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The elements of an orbit are the parameters needed to specify that orbit uniquely, given a model of two point masses obeying the Newtonian laws of motion and the inverse-square law of gravitational attraction. Because there are multiple ways of parameterising a motion, depending on which set of variables you choose to measure, there are several different ways of defining sets of orbital elements, sets each of which will specify the same orbit. This problem contains three degrees of freedom (the three Cartesian coordinates of the orbiting body). Therefore, each particular Keplerian ( = unperturbed) orbit is fully defined by six quantities - the initial values of the Cartesian components of the body's position and velocity. For this reason, all sets of orbital elements contain exactly six parameters. For a mathematically accurate explanation of this fact see the Discussion and references therein. (See also: orbital state vectors).
Keplerian elements The traditionally used set of orbital elements is called the set of Keplerian elements, after Johannes Kepler and his Kepler's laws. The Keplerian elements are six: Keplerian elements can be obtained from orbital state vectors using VEC2TLE software or by some direct computations. We see that the first three orbital elements are simply the Eulerian angles defining the orientation of the orbit relative to some fiducial coordinate system. The next two establish the shape of the orbit, while the last establishes the location of the orbiting body at a particular time. Altogether, the Keplerian elements parameterise a conic orbit emerging in an unperturbed two-body problem — an ellipse, a parabola, or a hyperbola. In a more realistic setting, a perturbed trajectory is represented as a sequence of such instantaneous conics that share one of their foci. In case the orbital elements are postulated to parameterise a sequence of conics that are always tangent to the trajectory, these orbital elements are called osculating. Alternative expressions Instead of the mean anomaly at epoch, , the mean anomaly , mean longitude, true anomaly or, rarely, the eccentric anomaly may also be used. Sometimes the epoch itself is used as the sixth orbital element. Other orbital parameters, such as the period, can then be calculated from the Keplerian elements. In some cases, the period is used as an orbital element instead of the semi-major axis. Visualizing an Orbit In Fig. 1, the orbital plane (yellow) intersects a reference plane called the plane of ecliptic (grey). The intersection is called the line of nodes, as it connects the center of mass with the ascending and descending nodes. This plane, together with the Vernal Point, (♈) establishes a reference frame. The elements can be seen as defining the orbit in this frame by degrees: Variance Among Keplerian Elements and Trajectories of Orbiting Bodies Because the simple Newtonian model of orbital motion of idealised points in free space is not exact, the orbital elements of real objects tend to change over time. Evolution of the orbital elements takes place due to the gravitational pull of bodies other than the primary, due to the nonsphericity of the primary, due to the atmospheric drag, relativistic effects, radiation pressure, electromagnetic forces, and so on. This evolution is described by the so-called planetary equations, which come in the form of Lagrange, or in the form of Gauss, or in the form of Delaunay, or in the form of Poincaré, or in the form of Hill. (The latter is a very exotic option, emerging in the case when the true anomaly enters the set of six orbital elements. Hill considered this kind of orbit parameterisation back in 1913.) For more information, see the Discussion. Two line elements Keplerian elements parameters can be encoded as text in a number of formats. The most common of them is the NASA/NORAD "two-line elements"(TLE) format* , originally designed for use with 80-column punched cards, but still in use because it is the most common format, and works as well as any other. Depending on the application and object orbit, the data derived from TLEs older than 30 days can become unreliable. Orbital positions can be calculated from TLEs through the SGP/SGP4/SDP4/SGP8/SDP8 algorithms.
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