2nd International Mars Sample Return 2018 (LPI Contrib. No. 2071) 6004.pdf

THE DESIGNING OF THE MISSION TO MARS FOR THE SOIL DELIVERY BY THE ELECTRIC PROPULSION SPACECRAFT. O. L. Starinova1, Ch. Gao2, D. V. Kurochkin3 and H. Yudong4, 1Professor of Space Engineering Department, Samara National Research University, Russian Federation, [email protected], 2Accosiate professor of Harbin University of Technology, People's Republic of China, [email protected], 3Engineer of Space Engineering Department, Samara National Research University, Russian Federation, 4Master student , Sa- mara National Research University, Russian Federation.

Introduction: Mission to return soil from Mars is choosing the design parameters corresponding to the a typical kind of transfer with the return to the source minimum spacecraft launching mass. orbit. The effectiveness of such a flight depends on The movement of the electric propulsion spacecraft many parameters describing the design scheme of is described by the system of the differential equations spacecraft and the features of the trajectory. It is well take into account the following assumptions. The known [1]-[4] that the electric propulsion spacecraft is spacecraft contemporaneously is subject to the gravita- the most useful for interplanetary flights at the present tion of the one-body (a planet or the Sun). The gravita- stage of space technology development. tional field of all bodies isn't central. The Earth and The problem of comprehensive optimization of the Mars atmospheres are standard. The planet orbits rela- trajectory and the ballistic parameters of a low-thrust tive to the Sun are known. mission Earth-Mars-Earth is formulated and solved in A design vector which uses to create the design this paper. scheme consists of the independent variables which Models and Methods: The problem of the Earth- significantly impact on the launched mass. A change in Mars-Earth mission optimization is formulated with each of these variables leads to a different design of the restrictions on the total mission duration and the dis- spacecraft. In this study, we chose specific impulse and tance to the Sun. Besides this, the heliocentric parts of initial acceleration as design variables. The design trajectory are calculated by taking the ellipticity of model of the spacecraft can be obtained based on a planetary orbits into account. The aim of considering mass balance equation in the parking orbit. This equa- dynamic space maneuver is the payload transportation tion represents the initial mass as the sum of the masses to the target areocentric orbit and returning its part to of its general subsystems, the fuel, and the payload. the initial geocentric orbit. The primary optimality cri- Results: With the use of the described technique, terion is a launching mass of spacecraft given the lim- the results of the ballistic optimization of the missions ited flight duration. to the Mars were received. A large number of the This problem formulation is a protracted process solved optimal control problems allowed to obtain the which requires the vast computing resources. Common- approximation dependences between the intermediary ly, this problem is solved by either direct or indirect criterion and the design-ballistic parameters of the optimization methods using numerical integration [1], flights Earth-Mars and Mars-Earth [2]. These depend- [2], [4] of the motion equations. This approach results encies enable automating the process of optimizing the in significant computational difficulties and doesn't design scheme of spacecraft.The proposed method is allow to analyze the various design and ballistic pa- especially useful on pre-conceptual design stages when rameters of the mission. However, the decision of op- the possibility of the fast analysis is more important timization problem can be considerably simplified if than the accurate optimization on more complex and the general problem is divided into the dynamic and the exact models. parametric parts. The basic idea of this division is the References: choice of such an intermediate criterion of optimization [1] Grodzovskiĭ G. L., Ivanov I. N., and Tokarev (the flight characteristics), which would allow one to V. V. (1969). Mechanics of spaceflight low-thrust: determine the dependence of mission costs on the set of (Mekhanika kosmicheskogo poleta s maloi tyagoi) spacecraft design parameters. (Vol. 507). Israel Program for Scientific Translations. This intermediate criterion may be different for dif- [2] Ishkov S. A., Milokumova O. L., and Salmin V.V. ferent kinds of spacecraft, but in any case, it allows to (1995) Optimization of Low-Thrust Roundtrip Earth- split the problem into two components. The first one is Mars-Earth Flights, Kosm. Issled., 33(2), 210-219. [3] the dynamic part of the optimization problem which Braun R. D. and Blersch D. J. (1991). Propulsive op- consists in determining the control functions and the tions for a manned Mars transportation system. JSR, ballistic parameters providing for a minimum flight 28(1), 85-92. [4] Landau D. F. and Longuski J. M. characteristic at fixed design parameters. The second (2004). A reassessment of trajectory options for human one is the parametric part of the problem which allows missions to Mars. AIAA Paper, 5095, 16-19.