Index of Functions: A B C D E F G H I J K L M N O P Q R S T U V W X
Index Page prop2b_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
       prop2b_c ( Propagate a two-body solution )

       void prop2b_c ( SpiceDouble         gm,
                       ConstSpiceDouble    pvinit[6],
                       SpiceDouble         dt,
                       SpiceDouble         pvprop[6] )
#### Abstract
       Compute the state of a massless body at time t_0 + dt by applying
       the two-body force model to a given central mass and a given body
       state at time t_0.
#### Required_Reading
       None.
#### Keywords
       CONIC
       EPHEMERIS
       UTILITY
#### Brief_I/O
       VARIABLE  I/O  DESCRIPTION
       --------  ---  --------------------------------------------------
       gm         I   Gravity of the central mass.
       pvinit     I   Initial state from which to propagate a state.
       dt         I   Time offset from initial state to propagate to.
       pvprop     O   The propagated state.
#### Detailed_Input
       gm          is the gravitational constant G times the mass M of the
                   central body.

       pvinit      is the state at some specified time relative to the
                   central mass. The mass of the object is assumed to
                   be negligible when compared to the central mass.

       dt          is an offset in time from the time of the initial
                   state to which the two-body state should be
                   propagated. (The units of time and distance must be
                   the same in `gm', `pvinit', and `dt').
#### Detailed_Output
       pvprop      is the two-body propagation of the initial state
                   `dt' units of time past the epoch of the initial state.
#### Parameters
       None.
#### Exceptions
       1)  If `gm' is not positive, the error SPICE(NONPOSITIVEMASS) is
           signaled by a routine in the call tree of this routine.

       2)  If the position of the initial state is the zero vector, the error
           SPICE(ZEROPOSITION) is signaled by a routine in the call tree of this
           routine.

       3)  If the velocity of the initial state is the zero vector, the error
           SPICE(ZEROVELOCITY) is signaled by a routine in the call tree of this
           routine.

       4)  If the cross product of the position and velocity of `pvinit'
           has squared length of zero, the error SPICE(NONCONICMOTION)
           is signaled by a routine in the call tree of this routine.

       5)  If `dt' is so large that there is a danger of floating point overflow
           during computation, the error SPICE(DTOUTOFRANGE) is signaled by a
           routine in the call tree of this routine and a message is generated
           describing the problem. The value of `dt' must be "reasonable". In
           other words, `dt' should be less than 10**20 seconds for realistic
           solar system orbits specified in the MKS system. (The actual bounds
           on `dt' are much greater but require substantial computation.) The
           "reasonableness" of `dt' is checked at run-time.
#### Files
       None.
#### Particulars
       This routine uses a universal variables formulation for the
       two-body motion of an object in orbit about a central mass. It
       propagates an initial state to an epoch offset from the
       epoch of the initial state by time `dt'.

       This routine does not suffer from the finite precision
       problems of the machine that are inherent to classical
       formulations based on the solutions to Kepler's equation:

             n( t - T ) = E - e * sin(E)         elliptic case
             n( t - T ) = e * sinh(F) - F        hyperbolic case

       The derivation used to determine the propagated state is a
       slight variation of the derivation in Danby's book
       "Fundamentals of Celestial Mechanics" [1].
#### 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) Use the two-body force model to propagate the state of a
          massless body orbiting the Earth at 100,000,000 km after half
          a period.

          In circular two-body motion, the orbital speed is

             s     = sqrt(mu/r)

          where mu is the central mass. After tau/2 = pi*r/s seconds
          (half period), the state should equal the negative of the
          original state.

          Example code begins here.

          /.
             Program prop2b_ex1
          ./
          #include <stdio.h>
          #include <math.h>
          #include "SpiceUsr.h"

          int main( )
          {
             /.
             Local variables.
             ./
             SpiceDouble            mu;
             SpiceDouble            pvinit [ 6 ];
             SpiceDouble            r;
             SpiceDouble            speed;
             SpiceDouble            state  [ 6 ];
             SpiceDouble            t;

             /.
             Initial values.
             ./
             mu    =  3.9860043543609598E+05;
             r     =  1.0e+08;
             speed =  sqrt( mu / r );
             t     =  pi_c()*r/speed;

             pvinit[0] =  0.0;
             pvinit[1] =  r/sqrt(2.0);
             pvinit[2] =  r/sqrt(2.0);
             pvinit[3] =  0.0;
             pvinit[4] = -speed/sqrt(2.0);
             pvinit[5] =  speed/sqrt(2.0);

             /.
             Calculate the state of the body at 0.5 period
             after the epoch.
             ./
             prop2b_c ( mu, pvinit, t, state );

             /.
             The `state' vector should equal -pvinit
             ./
             printf( "State at t0:\n" );
             printf( "   R   (km): %16.5f %16.5f %16.5f\n",
                            pvinit[0], pvinit[1], pvinit[2] );
             printf( "   V (km/s): %16.5f %16.5f %16.5f\n",
                            pvinit[3], pvinit[4], pvinit[5] );

             printf( "\n" );
             printf( "State at tau/2:\n" );
             printf( "   R   (km): %16.5f %16.5f %16.5f\n",
                            state[0], state[1], state[2]    );
             printf( "   V (km/s): %16.5f %16.5f %16.5f\n",
                            state[3], state[4], state[5]    );

             return ( 0 );
          }

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

          State at t0:
             R   (km):          0.00000   70710678.11865   70710678.11865
             V (km/s):          0.00000         -0.04464          0.04464

          State at tau/2:
             R   (km):         -0.00000  -70710678.11865  -70710678.11865
             V (km/s):          0.00000          0.04464         -0.04464

       2) When the eccentricity of an orbit is near 1, and the epoch
          of classical elements is near the epoch of periapse, classical
          formulations that propagate a state from elements tend to
          lack robustness due to the finite precision of floating point
          machines. In those situations it is better to use a universal
          variables formulation to propagate the state.

          By using this routine, you need not go from a state to
          elements and back to a state. Instead, you can get the state
          from an initial state.

          If `pvinit' is your initial state and you want the state 3600
          seconds later, the following call will suffice.

             Look up gm somewhere

             dt = 3600.0;

             prop2b_c ( gm, pvinit, dt, pvprop );

          After the call, `pvprop' will contain the state of the
          object 3600 seconds after the time it had state `pvinit'.
#### Restrictions
       1)  Users should be sure that `gm', `pvinit' and `dt' are all in the
           same system of units ( for example MKS ).
#### Literature_References
       [1]  J. Danby, "Fundamentals of Celestial Mechanics," 2nd Edition,
            pp 168-180, Willman-Bell, 1988.
#### Author_and_Institution
       N.J. Bachman        (JPL)
       J. Diaz del Rio     (ODC Space)
       W.L. Taber          (JPL)
       E.D. Wright         (JPL)
#### Version
       -CSPICE Version 1.1.1, 01-NOV-2021 (JDR)

           Edited -Examples section to comply with NAIF standard. Added
           complete code example.

       -CSPICE Version 1.1.0, 24-JUL-2001 (NJB)

           Changed prototype: input pvinit is now type
           (ConstSpiceDouble [6]). Implemented interface macro for
           casting input pvinit to const.

       -CSPICE Version 1.0.1, 20-MAR-1998 (EDW)

           Minor correction to header.

       -CSPICE Version 1.0.0, 08-FEB-1998 (EDW) (WLT)
#### Index_Entries
       Propagate state vector using two-body force model
#### Link to routine prop2b_c source file prop2b_c.c

    Fri Dec 31 18:41:10 2021