Dynamics And Application Of Spacecraft Formation Flying In Eccentric Orbits

With the development of formation flying technology and the expansion of its orbital applications, spacecraft formations in eccentric orbits attract many aerospace specialists'attention gradually. In comparison with circular orbits, it is a challenge for us to study the general relative motion of two spacecrafts along elliptic orbits. The relative dynamics, perturbed relative motion analysis, configuration design and applications are discussed respectively, involved in formation flying on eccentric orbits.Firstly, for closely and loosely spaced vehicles, two approximate models expressed in orbital elements are established to describe the relative motion using an elliptic reference orbit respectively. The analysis of their modeling errors shows that if a relative errorε≤10-2is needed, the former fits a formation with a small interval d≤100km and the latter fits that with a large interval d≤1000km. And the dynamic model for loosely spaced vehicle formations extends the extant research and may be worthy of such missions as space exploration, electronic reconnaissance, distributed navigation.Secondly, under J2 perturbation, two kinds of stable formation designs described in mean and osculating orbital elements respectively are presented and the stability of formations on elliptic orbits is discussed deeply. In two mean orbital element designs, a stable relative motion in orbital plane can be realized, but the cross-track motion is stable only in special instances. Numerical simulation results show that five kinds of basic perturbed relative motions are sensitive to the chief spacecraft's osculating inclination, argument of perigee and mean anomaly. And these three osculating orbital elements may lead to minimum drifts of the relative motion at some special values. Moreover, the existence of these minimum drifts is validated using mean orbital element theory. When the real relative motion is the coupling of five basic formations, the (Δa ,Δe,Δi ,Δω,ΔM) or ( )ic ,ωc ,Mc0 of the corresponding stable formation can be obtained using mean orbital element and numerical methods. On the basis of these, three stable formation designs expressed in osculating elements are brought forward and can acquire stable relative motion in orbital plane, but may not in cross-track direction.Thirdly, according to the relative motion model for closely spaced vehicles, five characteristics of the relative trajectory between two spacecrafts on eccentric orbit are summarized and nine applicable configurations are proposed for one/ two/ multiple deputy spacecrafts. The configurations with only one deputy vehicle can be used as a component unit of a combined configuration, those with two vehicles can be applied in on-orbit or ground object's observation and measurement, and those with multiple vehicles can be suited to deep space exploration and space object surveillance. Fourthly, the stabilities of the straight line configuration in space information counterwork and the tetrahedron formation in Earth magnetosphere measurement are resolved with the above stable formation investigation. For the former mission, the performance of the straight line configuration is promoted remarkably in two designs using mean and osculating orbital elements. For the latter mission, one stable formation design expressed in mean elements makes a little impact on the line formation at apogee; and the other design expressed in osculating elements can not only increase the average performance of the tetrahedron configuration, but also improve the stability of the relative motion to some extent.Finally, when two spacecrafts have a small difference in area-mass ratio, the characteristics of the relative dynamics under drag and solar radiation pressure are analyzed numerically and two stable formation designs are proposed under the influence of one or three perturbative forces. The difference in semi-major axis or eccentricity is the optimization variable in the first design and the chief spacecraft's inclination, RAAN, argument of perigee and mean anomaly are all used in optimization for the second design. The former is suitable to the mission with given reference orbit, but the latter can't satisfy this requirement. Considering three kinds of perturbations, two stable designs are both able to utilize the effect of drag and solar radiation pressure to minimize that of J₂.