Spacecraft escape and capture trajectories from or to Halo orbits about the L1 or L2 points using impulsive maneuvers at periapsis of the manifolds for interplanetary transfers are analyzed in the restricted Hill three-body problem. This application is motivated by future proposals to place deep-space ports at the Earth andMars L1 or L2 points. First, the feasibility of interplanetary trajectories between Earth Halo orbits and Mars Halo orbits is investigated. In this study, unstable and stable manifolds associated with the Halo orbits are used to approach the vicinity of the planet’s surface, and use impulsive maneuvers at periapsis for escape and capture trajectories to and from Halo orbits. Interplanetary trajectories between Earth and Mars Halo orbits with reasonable V and flight time are found. Next, applying these dynamics to an Earth–Mars transportation system using spaceports on Earth and Mars Halo orbits, the system is evaluated in terms of the spacecraft mass of round-trip transfer. As a result, transfer between low Earth orbits and low Mars orbits via the planets’ Halo orbits can reduce spacecraft wet mass compared with a direct round-trip transfer, by leaving propellant for the return.
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.
Two spacecraft or more are assumed to be in a state of loose formation flying around a collinear Lagrangian point in the Sun-Earth circular restricted three-body problem (CR3BP) system. The orbit reference of choice for the leader is a halo orbit, and the followers are assumed to follow nearby and be constrained either geometrically or in size. This type of formation could be useful in the future for constructing space ports, space telescopes, astronomical spacecraft requiring sun shields and, with greater numbers, spacecraft swarm missions. The formation design method is constructed by firstly seeking the local coordinate system from the monodromy matrix through extraction of the independent bases that span the space of the halo orbit. To nullify diverging and converging motion, we confine the relative motion to within the periodic subspaces. We observe two modes of relative motion within these subspaces, long-term and short-term motion. In this study, we approximate the long-term motion by deriving a discrete formulation of independent directions based on the eigenvectors of the monodromy matrix, while for the short-term motion we approximate the fundamental set solutions using Fourier series and additional linear functions. Since the size of the formation discussed is significantly smaller than that of the halo orbit, the formation design method can fundamentally be stated as a process of linearly combining these approximations to achieve the desired formation. Consequently, use of this approach transforms formation design from a differential equation problem into an algebraic one, and furthermore enables the long-term and short-term motion design problems to be handled either jointly or separately. A set of design examples is presented to demonstrate the validity of the design method.
This paper reports on the conceptual design of a three-stage launch vehicle (LV) with a clustered hybrid rocket engine (HRE) through multi-disciplinary design optimization. This LV is a space transportation concept that can deliver micro-satellites to sun-synchronous orbits (SSOs). To design a high-performance LV with HRE, the optimum size of each component, such as an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank, and a nozzle, should be determined. In this study, paraffin (FT-0070) is used as a propellant for the HRE, and three cases are compared: In the first case, HREs are optimized for each stage. In the second case, HREs are optimized together for the first and second stages but separately for the third stage. In the third case, HREs are optimized together for each stage. The optimization results show that the performance of the design case that uses the same HREs in all stages is 40% reduced compared with the design case that uses optimized HREs for each stage.
This paper investigates the preliminary orbit maintenance for the next-generation infrared astronomical mission, Space InfraredTelescope forCosmology andAstrophysics, which is to be launched into the sun–Earth L2 halo orbit. Particularly, the impact of the reaction wheel unloading V on the spacecraft trajectory from the viewpoint of orbit maintenance is analyzed. In addition, the nondivergent and suppressive methods using the characteristics of the dynamical theory for unloading V are proposed and validated.
Spacecraft capture trajectories to the periodic orbits of the L1 and L2 points in the restricted Hill three-body problem are studied. The specific focus is on transfer to these vicinities from interplanetary trajectories. This application is motivated by future proposals to place Deep Space ports at the Earth and Mars L1 or L2 points. These spaceports are considered as candidate gateways for interplanetary transfers in the future. We utilize stable manifolds for capture trajectories to periodic orbits around the libration points. As a result, the cost of capture into a periodic orbit is also reduced relative to direct capture into a parabolic orbit. The way of linking between interplanetary transfer trajectories and the stable manifold is also discussed.
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