초록 ▼

Methods are provided for accurate insertion of a constellation of satellites into a common orbit plane with a mean inclination .theta.. In an embodiment of the invention, an initial satellite is inserted on an initial insertion date and, for each of subsequent satellites, a respective orbit insertio

Methods are provided for accurate insertion of a constellation of satellites into a common orbit plane with a mean inclination .theta.. In an embodiment of the invention, an initial satellite is inserted on an initial insertion date and, for each of subsequent satellites, a respective orbit insertion date is selected. With reference to the latter date, the initial insertion RAAN is updated to find a current RAAN. This updating is accomplished with knowledge of a regression rate of the initial satellite's orbit normal about a Q vector. With reference to the Q vector's inertial position, an instantaneous inclination that corresponds to the mean inclination .theta. and the updated RAAN is then derived. Finally, the subsequent satellite is injected into an orbit having the instantaneous inclination and the current RAAN.

대표청구항 ▼

[ What is claimed is:] [1.] A method of accurately inserting a plurality of satellites into a common orbit plane that has a mean inclination .theta. wherein a Q vector repetitively traces a path with a period of .about.18.6 years and represents orbit perturbations effected by the earth, the sun and

[ What is claimed is:] [1.] A method of accurately inserting a plurality of satellites into a common orbit plane that has a mean inclination .theta. wherein a Q vector repetitively traces a path with a period of .about.18.6 years and represents orbit perturbations effected by the earth, the sun and the moon and the orbit of each of said satellites has an orbit normal that is tilted from said Q vector by a respective constant angle and regresses about said Q vector with a respective regression rate equal to the cosine of its constant angle times the magnitude of said Q vector, the method comprising the steps of:choosing a nominal inertial position Q.sub.nip of said Q vector;selecting, for a first one of said satellites, a first orbit insertion date D.sub.f and a first right ascension of the ascending node (RAAN) .OMEGA..sub.f ;with reference to said nominal inertial position Q.sub.nip, deriving a first instantaneous inclination I.sub.f of an orbit that has said mean inclination .theta. and said first RAAN .OMEGA..sub.f ;inserting, on said first orbit insertion date D.sub.f, said first satellite into a first orbit having said first instantaneous inclination I.sub.f and said first RAAN .OMEGA..sub.f ; andfor each subsequent one of said satellites:a) scheduling an orbit insertion date of said subsequent satellite that is delayed from said first orbit insertion date D.sub.f by an elapsed time T.sub.e ;b) updating said first RAAN .OMEGA..sub.f with the first orbit's respective regression rate and said elapsed time Te to realize a subsequent RAAN .OMEGA..sub.s ;c) with reference to said nominal inertial position Q.sub.nip, obtaining a subsequent instantaneous inclination I.sub.s of an orbit that has said mean inclination .theta. and said subsequent RAAN .OMEGA..sub.s ; andd) inserting said subsequent satellite into an orbit that has said subsequent instantaneous inclination I.sub.s and said subsequent RAAN .OMEGA..sub.s.