Martian Moons and Space Transportation Using Chemical and Electric Propulsion Options

Using chemical and nuclear electric propulsion for the exploration of the Martian moons will be investigated. Both oxygen/hydrogen chemical propulsion and nuclear electric propulsion with 500 kilowatt electric (kWe) to 10 megawatt electric (MWe) reactors will be assessed. The initial masses, propellant masses, and trip times for a variety of space vehicle payload masses will be compared. For high energy orbital transfer, the nuclear electric propulsion vehicles required a small fraction of the propellant mass over oxygen/hydrogen orbital transfer vehicles (OTVs). The moons, Phobos and Deimos, may hold resources for refueling future space vehicles. In-situ resource utilization (ISRU) can be a powerful method of reducing Earth dependence on space vehicle propellants, liquid water, and breathing gases. Historical studies have identified the potential of water in carbonaceous chondrites on the moons. The moon-derived propellants OTVs that move payloads between the moons and to other important operational Mars orbits. Also, the propellants have been suggested to support reusable Mars landers. To extract the water, the mined mass, its volume and the mining time were estimated. The water mass fraction may be as low as 2x10−4. Very large masses were needed to be extracted for up to 100 MT of water.

On the choice of a parking orbit for a service spacecraft

At present, the requirements for increasing spacecraft active life and operational reliability and reducing spacecraft operation costs become more and more stringent. Because of this, on-orbit servicing becomes more and more attractive. One of the most promising ways to increase the efficiency of transport operations in space is to carry out on-orbit servicing using reusable spacecraft with low-thrust solar electrojet engines. The aim of this paper is to develop a mathematical model for the choice of an optimal low near-Earth parking orbit for a reusable service spacecraft. The case of noncoplanar near-circular orbits of spacecraft and a shuttle scenario of their servicing is considered. The solution of the problem of choosing an optimal parking orbit for a reusable service spacecraft involves repeated solutions of the problem of determining the delta-velocity of the service spacecraft’s orbital transfers between its parking orbit and the orbits of the serviced spacecraft. In this connection, using the averaging method, a mathematical model is developed for the analytical determination of orbital transfer program controls and trajectories and assessing orbital transfer energy expenditures. With its use, a mathematical model is developed for the choice of a service spacecraft’s optimal parking orbit. The objective function is the total delta-velocity of the service spacecraft’s orbital transfers from its parking orbit to the orbits of the serviced spacecraft and vice versa with the inclusion of the orbital transfer frequency. The optimizable parameters are the service spacecraft parking orbit parameters. The use of the proposed models is illustrated by an example of service spacecraft parking orbit optimization. What is new is the mathematical models developed. The results obtained may be used in the preliminary planning of on-orbit servicing operations.

Reconfigurable Satellite Constellation Design for Disaster Monitoring Using Physical Programming

Data collection by satellites during and after a natural disaster is of great significance. In this work, a reconfigurable satellite constellation is designed for disaster monitoring, and satellites in the constellation are made to fly directly overhead of the disaster site through orbital transfer. By analyzing the space geometry relations between satellite orbit and an arbitrary disaster site, a mathematical model for orbital transfer and overhead monitoring is established. Due to the unpredictability of disasters, target sites evenly spaced on the Earth are considered as all possible disaster scenarios, and the optimal reconfigurable constellation is designed with the intention to minimize total velocity increment, maximum and mean reconfiguration time, and standard deviation of reconfiguration times for all target sites. To deal with this multiobjective optimization, a physical programming method together with a genetic algorithm is employed. Numerical results are obtained through the optimization, and different observation modes of the reconfigurable constellation are analyzed by a specific case. Superiority of our design is demonstrated by comparing with the existing literature, and excellent observation performance of the reconfigurable constellation is demonstrated.

Optimization of Interplanetary Trajectories Using the Colliding Bodies Optimization Algorithm

In this paper, a recent physics-based metaheuristic algorithm, the Colliding Bodies Optimization (CBO), already employed to solve problems in civil and mechanical engineering, is proposed for the optimization of interplanetary trajectories by using both indirect and direct approaches. The CBO has an extremely simple formulation and does not depend on any initial conditions. To test the performances of the algorithm, missions with remarkably different orbital transfer energies are considered: from the simple planar case, as the Earth-Mars orbital transfer, to more energetic ones, like a rendezvous with the asteroid Pallas.

Magnetic attitude tracking control of gravity gradient microsatellite in orbital transfer

ABSTRACTThe problem of the magnetic attitude tracking control is studied for a gravity gradient microsatellite in orbital transfer. The contributions of the work are mainly shown in two aspects: (1) the design of an expected attitude trajectory; (2) a method of the magnetic attitude tracking control. In orbital transfer, the gravity gradient microsatellite under a constant thrust shows complicated dynamic behaviours. In order to damp out the pendular motion, the gravity gradient microsatellite is subject to the the attitude tracking problem. An expected attitude trajectory is designed based on dynamic characteristics revealed in the paper, which not only ensures the flight safety of the system, but also reduces the energy consumption of the controller. Besides, the control torque produced by a magnetorquer is constrained to lie in a two-dimensional plane orthogonal to the magnetic field, so an auxiliary compensator is proposed to improve the control performance, which is different from existing magnetic control methods. In addition, a sliding mode control based on the compensator is presented, and the Lyapunov stability analysis is performed to show the global convergence of the tracking error. Finally, a numerical case of the gravity gradient microsatellite is studied to demonstrate the effectiveness of the proposed tracking control.