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Index Page oscelt_c
Table of contents
Procedure Abstract Required_Reading Keywords Brief_I/O Detailed_Input | Detailed_Output Parameters Exceptions Files Particulars Examples | Restrictions Literature_References Author_and_Institution Version Index_Entries
#### Procedure
       oscelt_c ( Determine conic elements from state )

       void oscelt_c ( ConstSpiceDouble   state[6],
                       SpiceDouble        et,
                       SpiceDouble        mu,
                       SpiceDouble        elts[8]   )
#### Abstract
       Determine the set of osculating conic orbital elements that
       corresponds to the state (position, velocity) of a body at
       some epoch.
#### Required_Reading
       None.
#### Keywords
       CONIC
       EPHEMERIS
#### Brief_I/O
       VARIABLE  I/O  DESCRIPTION
       --------  ---  --------------------------------------------------
       state      I   State of body at epoch of elements.
       et         I   Epoch of elements.
       mu         I   Gravitational parameter (GM) of primary body.
       elts       O   Equivalent conic elements.
#### Detailed_Input
       state       is the state (position and velocity) of the body
                   at some epoch. Components are x, y, z, dx/dt, dy/dt,
                   dz/dt. `state' must be expressed relative to an
                   inertial reference frame. Units are km and km/sec.

       et          is the epoch of the input state, in ephemeris seconds
                   past J2000.

                                                        3    2
       mu          is the gravitational parameter (GM, km /sec ) of
                   the primary body.
#### Detailed_Output
       elts        are equivalent conic elements describing the orbit
                   of the body around its primary. The elements are,
                   in order:

                      rp      Perifocal distance.
                      ecc     Eccentricity.
                      inc     Inclination.
                      lnode   Longitude of the ascending node.
                      argp    Argument of periapsis.
                      m0      Mean anomaly at epoch.
                      t0      Epoch.
                      mu      Gravitational parameter.

                   The epoch of the elements is the epoch of the input
                   state. Units are km, rad, rad/sec. The same elements
                   are used to describe all three types (elliptic,
                   hyperbolic, and parabolic) of conic orbit.
#### Parameters
       None.
#### Exceptions
       1)  If `mu' is not positive, the error SPICE(NONPOSITIVEMASS)
           is signaled by a routine in the call tree of this routine.

       2)  If the specific angular momentum vector derived from `state'
           is the zero vector, the error SPICE(DEGENERATECASE)
           is signaled by a routine in the call tree of this routine.

       3)  If the position or velocity vectors derived from `state'
           is the zero vector, the error SPICE(DEGENERATECASE)
           is signaled by a routine in the call tree of this routine.

       4)  If the inclination is determined to be zero or 180 degrees,
           the longitude of the ascending node is set to zero.

       5)  If the eccentricity is determined to be zero, the argument of
           periapse is set to zero.

       6)  If the eccentricity of the orbit is very close to but not
           equal to zero, the argument of periapse may not be accurately
           determined.

       7)  For inclinations near but not equal to 0 or 180 degrees,
           the longitude of the ascending node may not be determined
           accurately. The argument of periapse and mean anomaly may
           also be inaccurate.

       8)  For eccentricities very close to but not equal to 1, the
           results of this routine are unreliable.

       9)  If the specific angular momentum vector is non-zero but
           "close" to zero, the results of this routine are unreliable.

       10) If `state' is expressed relative to a non-inertial reference
           frame, the resulting elements are invalid. No error checking
           is done to detect this problem.
#### Files
       None.
#### Particulars
       The CSPICE routine conics_c is the inverse of this routine:
       conics_c maps a set of osculating elements and a time to a state
       vector.
#### Examples
       The numerical results shown for these examples may differ across
       platforms. The results depend on the SPICE kernels used as
       input, the compiler and supporting libraries, and the machine
       specific arithmetic implementation.

       1) Determine the osculating elements of Phobos with respect to
          Mars at some arbitrary time in the J2000 inertial reference
          frame.

          Use the meta-kernel shown below to load the required SPICE
          kernels.

             KPL/MK

             File name: oscelt_ex1.tm

             This meta-kernel is intended to support operation of SPICE
             example programs. The kernels shown here should not be
             assumed to contain adequate or correct versions of data
             required by SPICE-based user applications.

             In order for an application to use this meta-kernel, the
             kernels referenced here must be present in the user's
             current working directory.

