Spacecraft capture trajectories to Lyapunov/Halo periodic orbits of the L1 and L2 points in the restricted Hill three-body problem are analyzed. The specific focus is on transfer to these orbits from interplanetary trajectories. This application is motivated by future proposals to place "deep space ports" at the Earth and Mars L1 or L2 points. We use stable manifolds for capture trajectories to periodic orbits around the libration points. Numerical results show that the stable and unstable manifolds from periodic orbits around the libration points can intersect the surface of any of the planets of the solar system by changing the size of periodic orbits. Applying this to Earth-Mars transfers, the cost of capture into a periodic orbit is reduced compared with direct capture into a parabolic orbit. Moreover, if a spaceport is built on a periodic orbit in the vicinity of a sun-Mars libration point and propellant can be supplied there to the spacecraft, the required V for entry into a circular orbit about Mars from an interplanetary trajectory can be considerably reduced compared with a direct entry into a circular orbit.

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The saddle points are locations where the net gravitational accelerations balance. These regions are gathering more attention within the astrophysics community. Regions about the saddle points present clean, close-to-zero background acceleration environments where possible deviations from General Relativity can be tested and quantified. Their location suggests that flying through a saddle point can be accomplished by leveraging highly nonlinear orbits.In this paper, the geometrical and dynamical properties of the Sun-Earth saddle point are characterized. A systematic approach is devised to find ballistic orbits that experience one or multiple passages through this point. A parametric analysis is performed to consider spacecraft initially on L 1,2 Lagrange point orbits. Sun-Earth saddle point ballistic fly-through trajectories are evaluated and classified for potential use. Results indicate an abundance of short-duration, regular solutions with a variety of characteristics.

A single-stage launch vehicle with hybrid rocket engine has been conceptually designed by using design informatics, which has three points of view as problem definition, optimization, and data mining. The primary objective of the design in the present study is that the sufficient down range and duration time in the lower thermosphere are achieved for the aurora scientific observation whereas the initial gross weight is held down. The multidisciplinary design optimization and data mining were performed by using evolutionary hybrid computation under the conditions that polypropylene as solid fuel and liquid oxygen as liquid oxidizer were adopted and that single-time ignition is implemented in sequence. Consequently, the design information regarding tradeoffs and the behavior of the design variables was obtained.