Methods, systems, and apparatus, including computer programs encoded on a computer storage medium, for receiving signals that represent inertial forces experienced by the aircraft from an inertial sensor mounted in an aircraft. Identifying a pattern within the signals that represents a reduction of
Methods, systems, and apparatus, including computer programs encoded on a computer storage medium, for receiving signals that represent inertial forces experienced by the aircraft from an inertial sensor mounted in an aircraft. Identifying a pattern within the signals that represents a reduction of thrust on one side of the aircraft caused by an engine failure. Causing an indication within the aircraft to be activated in response to identifying the pattern, and thereby, alerting a pilot to a possible engine failure.
대표청구항▼
1. A computer-implemented method for detecting aircraft engine failure in a multi-engine aircraft, the method executed by one or more processors and comprising: receiving, from an inertial sensor mounted in an aircraft, signals that represent inertial forces experienced by the aircraft, the signals
1. A computer-implemented method for detecting aircraft engine failure in a multi-engine aircraft, the method executed by one or more processors and comprising: receiving, from an inertial sensor mounted in an aircraft, signals that represent inertial forces experienced by the aircraft, the signals including at least one signal that represents an acceleration of the aircraft and at least one signal that represents a rotation of the aircraft;identifying, by the one or more processors, a pattern, within the signals, that represents a reduction of thrust on one side of the aircraft caused by an engine failure; andin response to identifying the pattern, causing an indication within the aircraft to be activated, and thereby, alerting a pilot to a possible engine failure. 2. The method of claim 1, wherein the inertial sensor is a 6-axis inertial sensor, and wherein the signals comprise three linear acceleration signals, each being along a different orthogonal axis, and three angular acceleration signals, each being about a different orthogonal axis. 3. The method of claim 1, wherein the pattern comprises: a change in magnitude for three of the signals that occurs, for each of the three signals, within a first time period, anda persistence in the magnitude change of two of the three signals for a duration of a second time period, after the first time period. 4. The method of claim 3, wherein the change in magnitude for each of the three signals is greater than or equal to a, respective, threshold value. 5. The method of claim 3, wherein the first time period is less than or equal to 200 ms. 6. The method of claim 3, wherein the second time period is between 1 second and 5 seconds. 7. The method of claim 3, wherein the persistence in the magnitude change of the two of the three signals is indicated by the respective magnitudes of the two signals remaining above a, respective, threshold value for the duration of the second time frame. 8. The method of claim 1, wherein identifying the pattern comprises: combining three of the signals into a combined signal; anddetermining whether the combined signal exceeds a threshold value. 9. The method of claim 1, further comprising determining, based on the signals, on which side of the aircraft the reduction of thrust occurred, wherein causing an indication within the aircraft to be activated comprises causing a first indication to be activated based on determining that the reduction in thrust occurred on the port side, and causing a second indication to be activated based on determining the reduction in thrust occurred on the starboard side. 10. The method of claim 1, further comprising: determining, from the signals, an average gravitational acceleration signal;determining, based on the average gravitational acceleration signal, an offset between a position of a z-axis of the aircraft relative to a z-axis of the inertial sensor; andmodifying the signals, based on the position of the z-axis of the aircraft relative to the z-axis of the inertial sensor, to account for the offset. 11. The method of claim 1, further comprising: receiving a user input to enter a calibration mode; andin response to identifying the pattern, modifying one or more threshold values and one or more time periods based on the pattern. 12. An engine failure detection system for multi-engine aircraft comprising: one or more processors;a data store coupled to the one or more processors having instructions stored thereon for execution by the one or more processors;an inertial sensor coupled to the one or more processors and configured to be mounted to an aircraft;a first warning light coupled to the one or more processors; anda second warning light coupled to the one or more processors,wherein the instructions, when executed by the one or more processors, causes the one or more processors to perform operations comprising: receiving, from the inertial sensor, signals that represent inertial forces experienced by the aircraft, the signals including at least one signal that represents an acceleration of the aircraft and at least one signal that represents a rotation of the aircraft;identifying a pattern, within the signals, that represents a reduction of thrust on one side of the aircraft caused by an engine failure; andin response to identifying the pattern, causing one of the first or second warning light to be activated, and thereby, alert a pilot to a possible engine failure. 