Examples of perturbations
Some of the variations in the orbital parameters caused by perturbations can be understood in simple terms. The lunar orbit is inclined to the ecliptic plane by about 5°, and the longitude of its ascending node on the ecliptic plane (Ω in ) is observed to regress (Ω decreasing) a complete revolution in 18.61 years. The Sun is the dominant cause of this regression of the lunar node. When the Moon is closer to the Sun than Earth, the Sun accelerates the Moon slightly more than it accelerates Earth. This difference in the accelerations is what perturbs the lunar motion around Earth. The Moon does not fly off in this situation, since the acceleration of the Moon toward Earth is much larger than the difference between the Sun’s accelerations of Earth and the Moon.
The Sun, of course, is always in the ecliptic plane, since its apparent path among the stars defines the plane. This means that the perturbing acceleration just defined will always be pointed slightly toward the ecliptic plane whenever the Moon is below or above this plane in its orbital motion about Earth. This tendency to pull the Moon toward the ecliptic plane means that the Moon will cross the plane on each half orbit at a longitude that is slightly smaller than the longitude at which it would have crossed if the Sun had not been there. Thus, the line of nodes will have regressed. The instantaneous rate at which the node regresses varies as the geometry changes during the Moon’s motion around Earth, and during the Earth-Moon system’s motion around the Sun, but there is always a net regression. Such a change that is always in the same direction as time increases is called a secular perturbation. Superposed on the secular perturbation of the longitude of the node are periodic perturbations (periodically changing their direction), which are revealed by the fact that the rate of secular regression of the node is not constant in time. The Sun causes a secular increase in the longitude of the lunar perigee (Ω + ω in ) of one complete revolution in 8.85 years, as well as periodic perturbations in the inclination, eccentricity, and mean motion.
For near-Earth artificial satellites, the deviation of Earth’s mass distribution from spherical symmetry is the dominant cause of the perturbations from pure elliptic motion. The most important deviation is the equatorial bulge of Earth due to its rotation. If, for example, Earth were a sphere with a ring of mass around its Equator, the ring would give to a satellite whose orbit is inclined to the Equator a component of acceleration toward the Equator plane whenever the satellite was above or below this plane. By an argument similar to that for the Moon acted on by the Sun, this acceleration would cause the line of nodes of a close satellite orbit to regress a little more than 5° per day.
As a final example, the distribution of continents and oceans and the varying mass densities in Earth’s mantle (the layer underlying the crust) lead to a slight deviation of Earth’s gravitational force field from axial symmetry. Usually this causes only short-period perturbations of low amplitude for near-Earth satellites. However, communications or weather satellites that are meant to maintain a fixed longitude over the Equator (i.e., geostationary satellites, which orbit synchronously with Earth’s rotation) are destabilized by this deviation except at two longitudes. If the axial asymmetry is represented by a slightly elliptical Equator, the difference between the major and minor axis of the ellipse is about 64 metres, with the major axis located about 35° W. A satellite at a position slightly ahead of the long axis of the elliptical Equator will experience a component of acceleration opposite its direction of orbital motion (as if a large mountain were pulling it back). This acceleration makes the satellite fall closer to Earth and increases its mean motion, causing it to drift further ahead of the axial bulge on the Equator. If the satellite is slightly behind the axial bulge, it experiences an acceleration in the direction of its motion. This makes the satellite move away from Earth with a decrease in its mean motion, so that it will drift further behind the axial bulge. The synchronous Earth satellites are thus repelled from the long axis of the equatorial ellipse and attracted to the short axis, and compensating accelerations, usually from onboard jets, are required to stabilize a satellite at any longitude other than the two corresponding to the ends of the short axis of the axial bulge. (The jets are actually required for any longitude, as they must also compensate for other perturbations such as radiation pressure.)
