Method and system for in-flight fault monitoring of flight...

Details Method and system for in-flight fault monitoring of flight... Method and system for in-flight fault monitoring of flight...

2001-10-31

2003-09-23

Jordan, Charles T. (Department: 3644)

Aeronautics and astronautics

Aircraft control

Automatic

C244S195000

Reexamination Certificate

active

06622972

ABSTRACT:

FIELD OF THE INVENTION
This invention relates generally to fault detection of flight control actuators and, more specifically, to a method and system for monitoring and testing in-flight fault characteristics of flight control actuators without producing adverse in-flight motion.
BACKGROUND OF THE INVENTION
Aircraft actuators are integral components for the maintenance of proper operational flight control. The actuators control moveable flight surfaces such as ailerons, flaps, leading-edge slats, spoilers, horizontal stabilizer, elevators, and rudders. Without accurate and timely fault detection and diagnostics of the actuators, catastrophic results may occur. Correctly repairing a damaged actuator depends on such accuracy and timeliness.
Although actuator fault detection is crucial, it remains problematic because in most cases actuators must be tested post-flight, on the ground. The data collected in post-flight diagnostics is not as reliable as in-flight diagnostics in that it does not allow for operating flight loads on the actuation system, leading to incorrect diagnostics and costly replacement of the wrong actuator or actuator component. The problem is more acute for Uninhabited Air Vehicles where a flight crew is not present to deal with the failure situation and the failure could cause the loss of the vehicle.
SUMMARY OF THE INVENTION
This invention comprises a method and system for monitoring in-flight fault characteristics of flight control actuators. The system includes an electronic flight control system in communication with a fault detection system. The electronic flight control system includes a flight computer configured to send command signals to at least one actuator in response to flight commands. The electronic flight control system further includes a plurality of aircraft response sensors that measure aircraft attitude such as pitch, roll, yaw, or other aspects.
The electronic flight control system sends command signals to the actuators to affect a flight surface associated with each actuator. As the electronic flight control system issues each command, the computer processors associated with the electronic flight control system compute the appropriate signal command and further send the command signal to the, affected actuator. The actuator responds by adjusting, if necessary, the affected flight surface.
The flight surfaces include ailerons, flaps, leading-edge slats, spoilers, horizontal stabilizer, elevators, and rudders or any other flight surface capable of affecting the aerodynamics of an aircraft.
The fault detection system includes a computer processor configured to perform inflight operations on monitored aircraft data. The fault detection system computes and sends its own signals to at least one actuator through the electronic flight control system. If the actuators are functioning properly, the actuators will move in a canceling manner, with no net motion to the aircraft. However, when a response sensor detects an unexpected aircraft motion it is indicative of a possible actuator failure.
In accordance with other aspects of the invention that detects a possible actuator failure, the various aircraft actuators are tested sequentially, one or more at a time, cycling through the aircraft until all actuators have been tested.
In accordance with further aspects of the invention, upon identifying a potentially failed actuator, an additional test is performed to verify the failure. Preset test commands are sent to one or more suspect actuators and surface position data, as well as aircraft motion sensors, are monitored. During these tests, the fault detection system sends signals to reset the remaining aircraft flight surfaces to compensate for the preset test commands on the suspect actuator. This compensation is achieved by activating a reconfigurable control mode that reconfigures other actuators so that the test commands cause no perceptible aircraft motion beyond those commanded by the pilot (or autopilot). This subsequent confirmation test allows confirmation that the suspected actuator has actually failed, without disrupting the mission capability of the vehicle.
In accordance with still other aspects of the invention, the fault detection system isolates and removes the failed actuator from the rest of the electronic flight control system. Further, the fault detection system sends command signals to reset the remaining aircraft flight surfaces to compensate for the removed actuator by activating a reconfigurable control mode that reconfigures other actuators to maintain safe and controllable handling qualities for the aircraft.


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