초록 ▼

A method is disclosed for designing an orbit of a spacecraft which allows the spacecraft to take a small-radius halo orbit near a Lagrange point while avoiding the prohibited zone where the spacecraft may be shadowed or might be prevented from making communication. The method makes it possible to ha

A method is disclosed for designing an orbit of a spacecraft which allows the spacecraft to take a small-radius halo orbit near a Lagrange point while avoiding the prohibited zone where the spacecraft may be shadowed or might be prevented from making communication. The method makes it possible to have a closed orbit although being similar to a Lissajous orbit, under a restricted condition where a propulsion force magnitude applied to a spacecraft is fixed, and where it rotates at a constant angular velocity, based on the equation of motion close to a Lagrange point. The method also provide a theory for guiding/controlling the orbit of a spacecraft, that is, the embodiment of the above orbit design method.

대표청구항 ▼

What is claimed: 1. An orbit design method for placing a spacecraft in an orbit comprising: identifying a celestial system comprising a first body and a second body in a cosmic space where both bodies turn around a barycenter, the barycenter being a center of masses of the two bodies, both bodies i

What is claimed: 1. An orbit design method for placing a spacecraft in an orbit comprising: identifying a celestial system comprising a first body and a second body in a cosmic space where both bodies turn around a barycenter, the barycenter being a center of masses of the two bodies, both bodies in a circular orbit, and a spacecraft equipped with a propulsion mechanism residing on a first line connecting the two bodies, around an equilibrium point near the second body; and implementing a coordinate system with respect to the celestial system such that the first line connecting the first and second celestial bodies is made a fixed axis, a second axis perpendicular to the fixed axis is included on a revolution plane of the two bodies, and a second line perpendicular to the revolution plane is made a third axis, with a plane spanned by both the fixed and third axes being made a first plane, where the spacecraft is located close to a Lagrange point on the fixed axis, where a gravity from the two bodies is balanced with a centrifugal force resulting from a revolution motion, wherein the propulsion mechanism of the spacecraft is continuously activated during orbit such that a propulsion force vector having a specified fixed magnitude is allowed to rotate at a specified fixed angular velocity so as to make the orbit of the spacecraft projected onto a second plane depict a non-Keplerian and circular halo orbit in a form of a closed small-radius circuit maintaining a passive stability location characteristic associated with a Lissajous orbit, the second plane being a plane perpendicular to any arbitrary line on the first plane spanned by both the fixed axis and the third axis. 2. The orbit design method according to claim 1, wherein a control of the orbit of the spacecraft is performed such that a direction of the propulsive force vector applied to the spacecraft coincides with a radial direction of the circular orbit obtained by projection of an original orbit to the second plane perpendicular to the arbitrary line on the first plane, wherein the control is a maintenance free maneuver that is capable of correcting the trajectory continuously. 3. An orbit design method for placing a spacecraft in an orbit comprising: identifying a celestial system comprising a first body and a second body in a cosmic space where both bodies turn around a barycenter, the barycenter being a center of masses of the two bodies, both bodies in a circular orbit, and a spacecraft equipped with a propulsion mechanism residing close to on a first line connecting the two bodies, around an equilibrium point near the second body; and implementing a coordinate system with respect to the celestial system such that the first line connecting the first and second celestial bodies is made a fixed axis, a second axis and a first axis are included on a revolution plane of the two bodies, and a second line perpendicular to the revolution plane is made a third axis with a plane spanned by both the fixed and third axes being made a first plane, where the spacecraft is located on the fixed axis, around an equilibrium point near the second body, where a gravity from the two bodies is balanced with a centrifugal force resulting from a revolution motion, wherein the propulsion mechanism of the spacecraft is continuously activated during orbit such that a propulsion force vector having a specified fixed magnitude is allowed to rotate at a specified fixed angular velocity so as to make the orbit of the spacecraft projected onto a second plane depict a non-Keplerian and circular halo orbit in a form of a closed small-radius circuit maintaining a passive stability location characteristic associated with a Lissajous orbit, the second plane being a plane perpendicular to the fixed axis on the first plane. 4. The orbit design method according to claim 3, wherein a control of the orbit of the spacecraft is performed such that a direction of the propulsive force vector applied to the spacecraft coincides with a radial direction of the circular orbit obtained by projection of an original orbit to the second plane perpendicular to the fixed axis on the first plane, wherein the control is a maintenance free maneuver that is capable of correcting the trajectory continuously. 5. The orbit design method according to claim 3, wherein the spacecraft is a space facility staying close to a Lagrange point. 6. The orbit design method according to claim 1, wherein the spacecraft is an observatory satellite staying close to a Lagrange point and aiming at scanning an entire heavenly space. 7. The orbit design method according to claim 3, wherein the spacecraft is an observatory satellite staying close to a Lagrange point and aiming at scanning an entire heavenly space. 8. The orbit design method according to claim 1, wherein the first body is the Sun and the second body is the Earth. 9. The orbit design method according to claim 3, wherein the first body is the Sun and the second body is the Earth.