High order symplectic integrators for perturbed Hamiltonian systems

A family of symplectic integrators adapted for the integration of perturbed Hamiltonian systems of the form H=A+B was given in (McLachlan, 1995). We give here a constructive proof that for all integer p, such integrator exists, with only positive steps, and with a remainder of order O(^p + ²²), where is the stepsize of the integrator. Moreover, we compute the analytical expressions of the leading terms of the remainders at all orders. We show also that for a large class of systems, a corrector step can be performed such that the remainder becomes O(^p +⁴²). The performances of these integrators are compared for the simple pendulum and the planetary three-body problem of Sun–Jupiter–Saturn.     

High precision symplectic integrators for the Solar System

Using a Newtonian model of the Solar System with all 8 planets, we perform extensive tests on various symplectic integrators of high orders, searching for the best splitting scheme for long term studies in the Solar System. These comparisons are made in Jacobi and heliocentric coordinates and the implementation of the algorithms is fully detailed for practical use. We conclude that high order integrators should be privileged, with a preference for the new $(10,6,4)$ method of Blanes et al. (2013).     

Explicit adaptive symplectic integrators for solving Hamiltonian systems

We consider Sundman and Poincaré transformations for the long-time numerical integration of Hamiltonian systems whose evolution occurs at different time scales. The transformed systems are numerically integrated using explicit symplectic methods. The schemes we consider are explicit symplectic methods with adaptive time steps and they generalise other methods from the literature, while exhibiting a high performance. The Sundman transformation can also be used on non-Hamiltonian systems while the Poincaré transformation can be used, in some cases, with more efficient symplectic integrators. The performance of both transformations with different symplectic methods is analysed on several numerical examples.     

Recent progress in the theory and application of symplectic integrators

In this paper various aspect of symplectic integrators are reviewed. Symplectic integrators are numerical integration methods for Hamiltonian systems which are designed to conserve the symplectic structure exactly as the original flow. There are explicit symplectic schemes for systems of the formH=T(p)+V(q), and implicit schemes for general Hamiltonian systems. As a general property, symplectic integrators conserve the energy quite well and therefore an artificial damping (excitation) caused by the accumulation of the local truncation error cannot occur. Symplectic integrators have been applied to the Kepler problem, the motion of minor bodies in the solar system and the long-term evolution of outer planets.     

Numerical experiments on the efficiency of second-order mixed-variable symplectic integrators for N-body problems

We discuss the efficiency of the so-called mixed-variable symplectic integrators for N-body problems. By performing numerical experiments, we first show that the evolution of the mean error in action-like variables is strongly dependent on the initial configuration of the system. Then we study the effect of changing the stepsize when dealing with problems including close encounters between a particle and a planet. Considering a previous study of the slow encounter between comet P/Oterma and Jupiter, we show that the overall orbital patterns can be reproduced, but this depends on the chosen value of the maximum integration stepsize. Moreover the Jacobi constant in a restricted three-body problem is not conserved anymore when the stepsize is changed frequently: over a 10⁵ year time span, to keep a relative error in this integral of motion of the same order as that given by a Bulirsch-Stoer integrator requires a very small integration stepsize and much more computing time. However, an integration of a sample including 10⁴ particles close to Neptune shows that the distributions of the variation of the elements over one orbital period of the particles obtained by the Bulirsch-Stoer integrator and the symplectic integrator up to a certain integration stepsize are rather similar. Therefore, mixed-variable symplectic integrators are efficient either for N-body problems which do not include close encounters or for statistical investigations on a big sample of particles.     

Gas-surface interactions and satellite drag coefficients

Information on gas-surface interactions in orbit has accumulated during the past 35 years. The important role played by atomic oxygen adsorbed on satellite surfaces has been revealed by the analysis of data from orbiting mass spectrometers and pressure gauges. Data from satellites of special design have yielded information on the energy accommodation and angular distributions of molecules reemitted from satellite surfaces. Consequently, it is now possible to calculate satellite drag coefficients from basic physical principles, utilizing parameters of gas-surface interactions measured in orbit. The results of such calculations are given. They show the drag coefficients of four satellites of different compact shapes in low-earth orbit with perigee altitudes in the range from about 150 to 300 km, where energy accommodation coefficients and diffuse angular distributions have been measured. The calculations are based on Sentman's analysis of drag forces in free-molecular flow. His model incorporates the random thermal motion of the incident molecules, and assumes that all molecules are diffusely reemitted The uncertainty caused by the assumption of diffuse reemission is estimated by using Schamberg's model of gas-surface interaction, which can take into account a quasi-specular component of the reemission. Such a quasi-specular component is likely to become more important at higher altitudes as the amount of adsorbed atomic oxygen decreases. A method of deducing accommodation coefficients and angular distributions at higher altitudes by comparing the simultaneous orbital decay of satellites of different shapes at a number of altitudes is suggested. The purpose is to improve thermospheric measurements and models, which are significantly affected by the choice of drag coefficients.     

