING OR INDICATING THE WORKING OF MACHINES; GENERATING RANDOM NUMBERS; VOTING OR LOTTERY APPARATUS; ARRANGEMENTS, SYSTEMS OR APPARATUS FOR CHECKING NOT PROVIDED FOR ELSEWHERE ## Abstract Systems and methods for attitude fault detection based on air data and aircraft control settings are provided. In one embodiment, a sensor monitor for an aircraft attitude measurement system comprises: an aircraft model configured to model a plurality of states, the plurality of states including at least an aircraft attitude state, an aircraft velocity state, a sink rate error state, and a wind velocity state; a propagator-estimator configured to utilize the plurality of states of the aircraft model to process air data measurements and attitude measurements from a first inertial measurement unit of the aircraft attitude measurement system; and a residual evaluator configured to input residual error values generated by the propagator-estimator, wherein the residual evaluator outputs an alert signal when the residual error values exceed a predetermined statistical threshold. ## Claims ### Claim 1 A sensor monitor for an aircraft attitude measurement system, the sensor monitor comprising: an aircraft model configured to model a plurality of states, the plurality of states including at least an aircraft attitude state, an aircraft velocity state, a sink rate error state, and a wind velocity state; a propagator-estimator configured to utilize the plurality of states of the aircraft model to process air data measurements and attitude measurements from a first inertial measurement unit of the aircraft attitude measurement system; and a residual evaluator configured to input residual error values generated by the propagator-estimator, wherein the residual evaluator outputs an alert signal when the residual error values exceed a predetermined statistical threshold; and wherein the alert signal initiates an alert on a device that indicates that the first inertial measurement unit is faulted. ### Claim 2 (depends on 1) The sensor monitor of claim 1, wherein the aircraft velocity state time derivative is calculated as a function of one or more of the aircraft's angle-of-attack, rudder settings, thrust setting, attitude and velocity. ### Claim 3 (depends on 1) The sensor monitor of claim 1, wherein the sink rate error state time derivative is calculated as a function of barometric altimeter measurements. ### Claim 4 (depends on 3) The sensor monitor of claim 3, wherein the sink rate error state defines a stochastic process representing an error in the sink rate measurement. ### Claim 5 (depends on 1) The sensor monitor of claim 1, wherein the wind velocity state time derivative is calculated as a function of True Air Speed as obtained from aircraft sensor data. ### Claim 6 (depends on 5) The sensor monitor of claim 5, wherein the wind velocity state defines a stochastic process representing an error in the true air speed measurement. ### Claim 7 (depends on 1) The sensor monitor of claim 1, wherein the alert signal produces an alert on a display that indicates that the first inertial measurement unit is faulted. ### Claim 8 (depends on 1) The sensor monitor of claim 1, wherein the sensor monitor is internal to the first inertial measurement unit. ### Claim 9 A fault detection system for aircraft attitude measurement system, the fault detection system comprising: a sensor monitor coupled to a first inertial measurement unit of the aircraft attitude measurement system, the sensor monitor comprising: an aircraft model of an aircraft, the aircraft model configured to model a plurality of aircraft states, the plurality of aircraft states including at least an aircraft attitude state, an aircraft velocity state, a sink rate error state, and a wind velocity state; a propagator-estimator configured to propagate and update the plurality of aircraft states of the aircraft model based on air data measurements and attitude measurements from the first inertial measurement unit; and a residual evaluator coupled to the propagator-estimator and configured to input measurement error residual values generated by the propagator-estimator, wherein the residual evaluator outputs an alert signal when the measurement error residual values exceed a predetermined statistical threshold; and a status device, wherein the alert signal produces an alert on the status device that indicates that the first inertial measurement unit is faulted. ### Claim 10 (depends on 9) The fault detection system of claim 9, wherein the sensor monitor is internal to the first inertial measurement unit. ### Claim 11 (depends on 9) The fault detection system of claim 9, wherein the propagator-estimator is a Kalman filter. ### Claim 12 (depends on 9) The fault detection system of claim 9, the status device comprising a display; wherein the alert signal produces an alert on the display that indicates that the first inertial measurement unit is faulted. ### Claim 13 (depends on 9) The fault detection system of claim 9, wherein aircraft attitude state includes one or both of an aircraft pitch position and an aircraft roll position. ### Claim 14 (depends on 9) The fault detection system of claim 9, wherein the aircraft attitude state time derivative is calculated as a function of one or more of the aircraft's angle-of-attack, rudder settings, thrust setting, attitude and velocity; wherein the aircraft velocity state time derivative is calculated as a function of one or more of the aircraft's angle-of-attack, rudder settings, thrust setting, attitude and velocity; wherein the sink rate error state time derivative is calculated as a function of barometric altimeter measurements; and wherein the wind velocity state time derivative is calculated as a function of True Air Speed as obtained from aircraft sensor data. ### Claim 15 A fault detection method for an aircraft attitude measurement system, method comprising: monitoring attitude solution data generated by a first inertial measurement unit of an aircraft attitude measurement system; executing a propagator-estimator configured with an aircraft model for a plurality of aircraft states based on an aircraft attitude state vector, a velocity state vector, a Sink Rate Error state vector, and a wind velocity state vector; generating measurement error residual values using the propagator-estimator, wherein the propagator-estimator is configured to iteratively predict and update the plurality of aircraft states of the aircraft model; and comparing the measurement error residual values against a predetermined statistical threshold and generating an alert signal on a device that indicates that the first inertial measurement unit is faulted when the measurement error residual values exceed the predetermined statistical threshold. ### Claim 16 (depends on 15) The method of claim 15, wherein the propagator-estimator is a Kalman filter. ### Claim 17 (depends on 15) The method of claim 15, wherein the alert signal produces an alert on a display that indicates that the first inertial measurement unit is faulted. ### Claim 18 (depends on 15) The method of claim 15, wherein the aircraft attitude state vector time derivative is calculated as a function of one or more of the aircraft's angle-of-attack, rudder settings, thrust setting, attitude and velocity; wherein the aircraft velocity state vector time derivative is calculated as a function of one or more of the aircraft's angle-of-attack, rudder settings, thrust setting, attitude and velocity; wherein the sink rate error state vector time derivative is calculated as a function of barometric altimeter measurements; and wherein the wind velocity state vector time derivative is calculated as a function of True Air Speed as obtained from aircraft sensor data. ### Claim 19 (depends on 18) The method of claim 18, wherein the wind velocity state vector defines a stochastic process representing error in the true air speed measurement; and wherein the sink rate error state vector defines a stochastic process representing error in the sink rate measurement. ### Claim 20 (depends on 15) The method of claim 15, wherein aircraft attitude state includes one or both of an aircraft pitch position and an aircraft roll position.