The Electric Solar Wind Sail (E-sail) is an innovative, propellantless space propulsion concept based on the interaction of a number of electrically biased tethers with the protons of the solar wind. The spacecraft and the tethers are kept at a high positive potential with respect to surrounding plasma by means of an electron gun in order to enhance the electrically effective sail area and generate a net thrust on the spacecraft. This paper deals with the preliminary design of a small and flexible E-sail test mission, considering the constraints imposed by small and medium class launchers for primary and piggyback payloads. In particular, the Ariane 5 ASAP (Ariane Structure for micro-Auxiliary Payloads) is assumed as a reference, considering a standard Kourou Geosynchronous Transfer Orbit (GTO) as the spacecraft release orbit. The target orbit for this technology demonstration mission is a highly elliptical orbit with the same perigee altitude of the initial orbit (600 km) and an apogee altitude of at least 25 Earth radii (~165800 km), so as to operate the payload, composed by the tethers, the high voltage system and the required diagnostics, out of the Earth’s magnetosphere. This target orbit guarantees a sufficient level of electrostatic interaction between the charged tethers and the solar wind, so that thrust can be measured by on-board accelerometers. Considering the relevant total impulse typical of such transfer maneuvers, an electric propulsion system, based on a low power Hall Effect Thruster, is envisaged to provide the spacecraft the required velocity increment with a small propellant mass fraction. The spacecraft is meant to be equipped with four triple junction GaAs/Ge solar arrays and the energy produced by solar panels is stored in a space proven Li-Ion battery pack. The mission design has been carried out considering a set of pericenter-centered tangential thrusting arcs, to raise the orbital apogee, and a final phase of continuous thrusting along orbit semi-minor axis, to increase the orbital eccentricity. Third body perturbations, Earth oblateness and eclipse phases (during which the thruster is supposed to be switched off to limit the battery pack mass) are accounted for. This approach results in a transfer maneuver requiring about 12% of propellant mass fraction with a total required velocity increment of 1.62 km/s. The total transfer time is of about 130 days with approximately 1000 hours of cumulative thruster firing time. The spacecraft total mass is estimated sizing each of its main subsystem and results to be of the order of 65 kg. For all the subsystems Commercial Off The Shelf (COTS) components have been preferred so as to reduce costs and increase system reliability.
Preliminary Design of the E-sail Test Mission
MARCUCCIO, SALVO;
2012-01-01
Abstract
The Electric Solar Wind Sail (E-sail) is an innovative, propellantless space propulsion concept based on the interaction of a number of electrically biased tethers with the protons of the solar wind. The spacecraft and the tethers are kept at a high positive potential with respect to surrounding plasma by means of an electron gun in order to enhance the electrically effective sail area and generate a net thrust on the spacecraft. This paper deals with the preliminary design of a small and flexible E-sail test mission, considering the constraints imposed by small and medium class launchers for primary and piggyback payloads. In particular, the Ariane 5 ASAP (Ariane Structure for micro-Auxiliary Payloads) is assumed as a reference, considering a standard Kourou Geosynchronous Transfer Orbit (GTO) as the spacecraft release orbit. The target orbit for this technology demonstration mission is a highly elliptical orbit with the same perigee altitude of the initial orbit (600 km) and an apogee altitude of at least 25 Earth radii (~165800 km), so as to operate the payload, composed by the tethers, the high voltage system and the required diagnostics, out of the Earth’s magnetosphere. This target orbit guarantees a sufficient level of electrostatic interaction between the charged tethers and the solar wind, so that thrust can be measured by on-board accelerometers. Considering the relevant total impulse typical of such transfer maneuvers, an electric propulsion system, based on a low power Hall Effect Thruster, is envisaged to provide the spacecraft the required velocity increment with a small propellant mass fraction. The spacecraft is meant to be equipped with four triple junction GaAs/Ge solar arrays and the energy produced by solar panels is stored in a space proven Li-Ion battery pack. The mission design has been carried out considering a set of pericenter-centered tangential thrusting arcs, to raise the orbital apogee, and a final phase of continuous thrusting along orbit semi-minor axis, to increase the orbital eccentricity. Third body perturbations, Earth oblateness and eclipse phases (during which the thruster is supposed to be switched off to limit the battery pack mass) are accounted for. This approach results in a transfer maneuver requiring about 12% of propellant mass fraction with a total required velocity increment of 1.62 km/s. The total transfer time is of about 130 days with approximately 1000 hours of cumulative thruster firing time. The spacecraft total mass is estimated sizing each of its main subsystem and results to be of the order of 65 kg. For all the subsystems Commercial Off The Shelf (COTS) components have been preferred so as to reduce costs and increase system reliability.I documenti in IRIS sono protetti da copyright e tutti i diritti sono riservati, salvo diversa indicazione.
