Patent EA-002665-B1 (EA)
Title: Satellite inclination changes using weak stability boundaries
Inventor: BELBRUNO EDWARD A
IPC: B64G1/24
Filed: 1998-04-24
Abstract: 1. A method of changing the at least one of an inclination and an altitude of a space vehicle in the earth's terrestial attraction of at least two heavenly bodies comprising following independent steps of: (a) traveling the space vehicle using a first transfer from a first heavenly body or a first heavenly body orbit to a weak capture in a weak stability boundary (WSB) or WSB orbit associated with said first and said secondary heavenly bodies; (b) performing, optionally, at least negligible maneuver at the WSB or the WSB orbit and ejecting therefrom, the negligible maneuver designed to be used in the changing of inclination and the altitude of the heavenly body and ejecting therefrom at the WSB orbit; and  (c) traveling using a second transfer from the WSB or rejecting therefrom the WSB orbit to at least a second heavenly body or to a second heavenly body orbit or to another orbit to the first heavenly body. 2. The method according to claim 1, wherein the first transfer is designed for using at least an operational ballistic capture transfer (BCT) and a Hohmann transfer. 3. The method according to claim 1 or 2, wherein the second transfer is designed for using at least an operational ballistic ejection transfer (BET) and a Hohmann transfer. 4. The method according to claims 1 to 3, wherein the first transfer is designed for using the operational ballistic capture transfer (BCT), wherein step (a) further comprises a step of travelling using (BCT) from the first heavenly body or from the first heavenly body orbit orbit to a weak capture in a weak stability boundary (WSB) or WSB orbit associated with said first heavenly body by implementing the forward targeting process by varying for convergence of the target variables at the (WSB) or (WSB) orbit. 5. The method according to claims 1 to 3, wherein the first transferis designed for the operational ballistic capture transfer (BCT), wherein step (a) further comprises a step of travelling using (BCT) from the first heavenly body or from the first heavenly body orbit orbit to a weak capture in a weak stability boundary (WSB) or WSB orbit associated with said first heavenly body by implementing the forward targeting process by varying at least two spherical parameters varying for convergence of the target variables at the (WSB) or (WSB) orbit, while maintaining at least one classical variable used in said forward targeting process substantially fixed. 6. The method according to claims 1 to 3, wherein the first transferis designed for the operational ballistic capture transfer (BCT), wherein step (a) further comprises a step of travelling using (BCT) from the first heavenly body or from the first heavenly body orbit orbit to a weak capture in a weak stability boundary (WSB) or WSB orbit associated with said first heavenly body by implementing the forward targeting process by varying velocity magnitude VE , and flight path angle γE for convergence of the target variables at the second heavenly object or the second heavenly object orbit, the target variables including radial distance, rM, and inclination iM. 7. The method according to claims 1 to 3, wherein the first transferis designed for the operational ballistic capture transfer (BCT), wherein step (a) further comprises a step of travelling using (BCT) from the first heavenly body or from the first heavenly body orbit orbit to a weak capture in a weak stability boundary (WSB) or WSB orbit associated with said first heavenly body by implementing the forward targeting process comprising a second order Newton algorithm, and wherein the second order Newton algorithm utilizes two control variables including velocity magnitude VE and flight path angle γE that are varied to achieve capture conditions at the second heavenly body or the second heavenly body orbit using two target variables including radial distance, rM, and inclination iM. 8. The method according to claim 7, wherein the velocity magnitude VE, and the flight path angle γE are decoupled from the second heavenly body or the second heavenly body orbit in the first transfer. 9. The method according to claim 7, wherein the velocity magnitude VE, and the flight path angle γE are decoupled from angular elements of the first heavenly body including inclination ie, ascending node relative to earth ΩE and argument of periapsis relative to the fast heavenly body ωE. 10. The method according to claims 1 to 9, wherein wherein said implementing step (b) further comprising a negligible maneuver of between 2-20 meters per second at a weak stability boundary (WSB) or WSB orbit for leaving said boundary or for rejecting therefrom. 11. The method according to claims 1 to 10, wherein said travelling step (a) further comprises the step of maneuvering around the second heavenly body or the second heavenly body orbit by performing a negligible maneuver of between 2-20 meters per second at a weak stability boundary (WSB) or WSB orbit. 