conics_c |

## Procedurevoid conics_c ( ConstSpiceDouble elts[8], SpiceDouble et, SpiceDouble state[6] ) ## AbstractDetermine the state (position, velocity) of an orbiting body from a set of elliptic, hyperbolic, or parabolic orbital elements. ## Required_ReadingNone. ## KeywordsCONIC EPHEMERIS ## 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 osculating 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 T0 is given in ephemeris seconds past J2000. T0 is the instant at which the state of the body is specified by the elements. 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 signaled. 2) If a non-positive periapse distance is supplied, the error SPICE(BADPERIAPSEVALUE) is signaled. 3) If a non-positive value for the attracting mass is supplied, the error SPICE(BADGM), is signaled. 4) Errors such as an out of bounds value for `et' are diagnosed by routines in the call tree of 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 oscelt_c ( 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, later epoch. The difference between this state and the state predicted by the nominal orbit at the same epoch can be computed as follows. ## RestrictionsNone. ## Literature_References[1] Roger Bate, Fundamentals of Astrodynamics, Dover, 1971. ## Author_and_InstitutionN.J. Bachman (JPL) I.M. Underwood (JPL) W.L. Taber (JPL) E.D. Wright (JPL) ## Version-CSPICE Version 1.1.1, 29-JUL-2003 (NJB) Various header corrections were made. -CSPICE Version 1.1.0, 24-JUL-2001 (NJB) Changed protoype: input elts is now type (ConstSpiceDouble *). Implemented interface macro for casting input array to const. -CSPICE Version 1.0.1, 08-FEB-1998 (EDW) Corrected and clarified header entries. -CSPICE Version 1.0.0, 10-NOV-1997 (EDW) ## Index_Entriesstate from conic elements |

Wed Apr 5 17:54:30 2017