Relativistic effects for near-earth satellite orbit determination

The relativistic formulations for the equations which describe the motion of a near-Earth satellite are compared for two commonly used coordinate reference systems (RS). The discussion describes the transformation between the solar system barycentric RS and both the non-inertial and inertial geocentric RSs. A relativistic correction for the Earth's geopotential expressed in the solar system barycentric RS and the effect of geodesic precession on the satellite orbit in the geocentric RS are derived in detail. The effect of the definition of coordinate time on scale is also examined. A long-arc solution using 3 years of laser range measurements of the motion of the Lageos satellite is used to demonstrate that the effects of relativity formulated in the geocentric RS and in the solar system barycentric RS are equivalent to a high degree of accuracy.     

The optimization of satellite orbit for Space-VLBI observation

By sending one or more telescopes into space,Space-VLBI(SVLBI)is able to achieve even higher angular resolution and is therefore the trend of the VLBI technique.For the SVLBI program,the design of satellite orbits plays an important role for the success of planned observation.In this paper,we present our orbit optimization scheme,so as to facilitate the design of satellite orbits for SVLBI observation.To achieve that,we characterize the uv coverage with a measure index and minimize it by finding out the corresponding orbit configuration.In this way,the design of satellite orbit is converted to an optimization problem.We can prove that,with an appropriate global minimization method,the best orbit configuration can be found within the reasonable time.Besides that,we demonstrate that this scheme can be used for the scheduling of SVLBI observations.     

Modified multirevolution integration methods for satellite orbit computation

Multirevolution methods allow for the computation of satellite orbits in steps spanning many revolutions. The methods previously discussed in the literature are based on polynomial approximations, and as a result they will integrate exactly (excluding round-off errors) polynomial functions of a discrete independent variable. Modified methods are derived that will integrate exactly products of linear and periodic functions. Numerical examples are given that show that these new methods provide better accuracy for certain satellite problems. It is also shown that information obtained from an approximate analytical solution of the satellite equations of motion, may be used to increase the accuracy and/or efficiency of the multirevolution integration.     

Variations in air density,satellite drag coefficient and atmospheric rotation rate from analysis of the orbit of 1966-92D

Variations in air density, the satellite drag coefficient, and the atmospheric rotation rate at 60°S lat and 120–130 km height during the period September 1968–June 1969 have been determined from analysis of the high-eccentricity orbit of the 4th Molyniya 1 upper-stage rocket body, 1966-92D. The results show good correlation between density increases and strong geomagnetic activity, although solar flares of equal geomagnetic index value do not consistently produce density changes of equal magnitude. A 30 per cent semi-annual variation was observed, but there was no indication of the 50 per cent lower thermosphere seasonal-latitudinal variation that was predicted from the CIRA 1972 atmosphere. The satellite drag coefficient was observed to begin decreasing with height at an altitude where the molecular mean free path, λ, was twice the satellite's length. The coefficient decreased to a value approaching 1.0 as the satellite's perigee height fell below the altitude where λ was one-half the length. A mean atmospheric rotation rate of 1.1 ± 0.1 Earth rot/day was obtained for the last 20 days of decay. However, variations were observed with west-to-east wind speeds of ?100 m/sec measured for a local time of 13 hr.     

Optimal periodic relative orbit and rectilinear relative orbits with eccentric reference orbits

The problem of two-body linearized periodic relative orbits with eccentric reference orbits is studied in this paper. The periodic relative orbit in the target-orbital coordinate system can be used in fly-around and formation-flying orbit design. Based on the closed-form solutions to the Tschauner–Hempel equations, the initial condition for periodic relative orbits is obtained. Then the minimum-fuel periodic-orbit condition with a single impulse is analytically derived for given initial position and velocity vectors. When considering the initial coasting time, the impulse position of the global minimum-fuel periodic orbit is proved to be near to the perigee of the target and can be obtained by numerical optimization algorithms. Moreover, the condition for a special periodic orbit, i.e., the rectilinear relative orbit in the target-orbital frame, is obtained. Numerical simulations are used to demonstrate the efficacy of the method, and show the geometry of the periodic relative orbit and the rectilinear relative orbit.     