12. The method according to claims 1 to 11, wherein the WSB or the WSB orbit is nonlinear and being substantially at a boundary of capture and escape, thereby allowing the capture and the escape to occur for a substantially zero or relatively small maneuver, and wherein solar gravitational perturbations influence the first and second transfers. 13. The method according to claims 1 to 12, wherein the WSB or the WSB orbit is substantially at a boundary of interaction between gravitational fields of the first heavenly body and the second heavenly body. 14. The method according to claims 1 to 13, wherein as at least one of a space vehicle, satellite and rocket moves in at least one of the WSB or the WSB orbit, a Kepler energy of the at least one of a space vehicle, satellite and rocket is slightly negative and substantially near to zero. 15. The method according to claims 1 to 14, wherein the at least one of the WSB or the WSB orbit is realizable at the predetermined arbitrary altitude by specifying a predetermined velocity magnitude of the at least one of a space vehicle, satellite and rocket, thereby defining a predetermined capture eccentricity. 16. The method according to claims 1 to 15, wherein said implementing step (a) further comprises the step of generating a trajectory around the second heavenly body or the second heavenly body orbit comprising at least a negligible maneuver of between 2-20 meters per second at the WSB or the WSB orbit for at least one of timing and positioning of at least one of a space vehicle, satellite and rocket, prior to ejection therefrom. 17. The method according to claims 1 to 16, wherein the second transfer is designed for using the operational ballistic ejection transfer (BET), wherein step (c) further comprises the step of travelling using ballistic ejection transfer (BET) from a weak capture in a weak stability boundary (WSB) or WSB orbit, with at least only to the first heavenly body or the first heavenly body orbit or to the second heavenly body or to second heavenly body orbit by implementing the forward targeting process by varying the second parameters of the method for convergence of the second predetermined parameters at the (WSB) or (WSB) orbit. 18. The method according to claim 17, wherein the second variables are decoupled from at least the first heavenly body or from the first heavenly body orbit said and from angular elements of the second heavenly body. 19. The method according to claims 1 to 18, wherein the first heavenly body or the first heavenly body orbit comprises earth or earth orbit, and wherein the second heavenly body or the second heavenly body orbit comprises moon or moon orbit. 20. The method according to claim 1, wherein implementing transfer for an space emanating substantially at a first heavenly body first heavenly body to arrive at a second heavenly body or second heavenly body orbit, comprising the independent steps of: (a) entering parameters for said method of generating operational ballistic capture transfer; (b) implementing a forward targeting process by varying the parameters for convergence of target variables at the second heavenly body or the second heavenly body orbit from the first heavenly body or the first heavenly body orbit; and (c) iterating step (b) until sufficient convergence to obtain me operational ballistic capture transfer from the first heavenly body or the first heavenly body orbit to the second heavenly body or the second heavenly body orbit. 21. A method according to claim 20, wherein said iterating step (c) further comprises the step of iterating step (b) until sufficient convergence to obtain the operational ballistic capture transfer from the first heavenly body or the first heavenly body orbit to the second heavenly body or the second heavenly body orbit via a weak stability boundary (WSB) orbit interposed therebetween. 22. A method according to claims 20 or 21, wherein said implementing step (b) further comprises the step of implementing the forward targeting process by varying at least two spherical parameters for convergence of the target variables at me second heavenly body or the second heavenly body orbit, while maintaining at least one classical variable used in said forward targeting process substantially fixed. 23. A method according to claims 20 or 21, wherein said implementing step (b) further comprises the step of implementing the forward targeting process by varying velocity magnitude VE, and flight path angle γE for convergence of the target variables at the second heavenly body or the second heavenly body orbit, the target variables including radial distance, rM, and inclination iM. 24. A method according to claims 20 or 21, wherein said implementing step (b) further comprises the step of implementing the forward targeting process comprising a second order Newton algorithm, and wherein the second order Newton algorithm utilizes two control variables including velocity magnitude VE and flight path ang