Geosynchronous spacecraft autonomous navigation

1. A spacecraft system comprising:

a measurement device configured to measure information associated with position of a spacecraft;

a filter configured to process the measured information and to provide estimated pre-maneuver information associated with position of the spacecraft and estimated post-maneuver information associated with position of the spacecraft;

a propagator configured to predict post-maneuver information associated with position of the spacecraft based upon the estimated pre-maneuver information associated with position of the spacecraft and a model of a maneuver of the spacecraft;

an error calculator configured to calculate an acceleration error using a state transition matrix, the acceleration error based upon the estimated post-maneuver information associated with position of the spacecraft and the predicted post-maneuver information associated with position of the spacecraft; and

a maneuver archive configured to store the acceleration error for updating the model of the maneuver,

wherein the state transition matrix is identified by Φ(t s ,t m ) and satisfies the following relationship:

[

δ

⁢

⁢

r

s

δ

⁢

⁢

v

s

]

=

Φ

⁡

(

t

s

,

t

m

)

⁡

[

δ

⁢

⁢

r

m

δ

⁢

⁢

v

m

]

,

 where

t m is the time of the maneuver of the spacecraft,

t s is a time after the maneuver of the spacecraft,

δr s is a difference between observed and predicted positions of the spacecraft at t s ,

δv s is a difference between observed and predicted velocities of the spacecraft at t s ,

δr m is a difference between observed and predicted positions of the spacecraft at t m , and

δv m is a difference between observed and predicted velocities of the spacecraft at t m .

2. The spacecraft system of claim 1 , wherein each of the measured information, the estimated pre-maneuver information, the estimated post-maneuver information and the predicted post-maneuver information includes one or more of: a position of the spacecraft and a velocity of the spacecraft.

3. The spacecraft system of claim 1 , wherein the filter is an Extended Kalman Filter (“EKF”).

4. The spacecraft system of claim 1 , wherein the propagator is an orbital propagator.

5. The spacecraft system of claim 1 , wherein the filter, the propagator, the error calculator and the maneuver archive are implemented in software.

6. The spacecraft system of claim 1 , further comprising the spacecraft.

7. The spacecraft system of claim 1 , wherein the filter, the propagator, the error calculator and the maneuver archive are implemented onboard the spacecraft.

8. The spacecraft system of claim 1 , wherein the spacecraft is in a geosynchronous orbit.

9. The spacecraft system of claim 1 , wherein the measurement device is a GPS receiver onboard the spacecraft.

10. The spacecraft system of claim 1 , wherein the error calculator is further configured to multiply the acceleration error by a weighting variable K before the model of the maneuver is updated.

11. The spacecraft system of claim 10 , wherein K is between 0.25 and 0.75.

12. The spacecraft system of claim 1 , wherein the spacecraft includes one or more thrusters, and wherein the model of the maneuver includes performance information about the one or more thrusters.

13. A system for calibrating spacecraft maneuvers comprising a processor configured to receive information associated with position of a spacecraft having one or more thrusters, to provide estimated pre-maneuver position-related information about the spacecraft and estimated post-maneuver position-related information about the spacecraft, to predict post-maneuver position-related information about the spacecraft based upon the estimated pre-maneuver position-related information about the spacecraft and a model of a maneuver of the spacecraft, the model of the maneuver of the spacecraft including performance information about the one or more thrusters, to calculate an acceleration error using a state transition matrix, and to update the model of the maneuver, wherein the acceleration error is based upon the estimated post-maneuver position-related information about the spacecraft and the predicted post-maneuver position-related information about the spacecraft,

wherein the state transition matrix is identified by Φ(t s ,t m ) and satisfies the following relationship:

[

δ

⁢

⁢

r

s

δ

⁢

⁢

v

s

]

=

Φ

⁡

(

t

s

,

t

m

)

⁡

[

δ

⁢

⁢

r

m

δ

⁢

⁢

v

m

]

,

 where

t m is the time of the maneuver of the spacecraft,

t s is a time after the maneuver of the spacecraft,

δr s is a difference between observed and predicted positions of the spacecraft at t s ,

δv s is a difference between observed and predicted velocities of the spacecraft at t s ,

δr m is a difference between observed and predicted positions of the spacecraft at t m , and

δv m is a difference between observed and predicted velocities of the spacecraft at t m .

14. The system of claim 13 , wherein the information associated with position of about the spacecraft includes one or more of: a position of the spacecraft and a velocity of the spacecraft.

15. A method for calibrating spacecraft maneuvers, comprising the steps of:

generating, with a filter, estimated pre-maneuver information associated with position of a spacecraft;

predicting, with a propagator, post-maneuver information associated with position of the spacecraft based upon the estimated pre-maneuver information associated with position and a model of a maneuver;

performing the maneuver;

generating, with the filter, estimated post-maneuver information associated with position of the spacecraft;

calculating an acceleration error using a state transition matrix, the acceleration error based upon a difference between the estimated post-maneuver information associated with position and the predicted post-maneuver information associated with position; and

updating the model of the maneuver based upon the calculated acceleration error,

wherein the state transition matrix is identified by Φ(t s ,t m ) and satisfies the following relationship:

[

δ

⁢

⁢

r

s

δ

⁢

⁢

v

s

]

=

Φ

⁡

(

t

s

,

t

m

)

⁡

[

δ

⁢

⁢

r

m

δ

⁢

⁢

v

m

]

,

 where

t m is the time of the maneuver of the spacecraft,

t s is a time after the maneuver of the spacecraft,

δr s is a difference between observed and predicted positions of the spacecraft at t s ,

δv s is a difference between observed and predicted velocities of the spacecraft at t s ,

δr m is a difference between observed and predicted positions of the spacecraft at t m , and

δv m is a difference between observed and predicted velocities of the spacecraft at t m .

16. The method of claim 15 , wherein each of the estimated pre-maneuver information, the estimated post-maneuver information and the predicted post-maneuver information includes one or more of: a position of the spacecraft and a velocity of the spacecraft.

17. The method of claim 15 , wherein the filter is an Extended Kalman Filter (“EKF”).

18. The method of claim 15 , wherein the propagator is an orbital propagator.

19. The method of claim 15 , wherein the estimated pre-maneuver position and velocity information and the estimated post-maneuver position and velocity information are generated based on GPS receiver measurements of position and velocity of the spacecraft.

20. The method of claim 15 , wherein the filter and the propagator are implemented in software.

21. The method of claim 15 , wherein the filter and the propagator are implemented onboard the spacecraft.

22. The method of claim 15 , wherein the calculating the acceleration error and the updating the model are performed onboard the spacecraft.

23. The method of claim 15 , wherein updating the model of the maneuver includes modifying the model by a weighting fraction K of the calculated acceleration error.

24. The method of claim 15 , wherein the model of the maneuver includes performance information about one or more thrusters of the spacecraft.