Real-time aircraft status detection system and method

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.