conics |

## ProcedureCONICS ( Determine state from conic elements ) SUBROUTINE CONICS ( ELTS, ET, STATE ) ## AbstractDetermine the state (position, velocity) of an orbiting body from a set of elliptic, hyperbolic, or parabolic orbital elements. ## Required_ReadingNone. ## KeywordsCONIC EPHEMERIS ## DeclarationsDOUBLE PRECISION ELTS ( 8 ) DOUBLE PRECISION ET DOUBLE PRECISION STATE ( 6 ) ## Brief_I/OVARIABLE I/O DESCRIPTION -------- --- -------------------------------------------------- ELTS I Conic elements. ET I Input time. STATE O State of orbiting body at ET. ## Detailed_InputELTS are conic elements describing the orbit of a body around a primary. The elements are, in order: RP Perifocal distance. ECC Eccentricity. INC Inclination. LNODE Longitude of the ascending node. ARGP Argument of periapse. M0 Mean anomaly at epoch. T0 Epoch. MU Gravitational parameter. Units are km, rad, rad/sec, km**3/sec**2. The epoch is given in ephemeris seconds past J2000. The same elements are used to describe all three types (elliptic, hyperbolic, and parabolic) of conic orbit. ET is the time at which the state of the orbiting body is to be determined, in ephemeris seconds J2000. ## Detailed_OutputSTATE is the state (position and velocity) of the body at time ET. Components are x, y, z, dx/dt, dy/dt, dz/dt. ## ParametersNone. ## Exceptions1) If the eccentricity supplied is less than 0, the error 'SPICE(BADECCENTRICITY)' is signalled. 2) If a non-positive periapse distance is supplied, the error 'SPICE(BADPERIAPSEVALUE)' is signalled. 3) If a non-positive value for the attracting mass is supplied, the error 'SPICE(BADGM)', is signalled. 4) Errors such as an out of bounds value for ET are diagnosed by routines called by this routine. ## FilesNone. ## ParticularsNone. ## ExamplesLet VINIT contain the initial state of a spacecraft relative to the center of a planet at epoch ET, and let GM be the gravitation parameter of the planet. The call CALL OSCELT ( VINIT, ET, GM, ELTS ) produces a set of osculating elements describing the nominal orbit that the spacecraft would follow in the absence of all other bodies in the solar system and non-gravitational forces on the spacecraft. Now let STATE contain the state of the same spacecraft at some other epoch, LATER. The difference between this state and the state predicted by the nominal orbit at the same epoch can be computed as follows. CALL ## RestrictionsNone. ## Literature_References[1] Roger Bate, Fundamentals of Astrodynamics, Dover, 1971. ## Author_and_InstitutionI.M. Underwood (JPL) W.L. Taber (JPL) ## VersionSPICELIB Version 4.0.0, 26-MAR-1998 (WLT) There was a coding error in the computation of the mean anomaly in the parabolic case. This problem has been corrected. SPICELIB Version 3.0.1, 15-OCT-1996 (WLT) Corrected a typo in the description of the units associated with the input elements. SPICELIB Version 3.0.0, 12-NOV-1992 (WLT) The routine was re-written to make use of NAIF's universal variables formulation for state propagation (PROP2B). As a result, several problems were simultaneously corrected. A major bug was fixed that caused improper state evaluations for ET's that precede the epoch of the elements in the elliptic case. A danger of non-convergence in the solution of Kepler's equation has been eliminated. In addition to this reformulation of |

Wed Apr 5 17:46:10 2017