# Orbital state vectors | Wikipedia audio article

In astrodynamics and celestial dynamics, the
orbital state vectors (sometimes state vectors) of an orbit are
Cartesian vectors of position ( r {\displaystyle \mathbf {r} }
) and velocity ( v {\displaystyle \mathbf {v} }
) that together with their time (epoch) ( t {\displaystyle t}
) uniquely determine the trajectory of the orbiting body in space.==Frame of reference==
State vectors are defined with respect to some frame of reference, usually but not always
an inertial reference frame. One of the more popular reference frames for the state vectors
of bodies moving near Earth is the Earth-centered equatorial system defined as follows:
The origin is Earth’s center of mass; The Z axis is coincident with Earth’s rotational
axis, positive northward; The X/Y plane coincides with Earth’s equatorial
plane, with the +X axis pointing toward the vernal equinox and the Y axis completing a
right-handed set.This reference frame is not truly inertial because of the slow, 26,000
year precession of Earth’s axis, so the reference frames defined by Earth’s orientation at a
standard astronomical epoch such as B1950 or J2000 are also commonly used.Many other
reference frames can be used to meet various application requirements, including those
centered on the Sun or on other planets or moons, the one defined by the barycenter and
total angular momentum of the solar system, or even a spacecraft’s own orbital plane and
angular momentum.==Position and velocity vectors==
The position vector r {\displaystyle \mathbf {r} }
describes the position of the body in the chosen frame of reference, while the
velocity vector v {\displaystyle \mathbf {v} }
describes its velocity in the same frame at the same time. Together, these two vectors
and the time at which they are valid uniquely describe the body’s trajectory.
The body does not actually have to be in orbit for its state vectors to determine its trajectory;
it only has to move ballistically, i.e., solely under the effects of its own inertia and gravity.
For example, it could be a spacecraft or missile in a suborbital trajectory. If other forces
such as drag or thrust are significant, they must be added vectorially to those of gravity
when performing the integration to determine future position and velocity.
For any object moving through space, the velocity vector is tangent to the trajectory. If u ^ t {\displaystyle {\hat {\mathbf {u} }}_{t}}
is the unit vector tangent to the trajectory, then v=
v u ^ t {\displaystyle \mathbf {v}=v{\hat {\mathbf
{u} }}_{t}}===Derivation===
The velocity vector v {\displaystyle \mathbf {v} \,}
can be derived from position vector r {\displaystyle \mathbf {r} \,}
by differentiation with respect to time: v=d r d
t {\displaystyle \mathbf {v}={d\mathbf {r}
\over {dt}}} An object’s state vector can be used to compute
its classical or Keplerian orbital elements and vice versa. Each representation has its
advantages. The elements are more descriptive of the size, shape and orientation of an orbit,
and may be used to quickly and easily estimate the object’s state at any arbitrary time provided
its motion is accurately modeled by the two-body problem with only small perturbations.
On the other hand, the state vector is more directly useful in a numerical integration
that accounts for significant, arbitrary, time-varying forces such as drag, thrust and
gravitational perturbations from third bodies as well as the gravity of the primary body.
The state vectors ( r {\displaystyle \mathbf {r} }
and v {\displaystyle \mathbf {v} }
) can be easily used to compute the angular momentum vector as h=r × v {\displaystyle \mathbf {h}=\mathbf {r} \times
\mathbf {v} } .
Because even satellites in low Earth orbit experience significant perturbations (primarily
from Earth’s non-spherical shape), the Keplerian elements computed from the state vector at
any moment are only valid at that time. Such element sets are known as osculating elements
because they coincide with the actual orbit only at that moment.==See also==
ECEF Earth-centered inertial