13. The system of claim 12, wherein the system is a stand-alone system that is not integrated with other sensor or systems of the aircraft. 14. The system of claim 12, further comprising a wireless communication interface, and wherein the first warning light and the second warning light are each coupled to a wireless receiver; andwherein causing one of the first or second warning light to be activated comprises sending, by the wireless communication interface to the wireless receiver of the one of the first or second warning light, a command to activate the one of the first or second warning light. 15. The system of claim 12, wherein the inertial sensor is a 6-axis inertial sensor, and wherein the signals comprise three linear acceleration signals, each being along a different orthogonal axis, and three angular acceleration signals, each being about a different orthogonal axis. 16. The system of claim 12, wherein the pattern comprises: a change in magnitude for three of the signals that occurs, for each of the three signals, within a first time period, anda persistence in the magnitude change of two of the three signals for a duration of a second time period, after the first time period. 17. The system of claim 16, wherein the change in magnitude for each of the three signals is greater than or equal to a, respective, threshold value. 18. The system of claim 16, wherein the first time period is less than or equal to 200 ms. 19. The system of claim 16, wherein the second time period is between 1 second and 5 seconds. 20. The system of claim 16, wherein the persistence in the magnitude change of the two of the three signals is indicated by the respective magnitudes of the two signals remaining above a, respective, threshold value for the duration of the second time frame. 21. The system of claim 12, wherein identifying the pattern comprises: combining three of the signals into a combined signal; anddetermining whether the combined signal exceeds a threshold value. 22. The system of claim 12, wherein the operations comprise determining, based on the signals, on which side of the aircraft the reduction of thrust occurred, wherein causing an indication within the aircraft to be activated comprises causing a first indication to be activated based on determining that the reduction in thrust occurred on the port side, and causing a second indication to be activate based on determining the reduction in thrust occurred on the starboard side. 23. The system of claim 12, wherein the operations comprise: determining, from the signals, an average gravitational acceleration signal;determining, based on the average gravitational acceleration signal, an offset between a position of a z-axis of the aircraft relative to a z-axis of the inertial sensor; andmodifying the signals, based on the position of the z-axis of the aircraft relative to the z-axis of the inertial sensor, to account for the offset. 24. The system of claim 12, wherein the operations comprise: receiving a user input to enter a calibration mode; andin response to identifying the pattern, modifying one or more threshold values and one or more time periods based on the pattern. 25. A non-transitory computer readable storage device storing instructions that, when executed by one or more processors, cause the one or more processors to perform operations comprising: receiving, from an inertial sensor mounted in a multi-engine aircraft, signals that represent inertial forces experienced by the aircraft, the signals including at least one signal that represents an acceleration of the aircraft and at least one signal that represents a rotation of the aircraft;identifying, by the one or more processors, a pattern, within the signals, that represents a reduction of thrust on one side of the aircraft caused by an engine failure; andin response to identifying the pattern, causing an indication within the aircraft to be activated, and thereby, alerting a pilot to a possible engine failure.
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이 특허에 인용된 특허 (7)
Middleton Robin (Bellevue WA), Apparatus and method for warning of a high yaw condition in an aircraft.
Rice Robert W. (Sandy Hook CT) Sweet David H. (Tequesta FL), Engine failure monitor for a multi-engine aircraft having partial engine failure and driveshaft failure detection.
Hernandez-Diaz Jorge H. (Minillas Sta. Box 41267 Santurce PR 00940), Method and apparatus for limiting adverse yaw-induced roll during engine failure in multiengine aircraft.
Eggold, David P.; Shapiro, Daniel R.; Gardner, Kyle J., Varying engine thrust for directional control of an aircraft experiencing engine thrust asymmetry.
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