Satellite de-orbiting by means of electrodynamic tethers. Part II: System configuration and performance (Articolo in rivista)

Type
Label
  • Satellite de-orbiting by means of electrodynamic tethers. Part II: System configuration and performance (Articolo in rivista) (literal)
Anno
  • 2002-01-01T00:00:00+01:00 (literal)
Alternative label
  • L. Iess, C. Bruno, C. Ulivieri, G. Vannaroni (2002)
    Satellite de-orbiting by means of electrodynamic tethers. Part II: System configuration and performance
    in Acta astronautica
    (literal)
Http://www.cnr.it/ontology/cnr/pubblicazioni.owl#autori
  • L. Iess, C. Bruno, C. Ulivieri, G. Vannaroni (literal)
Pagina inizio
  • 407 (literal)
Pagina fine
  • 416 (literal)
Http://www.cnr.it/ontology/cnr/pubblicazioni.owl#numeroVolume
  • 50 (literal)
Rivista
Http://www.cnr.it/ontology/cnr/pubblicazioni.owl#descrizioneSinteticaDelProdotto
  • Il lavoro descrive la configurazione di un sistema spaziale a filo elettrodinamico installato a bordo di satelliti in orbita bassa (LEO) per ottenerne la deorbitazione al termine della loro vita operativa. (literal)
Note
  • ISI Web of Science (WOS) (literal)
Http://www.cnr.it/ontology/cnr/pubblicazioni.owl#affiliazioni
  • Scuola di Ingegneria Aerospaziale, Universita' di Roma \"La Sapienza\", Istituto di Fisica dello Spazio Interplanetario CNR, Roma (literal)
Titolo
  • Satellite de-orbiting by means of electrodynamic tethers. Part II: System configuration and performance (literal)
Abstract
  • This paper aims to assess the efficiency of a de-orbiting system based upon conductive tethers under realistic assumptions for its interaction with the ionospheric environment. We analyze the configuration made up of a 2.5-10 km tether, a passive inflatable collector of 2.5-10m radius at the positive termination and a hollow cathode at the negative one. Voltages and current in the system are computed from the equation of the equivalent circuit, making use of the IRI-90 ionospheric model. The resulting electromagnetic drag forces have been used to compute the evolution of the orbital elements (especially the semi-major axis) and the re-entry times. Our results indicate that a typical satellite of 500 kg mass at 1300 km altitude can de-orbit in 20-100 days, for a broad range of orbital inclinations and solar activity. The validity of the concept is further strengthened by the comparison with alternative propulsion systems. (literal)
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