             The names and contents of the kernels referenced
             by this meta-kernel are as follows:

                File name                     Contents
                ---------                     --------
                mar097.bsp                    Mars satellite ephemeris
                gm_de431.tpc                  Gravitational constants
                naif0012.tls                  Leapseconds

             \begindata

                KERNELS_TO_LOAD = ( 'mar097.bsp',
                                    'gm_de431.tpc',
                                    'naif0012.tls'  )

             \begintext

             End of meta-kernel

          Example code begins here.

          /.
             Program oscelt_ex1
          ./
          #include <stdio.h>
          #include "SpiceUsr.h"

          int main( )
          {
             /.
             Local variables.
             ./
             SpiceDouble          elts   [8];
             SpiceDouble          et;
             SpiceDouble          lt;
             SpiceDouble          mu     [1];
             SpiceDouble          state  [6];

             SpiceInt             dim;

             /.
             Load the meta kernel listing the needed SPK, LSK and
             PCK with gravitational parameters kernels.
             ./
             furnsh_c ( "oscelt_ex1.tm" );

             /.
             Convert the time string to ephemeris time
             ./
             str2et_c( "Dec 25, 2007", &et );

             /.
             Retrieve the state of Phobos with respect to Mars in J2000.
             ./
             spkezr_c ( "PHOBOS", et, "J2000", "NONE", "MARS",
                        state,    &lt                         );

             /.
             Read the gravitational parameter for Mars.
             ./
             bodvrd_c ( "MARS", "GM", 1, &dim, mu );

             /.
             Convert the state 6-vector to the elts 8-vector. Note: the
             bodvrd_c returns data as arrays, so to access the gravitational
             parameter (the only value in the array), we use mu[0].
             ./
             oscelt_c ( state, et, mu[0], elts );

             /.
             Output the elts vector.
             ./
             printf( "Perifocal distance          (km): %21.10f\n", elts[0] );
             printf( "Eccentricity                    : %21.10f\n", elts[1] );
             printf( "Inclination                (deg): %21.10f\n",
                                                          elts[2] * dpr_c() );
             printf( "Lon of ascending node      (deg): %21.10f\n",
                                                          elts[3] * dpr_c() );
             printf( "Argument of periapsis      (deg): %21.10f\n",
                                                          elts[4] * dpr_c() );
             printf( "Mean anomaly at epoch      (deg): %21.10f\n",
                                                          elts[5] * dpr_c() );
             printf( "Epoch                        (s): %21.10f\n", elts[6] );
             printf( "Gravitational parameter (km3/s2): %21.10f\n", elts[7] );

             return ( 0 );
          }

          When this program was executed on a Mac/Intel/cc/64-bit
          platform, the output was:

          Perifocal distance          (km):       9232.5746716211
          Eccentricity                    :          0.0156113904
          Inclination                (deg):         38.1225231660
          Lon of ascending node      (deg):         47.0384055902
          Argument of periapsis      (deg):        214.1546430017
          Mean anomaly at epoch      (deg):        340.5048466068
          Epoch                        (s):  251812865.1837092042
          Gravitational parameter (km3/s2):      42828.3736206991

       2) Calculate the history of Phobos's orbit plane inclination
          with respect to Mars in the J2000 frame at intervals of six
          months for a time interval of 10 years.

          Use the meta-kernel from the first example.

          Example code begins here.

          /.
             Program oscelt_ex2
          ./
          #include <stdio.h>
          #include "SpiceUsr.h"

          int main( )
          {
             /.
             Local parameters.
             ./
             #define TIMLEN       41
             /.
             Local variables.
             ./
             SpiceChar            utcstr [TIMLEN];

             SpiceDouble          elts   [8];
             SpiceDouble          et;
             SpiceDouble          lt;
             SpiceDouble          mu     [1];
             SpiceDouble          state  [6];
             SpiceDouble          step;

             SpiceInt             dim;
             SpiceInt             i;

             /.
             Load the meta kernel listing the needed SPK, LSK and
             PCK with gravitational parameters kernels.
             ./
             furnsh_c ( "oscelt_ex1.tm" );

             /.
             Read the gravitational parameter for Mars.
             ./
             bodvrd_c ( "MARS", "GM", 1, &dim, mu );

             /.
             Convert the time string to ephemeris time
             ./
             str2et_c( "Jan 1, 2000 12:00:00", &et );

             /.
             A step of six months - in seconds.
             ./
             step = 180.0 * spd_c( );

