Real-time aircraft status detection system and method

• US 9,279,683 B2
• Filed: 03/02/2013
• Issued: 03/08/2016
• Est. Priority Date: 03/02/2012
• Status: Active Grant

• Current Assignee: Federal Express Corporation (FedEx Corporation)
• Original Assignee: Moog Incorporated
• Inventors: Turon, Martin A., Liccardo, Darren S.
• Primary Examiner(s): Caputo, Lisa
• Assistant Examiner(s): WILLIAMS, JAMEL E
• Application Number:
  US14/382,182
• Publication Number:
  US 20150027226A1
• Time in Patent Office: 1,102 Days
• Field of Search: 735/040.4
• US Class Current: 1/1
• CPC Class Codes: B64D 43/00 Arrangements or adaptations of instruments arrangements of cameras B64D47/08; aeronautical measuring instruments per se G01C
• CPC Class Codes: G01C 19/56 Turn-sensitive devices using vibrating masses, e.g. vibratory angular rate sensors based on Coriolis forces
• CPC Class Codes: H04Q 2209/00 Arrangements in telecontrol or telemetry systems
• CPC Class Codes: H04Q 9/00 Arrangements in telecontrol or telemetry systems for selectively calling a substation from a main station, in which substation desired apparatus is selected for applying a control signal thereto or for obtaining measured values therefrom

Abstract

A low power method for determining whether a cargo destined for air transport is in a flying state having the steps of: providing a housing for attachment to a cargo the housing having: an accelerometer for detecting a linear acceleration, a gyroscope for detecting an angular rate, a controller measuring a linear acceleration with the accelerometer, measuring an angular rate with the gyroscope, providing the measured linear acceleration and angular rate to the controller, and generating a flight status output signal indicating whether the housing is in a flying state as a function of the linear acceleration signal and angular rate signal.

Claims

1. A low power method for determining whether a cargo destined for air transport is in a flying state comprising the steps of:

providing a housing for attachment to a cargo, said housing comprising an accelerometer for detecting a linear acceleration, a gyroscope for detecting an angular rate and a controller;

measuring a linear acceleration with said accelerometer;

measuring an angular rate with said gyroscope;

providing said measured linear acceleration and angular rate to said controller; and

generating a flight status output signal indicating whether said housing is in a flying state as a function of said linear acceleration signal and said angular rate signal.

2. The method as set forth in claim 1, and further comprising the step of shutting off a radio as a function of said flight status output signal.

3. The method as set forth in claim 1, wherein said measured linear acceleration comprises a three independent axis linear acceleration signal and said measured angular rate comprises a three independent axis rotation rate signal.

4. The method as set forth in claim 1, wherein said flight status output signal comprises an in-flight status state and an on-the-ground status state.

5. The method as set forth in claim 4, and further comprising the step of providing a radio turn off signal when said flight status output signal comprises an in-flight status state.

6. The method as set forth in claim 1, wherein said function is a state machine algorithm comprising the steps of filtering said measured linear acceleration to produce a filtered linear acceleration, filtering said measured angular rate to produce a filtered angular rate, and changing a state as a function of said filtered linear acceleration and filtered angular rate.

7. The method as set forth in claim 1, and further comprising the step of powering said gyroscope as a function of said flight status output signal such that said gyroscope is powered off when said flight status output signal comprises an on-the-ground status state and said gyroscope is powered on when said flight status output signal comprises a verify-takeoff status state.

8. The method as set forth in claim 1, wherein said step of generating a flight status output signal comprises the steps of removing a gravitational component of said measured linear acceleration with a high pass filter having a cutoff frequency of about 0.01 Hz and saving said removed gravitational component.

9. The method as set forth in claim 1, wherein said step of generating a flight status output signal further comprises the step of detecting a linear acceleration characteristic of an aircraft takeoff.

10. The method as set forth in claim 9, wherein said linear acceleration characteristic of a takeoff is an acceleration signal with a frequency between 0.01 Hz and 0.1 Hz and a magnitude of between about 0.15 g and 0.5 g maintained for a duration of about two seconds.

11. The method as set forth in claim 9, and further comprising the step of saving a takeoff acceleration of said linear acceleration signal when said linear acceleration characteristic of an aircraft takeoff is detected and wherein said saved takeoff acceleration comprises a windowed average.

12. The method as set forth in claim 9, wherein said step of generating a flight status output signal further comprises the step of changing from an on-the-ground state to a verify-takeoff state when said linear acceleration characteristic of an aircraft takeoff is detected.

13. The method as set forth in claim 11, and further comprising the step of saving a gravitational acceleration component from said linear acceleration signal and wherein said gravitational acceleration is saved from a time period before a transition to a verity-takeoff state.

14. The method as set forth in claim 9, wherein said step of generating a flight status output signal further comprises the step of detecting an angular rate characteristic of an aircraft takeoff.

15. The method as set forth in claim 14, wherein said angular rate characteristic of an aircraft takeoff is an angular rate in which an x-axis angular rate, a y-axis angular rate, and a z-axis angular rate are all less than an angular rate limit threshold.