The three-body problem
The inclusion of solar perturbations of the motion of the Moon results in a “three-body problem” (Earth-Moon-Sun), which is the simplest complication of the completely solvable two-body problem discussed above. When Earth, the Moon, and the Sun are considered to be point masses, this particular three-body problem is called “the main problem of the lunar theory,” which has been studied extensively with a variety of methods beginning with Newton. Although the three-body problem has no complete analytic solution in closed form, various series solutions by successive approximations achieve such accuracy that complete theories of the lunar motion must include the effects of the nonspherical mass distributions of both Earth and the Moon as well as the effects of the planets if the precision of the predicted positions is to approach that of the observations. Most of the schemes for the main problem are partially numerical and therefore apply only to the lunar motion. An exception is the completely analytic work of the French astronomer Charles-Eugène Delaunay (1816–72), who exploited and developed the most elegant techniques of classical mechanics pioneered by his contemporary, the Irish astronomer and mathematician William R. Hamilton (1805–65). Delaunay could predict the position of the Moon to within its own diameter over a 20-year time span. Since his development was entirely analytic, the work was applicable to the motions of satellites about other planets where the series expansions converged much more rapidly than they did for the application to the lunar motion.
Delaunay’s work on the lunar theory demonstrates some of the influence that celestial mechanics has had on the development of the techniques of classical mechanics. This close link between the development of classical mechanics and its application to celestial mechanics was probably no better demonstrated than in the work of the French mathematician Henri Poincaré (1854–1912). Poincaré, along with other great mathematicians such as George D. Birkhoff (1884–1944), Aurel Wintner (1903–58), and Andrey N. Kolmogorov (1903–87), placed celestial mechanics on a more sound mathematical basis and was less concerned with quantitatively accurate prediction of celestial body motion. Poincaré demonstrated that the series solutions in use in celestial mechanics for so long generally did not converge but that they could give accurate descriptions of the motion for significant periods of time in truncated form. The elaborate theoretical developments in celestial and classical mechanics have received more attention recently with the realization that a large class of motions are of an irregular or chaotic nature and require fundamentally different approaches for their description.
The restricted three-body problem
The simplest form of the three-body problem is called the restricted three-body problem, in which a particle of infinitesimal mass moves in the gravitational field of two massive bodies orbiting according to the exact solution of the two-body problem. (The particle with infinitesimal mass, sometimes called a massless particle, does not perturb the motions of the two massive bodies.) There is an enormous literature devoted to this problem, including both analytic and numerical developments. The analytic work was devoted mostly to the circular, planar restricted three-body problem, where all particles are confined to a plane and the two finite masses are in circular orbits around their centre of mass (a point on the line between the two masses that is closer to the more massive). Numerical developments allowed consideration of the more general problem.
In the circular problem, the two finite masses are fixed in a coordinate system rotating at the orbital angular velocity, with the origin (axis of rotation) at the centre of mass of the two bodies. Lagrange showed that in this rotating frame there were five stationary points at which the massless particle would remain fixed if placed there. There are three such points lying on the line connecting the two finite masses: one between the masses and one outside each of the masses. The other two stationary points, called the triangular points, are located equidistant from the two finite masses at a distance equal to the finite mass separation. The two masses and the triangular stationary points are thus located at the vertices of equilateral triangles in the plane of the circular orbit.
There is a constant of the motion in the rotating frame that leads to an equation relating the velocity of the massless particle in this frame to its position. For given values of this constant it is possible to construct curves in the plane on which the velocity vanishes. If such a zero-velocity curve is closed, the particle cannot escape from the interior of the closed zero-velocity curve if placed there with the constant of the motion equal to the value used to construct the curve. These zero-velocity curves can be used to show that the three collinear stationary points are all unstable in the sense that, if the particle is placed at one of these points, the slightest perturbation will cause it to move far away. The triangular points are stable if the ratio of the finite masses is less than 0.04, and the particle would execute small oscillations around one of the triangular points if it were pushed slightly away. Since the mass ratio of Jupiter to the Sun is about 0.001, the stability criterion is satisfied, and Lagrange predicted the presence of the Trojan asteroids at the triangular points of the Sun-Jupiter system 134 years before they were observed. Of course, the stability of the triangular points must also depend on the perturbations by any other bodies. Such perturbations are sufficiently small not to destabilize the Trojan asteroids. Single Trojan-like bodies have also been found orbiting at leading and trailing triangular points in the orbits of Neptune and of Saturn’s satellite Tethys, at the leading triangular point in the orbit of another Saturnian satellite, Dione, and at the trailing point in the orbit of Mars.