On error bounds and initialization in satellite orbit theories

The order of magnitude of the error is investigated for a first-order von Zeipel theory of satellite orbits in an axisymmetric force field, i.e., first-order long period and short-period effects are included along with second order secular rates. The treatment is valid for zero eccentricity and/or inclination. In the case where initial position and velocity vectors are known, the in-track position error over time intervals of order 1/J ₂ is kept at 0(J ₂ ²), like the other position errors and velocity errors, by calibration of the mean motion with the aid of the energy integral. The results are specifically applicable to accuracy comparisons of the Brouwer orbit prediction method with numerical integration. A modified calibration is presented for the general asymmetric force field which includes tesseral harmonics.     

The motion of axisymmetric satellite with drag and radiation pressure

The axisymmetric satellite problem including radiation pressure and drag is treated. The equations of motion of the satellite are derived. The energy-like and Laplace-like invariants of motion have been derived for a general drag force function of the polar angle, and the Laplace-like invariant is used to find the orbit equation in the case of a spherical satellite. Then using the small parameter, the orbit of the satellite is determined for an axisymmetric satellite.     

Resonant attitude instabilities for a symmetric satellite in a circular orbit

The roll-yaw attitude motion of a spinning symmetric satellite in a circular orbit is investigated with particular emphasis on the behavior near resonance. Resonance in circular orbit occurs if there is a low-order commensurability between the coupled roll-yaw attitude frequencies. For the so-called Delp region where the Hamiltonian describing the linearized attitude oscillations is not positive definite, there can exist, near resonance, a simultaneous growth or decay of the energy of the two normal modes. Two sections of the resonance line ₂=3 ₁ permitting the largest effects are determined and the equations of motion are integrated numerically as a check on the resonance theory. In particular, resonance-induced instabilities are confirmed.     

An analytical iterative algorithm for the prediction of special satellite orbit points with the Brouwer orbit theory

Employing a direct recursive algorithm in relation with analytical theories will yield a considerable saving in computer time, as opposed to simulating a point by point integration through repeated evaluations of the orbit theory. As a case in point, we shall compute the set of osculating orbiting elements corresponding to special events within the revolution of an artificial satellite.     

Gas drag in primordial circumplanetary envelopes: A mechanism for satellite capture

Some natural satellites may have been captured due to the gas drag they experienced in passing through primordial circumplanetary nebulas. This paper models such an encounter and derives the testable parameters from the known properties of current solar system objects and Bodenheimer's (1977, Icarus 31) model of the earliest phases of Jupiter's evolution. We propose that the clusters of prograde and retrograde irregular satellites of Jupiter originated when two parent bodies were decelerated and fragmented as they passed through an extended primordial Jovian nebula. Fragmentation occured because the gas dynamic pressure exceeded the parent bodies' strengths. These events must have occurred only shortly before the primordial nebula experienced hydrodynamical collapse so that subsequently the fragments underwent only limited orbital evolution. Because self-gravity exceeded the relative drag force, the fragments initially remained together, only to be dispersed at a later time by a collision with a stray body. Predictions of this hypothesis, such as orbital distance of the irregular satellites and size of the parent bodies, are found to be consistent with the observed properties of Jupiter's irregular satellites. In addition nebular drag at a later time may have caused the inner three Galilean satellites to undergo a modest amount of orbital evolution, accounting for their present orbital resonance. Gas drag capture of Saturn's Phoebe and Iapetus and Neptune's Nereid and Triton may also be possible. Reasonable differences in properties could explain why these satellites, in contrast to the Jovian ones, did not fracture upon capture. The current irregular satellites represent only a tiny fraction of the bodies captured by primordial nebulas. The dominant fraction would have spiraled into the center of the nebula as a result of continued gas drag and thus offer one source for the heavy element cores of the outer planets. If one is willing to postulate the presence of a massive gaseous nebula around primordial Mars, then gas drag capture could account for the origin of the Martian moons. We hypothesize that a single parent body was captured in a region of the nebula where the gas velocity approached the Keplerian value, that it fragmented upon deceleration into at least two bodies, Phobos and Deimos, and that continued nebular drag led to the low eccentricity and inclination that characterize the satellites' current orbits. Following the dissipation of this nebula, the more massive Phobos tidally evolved to its current position.     

Launch window study for the highly eccentric orbit satellite HEOS-1

A satellite with a high eccentricitye0.95 is strongly perturbed by the sun and the moon. This fact and mission constraints restrict considerably the possible launch times for such a satellte. The launch window calculations can be performed in two steps in order to save computing time. An approximate analytical solution provides a general survey of the launch opportunities. An accurate numerical approach is then necessary for the exact definition of the launch window. In the case of the orbit of HEOS-1 (satellite 68 10901), moreover the consideration of the injection errors has been of great importance.Presented at the Conference on Celestial Mechanics, Oberwolfach, Germany, August 17–23, 1969.