Device for reducing the impact on a surface section by positively charged ions, and ion accelelerator arrangement
The invention relates to a surface section that is exposed to an ion flow, in particular for a drive arrangement in a spacecraft, comprising an electrostatic ion accelerator arrangement. According to the invention, in order to reduce erosion, an intermediate potential energy surface is provided, the surface being advantageous in that it allows a magnetic field that is essentially parallel to the surface section to be formed.
1. A spacecraft having a mass potential and comprising:
first outside surfaces lying at the mass potential; and
an ion accelerator arrangement comprising:
an ionization chamber having a beam exit opening and generating ions from a working gas introduced into the ionization chamber;
an electrode arrangement configured to electrostatically accelerate the generated ions via a static high voltage, to electrostatically eject from the beam exit opening the generated ions as a plasma beam, and to supply electrons to the ionization chamber for the ionization in the ionization chamber; and
a shielding surface laterally offset relative to the beam exit opening, surrounding the beam exit opening, and facing the ejected plasma beam during operation of the ion acceleration arrangement, said shielding surface having a funnel-shaped body with a smooth surface facing the ejected plasma beam, wherein said funnel-shaped body consists of metal that is not soft magnetic and does not influence a magnetic field, or of a dielectric;
wherein the shielding surface lies spatially between the ejected plasma beam and the first outside surfaces;
wherein the shielding surface lies electrically at a potential separated from the mass potential of the spacecraft; and
wherein the ion accelerator arrangement reduces impact of positively charged ions of the ejected plasma beam on a surface section of the spacecraft.
2. The spacecraft according to claim 1 , wherein the shielding surface is at least partly electrically non-conductive.
3. The spacecraft according to claim 1 , wherein the shielding surface is electrically insulated with regard to a mass potential of the ion accelerator arrangement.
4. The spacecraft according to claim 1 , wherein the shielding surface lies at sliding potential in a drive operation of the spacecraft.
5. The spacecraft according to claim 1 , wherein the shielding surface is offset, relative to a mass potential of the ion accelerator arrangement, by a defined voltage, in a drive operation of the spacecraft.
6. The spacecraft according to claim 1 , wherein the electrode arrangement comprises a cathode electrode disposed outside the ionization chamber and laterally offset relative to the beam exit opening.
7. The spacecraft according to claim 6 , wherein the shielding surface runs in a radial direction between the cathode electrode and the beam exit opening.
8. The spacecraft according to claim 1 , further comprising a magnet arrangement generating a magnetic field having a progression predominantly parallel to the shielding surface in a spatial region between the shielding surface and a spatial region of the ejected plasma beam.
9. The spacecraft according to claim 8 , wherein the magnetic field forms a cusp structure between a center longitudinal axis of the ionization chamber and an edge of the beam exit opening, and
wherein the cusp structure has a predominantly radial magnetic field progression.
10. The spacecraft according to claim 1 , wherein the shielding surface comprises
a carrier body having a first sputter-resistance, and
a coating on the carrier body having a second sputter-resistance, the second sputter-resistance being greater than the first sputter-resistance.
11. The spacecraft according to claim 1 , wherein a beam direction of the plasma beam is directed away from the spacecraft, and
wherein the ion accelerator arrangement operates as a drive device for the spacecraft.