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Electric Solar Wind Sail Optimal Transit in the Circular Restricted Three Body Problem
Get PDFThis paper analyzes the transfer orbits within a Sun-[Earth+Moon] system for a spacecraft whose primary propulsion system is an Electric Solar Wind Sail. The planetary system is approximated
through the Circular Restricted Three Body Problem and the spacecraft motion is studied in an optimal framework in which the performance index is the flight time. Minimum time transfers are studied using an indirect approach, and the optimal control law is found in analytical form as a function of the problem parameters. Optimal transfers between equilibrium points are discussed and interesting symmetries in the spacecraft trajectories are pointed out along with an analytical proof of their existence. A mission scenario consistent with the Geostorm concept is analyzed and the effectiveness of the propulsion system is emphasized for missions involving a tour through a subset of the classical Lagrangian points
Artificial Equilibrium Points for Electric Sail with Constant Attitude
Get PDFCreating and maintaining Artificial Equilibrium Points (AEPs) in the restricted threebody problem is a challenging mission scenario in which a propellantless propulsion system exploits its natural potential. Indeed, in such a problem the acceleration resulting from the sum of centrifugal and gravitational forces can be balanced, for a theoretically unlimited time period, by means of a suitable continuous propulsive thrust.
A thorough analysis involving the location and stability of AEPs has been addressed in a recent paper, under the assumption that the propulsion system provides a purely radial thrust with respect to the Sun, and the thrust modulus is a function of the Sun-spacecraft distance only . In that way, with a unified mathematical model, it is possible to analyze the performance of different propulsion systems, as, for example, a photonic solar sail and an
Electric Solar Wind Sail (E-Sail). In particular, an E-Sail is known to be able to provide a continuous propulsive acceleration by means of Coulomb’s interaction of a number of positively charged tethers with the solar wind plasma stream
Optimal Control Laws for Heliocentric Transfers with a Magnetic Sail
Get PDFA magnetic sail is an advanced propellantless propulsion system that uses the interaction between the solar wind and an artificial magnetic field generated by the spacecraft, to produce a propulsive thrust in interplanetary space. The aim of this paper is to collect the available experimental data, and the simulation results, to develop a simplified mathematical model that describes the propulsive acceleration of a magnetic sail, in an analytical form, for mission analysis purposes. Such a mathematical model is then used for estimating the performance of a magnetic sail-based spacecraft in a two-dimensional, minimum time, deep space mission scenario. In particular, optimal and locally optimal steering laws are derived using an indirect approach. The obtained results are then applied to a mission analysis involving both an optimal Earth-Venus (circle-to-circle) interplanetary transfer, and a locally optimal Solar System escape trajectory. For example, assuming a characteristic acceleration of 1 mm/s(2), an optimal Earth-Venus transfer may be completed within about 380 days
Artificial Periodic Orbits Around L1-Type Equilibrium Points for a Generalized Sail
Get PDFThe contribution of this note is to extend the available results for APOs maintained by a propellantless propulsion system to the case of purely radial (continuous) propulsive acceleration, whose modulus depends on a given power of the Sun–spacecraft distance
A Graphical Approach to Electric Sail Mission Design with Radial Thrust
No full textThis paper describes a semi-analytical approach to electric sail mission analysis under the assumption that the spacecraft experiences a purely radial, outward, propulsive acceleration. The problem is tackled by means of the potential well concept, a very effective idea that was originally introduced by Prussing and Coverstone in 1998. Unlike a classical procedure that requires the numerical integration of the equations of motion, the proposed method provides an estimate of the main spacecraft trajectory parameters, as its maximum and minimum attainable distance from the Sun, with the simple use of analytical relationships and elementary graphs. A number of mission scenarios clearly show the effectiveness of the proposed approach. In particular, when the spacecraft parking orbit is either circular or elliptic it is possible to find the optimal performances required to reach an escape condition or a given distance from the Sun. Another example is given by the optimal strategy required to reach a heliocentric Keplerian orbit of prescribed orbital period. Finally the graphical approach is applied to the preliminary design of a nodal mission toward a Near Earth Asteroid
Artificial Lagrange Points for Solar Sail with Electrochromic Material Panels
Get PDFThe aim of this Note is to explore the capabilities of the emerging EMP technology for the active stabilization of L1-type AEPs using a square solar sail with a fixed attitude. The problem is addressed within an elliptic restricted framework, which is a more realistic model with respect to the classical circular case [4]. The main spacecraft parameters, including the sail side and the total spacecraft mass, are defined, by means of a simplified mathematical model, as a function of the main mission requirements in terms of maximum allowed sail lightness number variation and AEP position
Multi-Revolution Transfer for Heliocentric Missions with Solar Electric Propulsion
Get PDFAn extension of the classical method by Alfano, for the analysis of optimal circle-to-circle two-dimensional orbit transfer, is presented for a deep space probe equipped with a solar electric primary propulsion system. The problem is formulated as a function of suitable design parameters, which allow the optimal transfer to be conveniently characterized in a parametric way, and an indirect approach is used to find the optimal steering law that minimizes the required propellant mass. The numerical results, obtained by solving a number of optimal control problems, are arranged into contour plots, characterized by different and well-defined behaviors depending on the value of the initial spacecraft propulsive acceleration, the final orbit radius, and the thruster's specific impulse. The paper presents also a semi-analytical mathematical model for preliminary mission analysis purposes, which is shown to give excellent approximations of the (exact) numerical solutions when the number of revolutions of the spacecraft around the Sun is greater than five. An Earth Mars cargo mission has been thoroughly investigated to validate the proposed approach. In this case, assuming a propulsion system with a specific impulse of 3000 s (comparable to that installed on the Deep Space 1 spacecraft), the results obtained with the semi-analytical model coincide, from an engineering point of view, with the numerical solutions both in terms of total mission time (about 8.3 years) and propellant mass fraction required (about 17.5%). By decreasing the value of the specific impulse, the differences between the results from the semi-analytical model and the numerical simulations tend to increase. However, good results are still possible if the number of revolutions of the spacecraft around the Sun is close to an integer number
Artificial Equilibrium Points for a Generalized Sail in the Elliptic Restricted Three-Body Problem
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