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Дата изменения: Sat Dec 17 06:09:25 2005 Дата индексирования: Mon Apr 11 00:06:03 2016 Кодировка: Поисковые слова: п п п п п п п п п п п п п п п п п р п |
Determine the set of osculating conic orbital elements that corresponds to the state (position, velocity) of a body at some epoch.
None.
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
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.
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.
None
None
Let 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.
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.
conics_c ( elts, later, nominal );
vsubg_c ( nominal, state, 6, diff );
printf( "Perturbation in x, dx/dt = %e %e\n", diff[0], diff[3] );
printf( " y, dy/dt = %e %e\n", diff[1], diff[4] );
printf( " z, dz/dt = %e %e\n", diff[2], diff[5] );
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 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.
1) If `mu' is not positive, the error SPICE(NONPOSITIVEMASS)
is signaled.
2) If the specific angular momentum vector derived from STATE
is the zero vector, the error SPICE(DEGENERATECASE)
is signaled.
3) If the position or velocity vectors derived from STATE
is the zero vector, the error SPICE(DEGENERATECASE)
is signaled.
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 eccentricy 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.
None
N.J. Bachman (JPL) K.R. Gehringer (JPL) I.M. Underwood (JPL) E.D. Wright (JPL)
[1] Roger Bate, Fundamentals of Astrodynamics, Dover, 1971.
-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)
conic elements from state