             /.
             10 years in steps of six months starting
             approximately Jan 1, 2000.
             ./
             printf( "        UCT Time          Inclination\n" );
             printf( "------------------------  -----------\n" );

             for ( i = 0; i < 20; i++ )
             {

                /.
                Retrieve the state; convert to osculating elements.
                ./
                spkezr_c ( "PHOBOS", et, "J2000", "NONE", "MARS",
                           state,    &lt                         );
                oscelt_c ( state, et, mu[0], elts );

                /.
                Convert the ephemeris time to calendar UTC.
                ./
                et2utc_c( et, "C", 3, TIMLEN, utcstr );

                printf( "%s %12.6f\n", utcstr, elts[2] * dpr_c( ) );

                et += step;

             }

             return ( 0 );
          }

          When this program was executed on a Mac/Intel/cc/64-bit
          platform, the output was:

                  UCT Time          Inclination
          ------------------------  -----------
          2000 JAN 01 12:00:00.000    36.055248
          2000 JUN 29 12:00:00.000    37.112144
          2000 DEC 26 12:00:00.000    38.152129
          2001 JUN 24 12:00:00.000    37.552071
          2001 DEC 21 12:00:00.000    36.242049
          2002 JUN 19 11:59:59.999    36.330470
          2002 DEC 16 12:00:00.000    37.674595
          2003 JUN 14 11:59:59.999    38.121191
          2003 DEC 11 12:00:00.001    36.973204
          2004 JUN 08 11:59:59.999    36.033732
          2004 DEC 05 12:00:00.001    36.844542
          2005 JUN 03 11:59:59.999    38.077365
          2005 NOV 30 12:00:00.001    37.786106
          2006 MAY 29 11:59:58.999    36.413540
          2006 NOV 25 11:59:59.001    36.171050
          2007 MAY 24 11:59:58.999    37.448015
          2007 NOV 20 11:59:59.001    38.189118
          2008 MAY 18 11:59:58.999    37.223573
          2008 NOV 14 11:59:59.001    36.084745
          2009 MAY 13 11:59:57.999    36.608971
#### Restrictions
       1)  The input state vector must be expressed relative to an
           inertial reference frame.

       2)  Osculating elements are generally not useful for
           high-accuracy work.

       3)  Accurate osculating elements may be difficult to derive for
           near-circular or near-equatorial orbits. Osculating elements
           for such orbits should be used with caution.

       4)  Extracting osculating elements from a state vector is a
           mathematically simple but numerically challenging task. The
           mapping from a state vector to equivalent elements is
           undefined for certain state vectors, and the mapping is
           difficult to implement with finite precision arithmetic for
           states near the subsets of R6 where singularities occur.

           In general, the elements found by this routine can have
           two kinds of problems:

           -  The elements are not accurate but still represent
              the input state accurately. The can happen in
              cases where the inclination is near zero or 180
              degrees, or for near-circular orbits.

           -  The elements are garbage. This can occur when
              the eccentricity of the orbit is close to but
              not equal to 1. In general, any inputs that cause
              great loss of precision in the computation of the
              specific angular momentum vector or the eccentricity
              vector will result in invalid outputs.

           For further details, see the -Exceptions section.

           Users of this routine should carefully consider whether
           it is suitable for their applications. One recommended
           "sanity check" on the outputs is to supply them to the
           CSPICE routine conics_c and compare the resulting state
           vector with the one supplied to this routine.
#### Literature_References
       [1]  R. Bate, D. Mueller, and J. White, "Fundamentals of
            Astrodynamics," Dover Publications Inc., 1971.
#### Author_and_Institution
       N.J. Bachman        (JPL)
       J. Diaz del Rio     (ODC Space)
       E.D. Wright         (JPL)
#### Version
       -CSPICE Version 1.0.3, 10-AUG-2021 (JDR)

           Edited the header to comply with NAIF standard. Added complete code
           examples to -Examples section.

       -CSPICE Version 1.0.2, 27-DEC-2007 (NJB)

          Updated -Index_Entries header section to use keywords
          "osculating" and "convert." Updated -Particulars header
          section to refer to conics_c.

       -CSPICE Version 1.0.1, 17-NOV-2005 (NJB)

          The -Exceptions and -Restrictions header sections were filled in.
          Some corrections were made to the code example.

       -CSPICE Version 1.0.0, 16-APR-1999 (EDW)
#### Index_Entries
       conic elements from state
       osculating elements from state
       convert state to osculating elements
#### Link to routine oscelt_c source file oscelt_c.c

    Fri Dec 31 18:41:10 2021