Routing the asteroid surface vehicle with detailed mechanics

The motion of a surface vehicle on/above an irregular object is inv.estigated for a potential interest in the insitu explorations to asteroids of the solar system. A global valid numeric method, including detailed gravity and geo- morphology, is developed to mimic the behaviors of the test particles governed by the orbital equations and surface cou- pling effects. A general discussion on the surface mechanical environment of a specified asteroid, 1620 Geographos, is presented to make a global evaluation of the surface vehicle＇s working conditions. We show the connections between the natural trajectories near the ground and differential features of the asteroid surface, which describes both the good and bad of typical terrains from the viewpoint of vehicles＇ dynamic performances. Monte Carlo simulations are performed to take a further look at the trajectories of particles initializing near the surface. The simulations reveal consistent conclusions with the analysis, i.e., the open- field flat ground and slightly concave basins/valleys are the best choices for the vehicles＇ dynamical security. The dependence of decending trajectories on the releasing height is studied as an application; the results show that the pole direction （where the centrifugal force is zero） is the most stable direction in which the shift of a natural trajectory will be well limited after landing. We present this work as an example for pre-analysis that provides guidance to engineering design of the exploration site and routing the surface vehicles.

Ivar asteroid rendezvous mission system scenario and trajectory design

The asteroid exploration opportunities are searched and calculated with launch dates in 2006 to2010, and with asteroid Ivar 1627 as the target, the spacecraft and its subsystems are designed and analyzed,and the transfer trajectory is designed using △VEGA technology for the asteroid rendezvous. The design resultssatisfied the energy requirements for small explorers.

Design of low-energy transfer from lunar orbit to asteroid in the Sun-Earth-Moon system

Asteroid exploration trajectories which start from a lunar orbit are investigated in this work.It is assumed that the probe departs from lunar orbit and returns to the vicinity of Earth,then escapes from the Earth by performing a perigee maneuver.A low-energy transfer in Sun-EarthMoon system is adopted.First,the feasible region of lowenergy transfer from lunar orbit to perigee within 5 000 km height above the Earth surface in Sun-Earth-Moon system is calculated and analyzed.Three transfer types are found,i.e.,large maneuver and fast transfers,small maneuver and fast transfers,and disordered and slow transfers.Most of feasibility trajectories belong to the first two types.Then,the lowenergy trajectory leg from lunar orbit to perigee and a heliocentric trajectory leg from perigee to asteroid are patched by a perigee maneuver.The optimal full-transfer trajectory is obtained by exploiting the differential evolution algorithm.Finally,taking 4179 Toutatis asteroid as the target,some low-energy transfer trajectories are obtained and analyzed.

Autonomous navigation of an asteroid orbiter enhanced by a beacon satellite in a high-altitude orbit

This study aims to assess the autonomous navigation performance of an asteroid orbiter enhanced using an inter-satellite link to a beacon satellite.Autonomous navigation includes the orbit determination of the orbiter and beacon,and asteroid gravity estimation without any ground station support.Navigation measurements were acquired using satellite-to-satellite tracking(SST)and optical observation of asteroid surface landmarks.This study presents a new orbiter-beacon SST scheme,in which the orbiter circumnavigates the asteroid in a low-altitude strongly-perturbed orbit,and the beacon remains in a high-altitude weakly-perturbed orbit.We used Asteroid 433 Eros as an example,and analyzed and designed low-and high-altitude orbits for the orbiter and beacon.The navigation measurements were precisely modeled,extended Kalman filters were devised,and observation configuration was analyzed using the Cramer-Rao lower bound(CRLB).Monte Carlo simulations were carried out to assess the effects of the orbital inclination and altitudes of the orbiter and beacon as key influencing factors.The simulation results showed that the proposed SST scheme was an effective solution for enhancing the autonomous navigation performance of the orbiter,particularly for improving the accuracy of gravity estimation.

Simulation and analysis of asteroid force closure anchoring performance based on discrete element method

Asteroid exploration is significant for studying the origin of the solar system,establishing planetary defenses,and alleviating the resource crisis of the Earth.Asteroid anchoring is the basis of in-situ exploration and resource development and utilization.Therefore,the performance of asteroid force-closure anchoring is investigated using the discrete element method.The micro parameters of the simulated materials are calibrated with angle of repose and uniaxial compression experiments,based on which the regional modeling method is adopted to establish the anchoring discrete element model.Asteroid anchoring experiments are conducted on a self-developed microgravity simulation platform to verify the accuracy of the simulation model.The asteroid anchoring simulations are performed to investigate the influence of external force on the anchoring performance.The analysis of anchoring force varying with time and the interaction between the anchor and regolith particles reveals the influence mechanism of external force direction on the anchoring performance.The external force direction affects the critical anchoring force by influencing the failure of the force-closure structure.The comprehensive analysis of simulation results clarifies the variation of the critical anchoring force with the external forces.Finally,a stable anchoring region is established,beneficial for asteroid anchoring device design.

Design and optimization of a trajectory for Moon departure Near Earth Asteroid exploration

The lunar probe often has some remaining fuel on completing the predefined Moon exploration mission and may carry out some additional tasks from the Moon orbit using the fuel.The possibility for the lunar probe to escape from the Moon and the Earth is analyzed.Design and optimization of the trajectory from the Moon orbit to the Near Earth Asteroids (NEAs) using the spacecraft's residual fuel is studied.At first,the semi-major axis,inclinations and the phase relations with the Earth of all the numbered NEAs are investigated to preliminarily select the possible targets.Based on the Sun-centered two-body problem,the launch window and the asteroid candidates are determined by calculating the minimum delta-v for two-impulse rendezvous mission and one-impulse flyby mission,respectively.For a precise designed trajectory,a full ephemeris dynamical model,which includes gravities of the Sun,the planets and the Moon,is adopted by reading the JPL ephemeris.The departure time,arrival time,burning time duration and thrust angles are set as variables to be designed and optimized.The optimization problem is solved via the Particle Swarm Optimization (PSO) algorithm.Moreover,two feasible NEA flyby missions are presented.

SmallSat swarm gravimetry:Revealing the interior structure of asteroids and comets

A growing interest in small body exploration has motivated research into the rapid char-acterization of near-Earth objects to meet economic or scientific objectives.Specifically,knowledge of the internal density structure can aid with target selection and enables an understanding of prehistoric planetary formation to be developed.To this end,multi-layer extensions to the polyhedral gravity model are suggested,and an inversion technique is implemented to present their effectiveness.On-orbit gravity gradiometry is simulated and employed in stochastic and deterministic algorithms,with results that imply robustness in both cases.