Systems and methods for attitude fault detection based on air data and aircraft control settings
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.
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.
2. 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.
3. The sensor monitor of claim 1 , wherein the sink rate error state time derivative is calculated as a function of barometric altimeter measurements.
4. The sensor monitor of claim 3 , wherein the sink rate error state defines a stochastic process representing an error in the sink rate measurement.
5. 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.
6. The sensor monitor of claim 5 , wherein the wind velocity state defines a stochastic process representing an error in the true air speed measurement.
7. 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.
8. The sensor monitor of claim 1 , wherein the sensor monitor is internal to the first inertial measurement unit.
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.
10. The fault detection system of claim 9 , wherein the sensor monitor is internal to the first inertial measurement unit.
11. The fault detection system of claim 9 , wherein the propagator-estimator is a Kalman filter.
12. 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.
13. 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.
14. 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.
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.
16. The method of claim 15 , wherein the propagator-estimator is a Kalman filter.
17. 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.
18. 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.
19. 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.
20. The method of claim 15 , wherein aircraft attitude state includes one or both of an aircraft pitch position and an aircraft roll position.