Rutherford’s extended formula and optimization of the gravity assists beam modelling in the Solar system

Rutherford’s formula for the scattering of charged α-particles in the Coulomb field can be easily generalized to the case of gravitational scattering. The extended Rutherford formula for the gravitational scattering is presented. One of the types of the gravitational scattering in the Solar system is the gravity assist maneuvers. In this paper, an effective gravitational scattering cross-section is introduced by analogy for them and the generalized Rutherford formula for gravitational scattering is presented out when performing gravity assists. Modern methods of the ballistic design of the interplanetary space flights using gravity assist maneuvers around planets [1-3] are associated with the need to calculate a lot of trajectories (i.e. of the phase beams). For their effective use it is necessary to study the structure of non-linear flyby gravitational scattering using the Rutherford’ formula and to construct the corresponding effective modelling using according regularized phase beams. It is shown that with using of such approach, it is possible to significantly increase the efficiency of the recurrent procedure for the gravity assists chains searching for ballistic scenarios of the modern interplanetary flights.

Gravity assists gravitational scattering and the perturbation rings in the Solar system

One of the types of gravitational scattering in the Solar system within the framework of the model of the restricted three-body problem (R3BP) is gravity assist maneuvers of the “particles of insignificant mass” [1] (spacecraft, asteroids, comets, etc.). For their description, a physical analogy with the beam scattering of charged α particles in a Coulomb field is useful. However, unlike the scattering of charged particles, there are external restrictions for the possibility of gravity assists executing related from the restricted size of planet’s sphere of influence. At the same time, internal restrictions for the gravity assists performance estimated by the effective radii of planets are known from the literature on R3BP [2] (gravitational capture by the planet, falling into it). They depend from the particle asymptotic velocity relative the planet. For obvious reasons, their influence cuts off the possibility of effective gravity assists performance [3]. In this work the generalized estimates of the sizes of the near-planetary regions (“perturbation rings”), falling into which is a necessary condition for the implementation of gravity assists, are presented. The detailed analysis shows that Neptune and Saturn have the characteristic “perturbation rings” of the largest sizes in the Solar system, and Jupiter occupies only the fourth place in this checklist.

Gravity assists maneuver in the problem of extension accessible landing areas on the Venus surface

This study discusses the usage of Venus gravity assist in order to choose and reaching any point on Venusian surface. The launch of a spacecraft to Venus during the launch windows of 2029 to 2031 is considered for this purpose. The constraints for the method are the re-entry angle and the maximum possible overload. The primary basis of the proposed strategy is to use the gravitational field of Venus to transfer the spacecraft to an orbit resonant to the Venusian one – with the aim of expanding accessible landing areas. Results of the current research show that this strategy provides an essential increase in accessible landing areas and, moreover, may provide an access to any point on the surface of Venus with a small increase in ∆V required for launch from the Earth and in the flight duration. The comparison with the landing without using gravity assist near planet is also given.

Design Of Lunar-Gravity-Assisted Escape Trajectories

Lunar gravity assist is a means to boost the energy and C3 of an escape trajectory. Trajectories with two lunar gravity assists are considered and analyzed. Two approaches are applied and tested for the design of missions aimed at Near-Earth asteroids. In the first method, indirect optimization of the heliocentric leg is combined to an approximate analytical treatment of the geocentric phase for short escape trajectories. In the second method, the results of pre-computed maps of escape C3 are employed for the design of longer Sun-perturbed escape sequences combined with direct optimization of the heliocentric leg. Features are compared and suggestions about a combined use of the approaches are presented. The techniques are efficiently applied to the design of a mission to a near-Earth asteroid.