16. The method as set forth in claim 14, wherein said angular rate is transformed through a rotation, said rotation configured and arranged to reorient said measured angular rate into a yaw rate, a roll rate, and a pitch rate.

17. The method as set forth in claim 16, wherein said rotation is a cross product of said saved gravitational acceleration and said saved takeoff acceleration.

18. The method as set forth in claim 16, and further comprises the steps of integrating said angular rate with respect to time to produce a pitch angle displacement, a yaw angle displacement, and a roll angle displacement and setting said flight status output signal to an on-the-ground state when said pitch angle displacement, yaw angle displacement, or roll angle displacement are greater than an angle displacement threshold.

19. The method as set forth in claim 18, and further comprising the steps of:

providing a pressure sensor;

detecting when said pressure sensor has a pressure decrease rate less than a takeoff pressure decrease threshold; and

detecting when said pressure sensor has a pressure increase rate greater than a landing pressure increase threshold.

20. A system for detecting an aircraft flight status comprising:

an accelerometer for detecting a linear acceleration and having a linear-acceleration output signal;

a gyroscope for detecting an angular rate and having an angular rate output signal; and

a controller configured and arranged to produce a flight status output signal as a function of said linear acceleration output signal and said angular rate output signal, said flight status output signal comprising an in-flight state and second state; and

wherein said accelerometer, gyroscope and controller are in a housing.

21. The system as set forth in claim 20, and wherein said controller is further configured to shut off a radio as a function of said flight status output signal.

22. The system as set forth in claim 20, wherein said linear acceleration output signal comprises a three independent axis linear acceleration signal and said angular rate output signal comprises a three independent axis rotation rate signal.

23. The system as set forth in claim 20, wherein said second state of said flight status output signal comprises an on-the-ground status state.

24. The system set forth in claim 23, wherein said in-flight status state comprises a verify-takeoff status state.

25. The method as set forth in claim 24, wherein said controller is configured and arranged to power off said gyroscope when said flight status output signal comprises said on-the-ground status state and to power on said gyroscope when said flight status output signal comprises said verify-takeoff status state.

26. The system as set forth in claim 20, wherein said function comprises a filter configured and arranged to produce a filtered linear acceleration, and a filter configured and arranged to produce a filtered angular rate.

27. The system as set forth in claim 20, wherein said controller is configured and arranged to remove a gravitational component of said linear acceleration output signal, said system further comprises a high pass filter configured and arranged to remove said gravitational component from said linear acceleration output signal, wherein said high pass filter has a cutoff frequency of about 0.01 Hz, and wherein said controller is configured and arranged to save said removed gravitational component.

28. The system as set forth in claim 20, wherein said controller is configured and arranged to detect a linear acceleration characteristic of an aircraft takeoff.

29. The system as set forth in claim 28, wherein said linear acceleration characteristic of takeoff is an acceleration signal with a frequency between 0.01 Hz and 0.1 Hz and a magnitude of between about 0.15 g and 0.5 g maintained for a duration of about two seconds.

30. The system as set forth in claim 28, wherein said controller is configured and arranged to save said acceleration output signal when a linear acceleration characteristic of an aircraft is detected.

31. The system as set forth in claim 30, wherein said controller is configured and arranged to change said flight status output signal from an on-the-ground state to a verify-takeoff state when said linear acceleration characteristic of an aircraft takeoff is detected.

32. The system as set forth in claim 30, wherein said controller is configured and arranged to save a gravitational acceleration component from said linear acceleration signal and wherein said gravitational acceleration is saved from a time period before a transition to a verify-takeoff state.

33. The system as set forth in claim 30, wherein said controller is configured and arranged to detect an angular rate characteristic of an aircraft takeoff.

34. The system as set forth in claim 33, wherein said angular rate characteristic of an aircraft takeoff is an angular rate in which an x-axis angular rate, a y-axis angular rate, and a z-axis angular rate are all less than an angular rate limit threshold.

35. The system as set forth in claim 33, and further comprising a rotation transformer configured and arranged to transformed said angular rate output signal through a rotation, said rotation configured and arranged to reorient said measured angular rate into a yaw rate, a roll rate, and a pitch rate.

36. The system as set forth in claim 35, wherein said rotation is a cross product of said saved gravitational acceleration and said saved takeoff acceleration.

37. The system as set forth in claim 35, and further comprising an integrator configured and arranged to integrate said angular rate with respect to time to produce a pitch angle displacement, a yaw angle displacement, and a roll angle displacement and wherein said controller is configured and arranged to set said flight status output signal to an on-the-ground state when said pitch angle displacement, yaw angle displacement, or roll angle displacement are greater than an angle displacement threshold.

38. The system as set forth in claim 37, and further comprising a pressure sensor.

39. The system as set forth in claim 38, and wherein said controller is configured and arranged to detect when said pressure sensor has a pressure decrease rate less than a takeoff pressure decrease threshold and to detect when said pressure sensor has a pressure increase rate greater than a landing pressure increase threshold.