Highly inclined elliptical orbit de-orbit techniques

1. A method of deorbiting an earth-orbiting satellite comprising:

executing a first orbit transfer maneuver that transfers the satellite from an operational orbit to a first interim orbit; wherein:

the operational orbit is substantially geosynchronous and has (i) an inclination of greater than 70 degrees; (ii) a nominal eccentricity in the range of 0.25 to 0.5; (iii) an argument of perigee of approximately 90 or approximately 270 degrees; (iv) a right ascension of ascending node of approximately 0; and (v) an operational orbit apogee altitude; and

the first interim orbit has an initial second apogee altitude that is at least 4500 km higher than the first apogee altitude, and the interim orbit naturally decays, subsequent to the orbit transfer maneuver, such that the satellite will reenter Earth's atmosphere no longer than 25 years after completion of the orbit transfer maneuver.

2. The method of claim 1 , wherein executing the first orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by more than 60 m/sec.

3. The method of claim 1 , wherein executing the orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by approximately 65 m/sec.

4. The method of claim 1 , wherein the first interim orbit has an initial second apogee altitude that is approximately 5000 km higher than the operational orbit apogee altitude.

5. The method of claim 1 , wherein the right ascension of ascending node is 0+/−20 degrees.

6. The method of claim 1 , wherein the operational orbit has an orbital period of approximately 23.93 hours.

7. The method of claim 1 , wherein executing the first orbit transfer maneuver includes at least one firing of a chemical or electric thruster proximate to orbit perigee.

8. The method of claim 7 , wherein executing the first orbit transfer maneuver includes a plurality of thruster firings.

9. The method of claim 1 , further comprising:

executing, following a period of time in which the first interim orbit is allowed to decay, a second orbit transfer maneuver that transfers the satellite from the decayed first interim orbit to a second interim orbit, wherein the decayed first interim orbit has an ascending node radius less than 42,160 km and the second interim orbit has an ascending node radius greater than 42,170 km.

10. The method of claim 9 , wherein executing the second orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by approximately 7 m/sec.

11. An earth-orbiting satellite comprising a propulsion subsystem and a spacecraft controller, the spacecraft controller configured to:

cause the propulsion subsystem to execute a first orbit transfer maneuver that transfers the satellite from an operational orbit to an interim orbit; wherein:

the operational orbit is substantially geosynchronous and has (i) an inclination of greater than 70 degrees; (ii) a nominal eccentricity in the range of 0.25 to 0.5; (iii) an argument of perigee of approximately 90 or approximately 270 degrees; (iv) a right ascension of ascending node of approximately 0; and (v) an operational orbit apogee altitude; and

the interim orbit has an initial second apogee altitude that is at least 4500 km higher than the first apogee altitude, and the interim orbit naturally decays, subsequent to the first orbit transfer maneuver, such that the satellite will reenter Earth's atmosphere no longer than 25 years after completion of the first orbit transfer maneuver.

12. The earth-orbiting satellite of claim 11 , wherein the first orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by more than 60 m/sec.

13. The earth-orbiting satellite of claim 11 , wherein the orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by approximately 65 m/sec.

14. The earth-orbiting satellite of claim 11 , wherein the first interim orbit has an initial second apogee altitude that is approximately 5000 km higher than the operational orbit apogee altitude.

15. The earth-orbiting satellite of claim 11 , wherein the right ascension of ascending node is 0+/−20 degrees.

16. The earth-orbiting satellite of claim 11 , wherein the operational orbit has an orbital period of approximately 23.93 hours.

17. The earth-orbiting satellite of claim 11 , wherein the first orbit transfer maneuver includes at least one firing of a chemical or electric thruster proximate to orbit perigee.

18. The earth-orbiting satellite of claim 17 , wherein the first orbit transfer maneuver includes a plurality of thruster firings.

19. The earth-orbiting satellite of claim 11 , wherein the spacecraft controller is further configured to:

execute, following a period of time in which the first interim orbit is allowed to decay, a second orbit transfer maneuver that transfers the satellite from the decayed first interim orbit to a second interim orbit, the decayed first interim orbit having an ascending node radius less than 42,160 km and the second interim orbit has an ascending node radius greater than 42,170 km.

20. The earth-orbiting satellite of claim 19 , wherein the second orbit transfer maneuver includes increasing the satellite velocity, proximate to orbit perigee, by approximately 7 m/sec.