IPC분류정보
국가/구분 |
United States(US) Patent
등록
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국제특허분류(IPC7판) |
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출원번호 |
US-0178546
(2002-06-25)
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발명자
/ 주소 |
- Julie, Thienel K.
- Richard, Harman R.
- Bar-Itzhack, Itzhack Y.
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출원인 / 주소 |
- The United States of America as represented by the Administrator of the National Aeronautics and Space Administration
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대리인 / 주소 |
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인용정보 |
피인용 횟수 :
3 인용 특허 :
7 |
초록
▼
This invention is drawn to an autonomous navigation system using Global Positioning System (GPS) and magnetometers for low Earth orbit satellites. As a magnetometer is reliable and always provides information on spacecraft attitude, rate, and orbit, the magnetometer-GPS configuration solves GPS init
This invention is drawn to an autonomous navigation system using Global Positioning System (GPS) and magnetometers for low Earth orbit satellites. As a magnetometer is reliable and always provides information on spacecraft attitude, rate, and orbit, the magnetometer-GPS configuration solves GPS initialization problem, decreasing the convergence time for navigation estimate and improving the overall accuracy. Eventually the magnetometer-GPS configuration enables the system to avoid costly and inherently less reliable gyro for rate estimation. Being autonomous, this invention would provide for black-box spacecraft navigation, producing attitude, orbit, and rate estimates without any ground input with high accuracy and reliability.
대표청구항
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1. An autonomous navigation system for determining a set of satellite navigation parameters including attitude, orbit, and rate, the system comprising:a GPS receiver;a magnetometer;processing means for processing measurements from the GPS and magnetometer to estimate attitude, orbit, and rate. 2. Th
1. An autonomous navigation system for determining a set of satellite navigation parameters including attitude, orbit, and rate, the system comprising:a GPS receiver;a magnetometer;processing means for processing measurements from the GPS and magnetometer to estimate attitude, orbit, and rate. 2. The autonomous navigation system as in claim 1, wherein the measurements from the GPS comprises:GPS pseudo-range; and GPS phase measurements. 3. The autonomous navigation system as in claim 2, wherein the measurements from the GPS further comprises GPS range rate. 4. The autonomous navigation system as in claim 1, the processing means comprising:an algorithm comprising a Kalman filter. 5. The autonomous navigation system as in claim 4, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate. 6. The autonomous navigation system as in claim 5, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements or magnetometer measurements. 7. The autonomous navigation system as in claim 5, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements and magnetometer measurements. 8. The autonomous navigation system as in claim 4, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , b T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate, and b is the magnetic field in body coordinates. 9. The autonomous navigation system as in claim 8, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements. 10. The autonomous navigation system as in claim 4, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , z T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate, and z is GPS phase measurement. 11. The autonomous navigation system as in claim 10, wherein the rate estimate is obtained by applying derivative approach to magnetometer measurements. 12. The autonomous navigation system as in claim 4, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , b T , z T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, b is the attitude quaternion, ω is the rotation rate, b is the magnetic field in body coordinates, and z is GPS phase measurement. 13. A method for estimating a set of navigation parameters of a satellite including attitude, orbit, and rate, the method comprising:providing measurements from GPS;providing measurements from magnetometer;executing an algorithm to estimate attitude, orbit, and rate of the satellite based on the measurements from the GPS and magnetometer. 14. The method as in claim 13, wherein the measurements from the GPS comprises:GPS pseudo-range; andGPS phase measurements. 15. The method as in claim 14, wherein the measurements from the GPS further comprises GPS range rate. 16. The method as in claim 13, wherein the step of executing the algorithm includes performing Kalman filtering. 17. The method as in claim 16, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T ],wherein R and V are the spacecraft position and velocity vectors, respectiv ely, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate. 18. The method as in claim 17, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements or magnetometer measurements. 19. The method as in claim 17, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements and magnetometer measurements. 20. The method as in claim 16, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , b T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate, and b is the magnetic field in body coordinates. 21. The method as in claim 20, wherein the rate estimate is obtained by applying derivative approach to GPS phase measurements. 22. The method as in claim 16, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , z T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors q is the attitude quaternion, ω is the rotation rate, and z is GPS phase measurement. 23. The method as in claim 22, wherein the rate estimate is obtained by applying derivative approach to magnetometer measurements. 24. The method as in claim 16, wherein output filter state of the Kalman filter includes: X T =[ R T , V T , C T , q T , ω T , b T , z T ],wherein R and V are the spacecraft position and velocity vectors, respectively, C is a vector of GPS receiver clock errors, q is the attitude quaternion, ω is the rotation rate, b is the magnetic field in body coordinates, and z is GPS phase measurement.
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