The feasibility of relocating a small (~500,000 kg) Near-Earth Asteroid (NEA) to High Earth Orbit via Solar Electric Propulsion (SEP) is evaluated with the orbital simulation software General Mission Analysis Tool (GMAT). Using prior research as a basis for the mission parameters, a retrieval mission to NEA 2008 HU4 is simulated in two parts: approach from Earth and return of the Asteroid Redirect Vehicle (ARV) with the asteroid in tow. Success of such a mission would pave the way for future missions to larger NEAs and other deep space endeavors. It is shown that for a hypothetical launch time of 24 May 2016, the ARV could arrive within 25 km of 2008 HU4 on 28 Jun 2017 with a Delta V of 0.406 km/s, begin return maneuver on 08 Dec 2017 and reach Earth altitude of 450,000 km by 23 Apr 2026 with a Delta V of 44.639 m/s.

Testing of hardware to be used in space sometimes involves using vacuum chambers. The need to test hollow cathodes, used as neutralizers for some Electric Propulsion Thrusters, and low power Hall Effect Thrusters at OHB Sweden requires a vacuum chamber to be upgraded. The thesis aims firstly to show the process of adapting the vacuum chamber at OHB to test a heaterless hollow cathode. The requirements of the test included a maximum temperature of a 100 °C in the cathode’s bracket and a pressure inside the chamber around 10¯6 mbar during the test. To accomplish the first, the mounting structure was subjected to a thermal simulation using CREO. Once, the requirement was fulfilled the structure was manufactured and mounted. For the second requirement, a vacuum chamber characterization was done to see the pressure evolution inside the facility. Lastly, the fluid line was designed and mounted in the facility. An upgrade of the current vacuum chamber was deemed insufficient to test low power Hall Effect Thrusters. A comparison between the current vacuum chamber at OHB Sweden and other vacuum chambers designed for this purpose was made. The conclusion was drawn that a new vacuum chamber is necessary. Proposals for the design of a new vacuum facility are presented with particular focus on dimensions, the pump system, sputter protection and thermal protection. / Testandet av hårdvara som ska användas i rymden involverar ibland vakuumkammare. Behovet av att testa hollow cathodes, som används som neutraliserare för vissa jonmotorer, och Hall Effect Thrusters med låg effekt hos OHB Sweden kräver att en vakuumkammare uppgraderas. Avhandlingen syftar först till att redogöra för hur en av vakuumkammarna hos OHB Sweden kan anpassas för att testa en heaterless hollow cathode. Kraven för testet inkluderar en maximal temperatur på 100 °C grader i katodens fäste och och ett vakuum runt 10 ¯6 mbar. En jig designades genom bland annat termisk simulering i CREO. När det termiska kravet ansågs uppfyllas tillverkades och installerades denna jig. För det andra kravet karakteriserades vakuumkammaren med dess vakuumpumpar i syfte att bedöma om dessa kunde uppfylla kravet. Slutligen designades och installerades ett rörsystem som möjliggör matning av bränsle till testobjekten. En uppgradering av nuvarande vakuumkammare bedömdes otillräcklig för att testa Hall Effect Thrusters med låg effekt. En jämförelse mellan nuvarande vakuumkammare på OHB Sweden och andra vakuumkammare designade för detta ändamål gjordes och slutsatsen drogs att en ny vakuumkammare är nödvändig. Förslag på design av ny vakuumkammare presenteras med särkilt fokus på dimensioner, pumpssystemet, sputtingskydd och termiskt skydd.

This work focuses on improving the thrust-to-power ratio of Hall effect thrusters using in-channel electrodes to reduce ion-wall neutralization and focus the ion beam. A higher thrust-to-power ratio would give Hall thrusters increased thrust with the limited power available on spacecraft. A T-220HT Hall thruster is modified in this work to include a pair of ring electrodes within inside the discharge channel. The electrodes are biased above anode potential to repel ions from the walls and toward the channel centerline. Theoretical analysis of ion loss factors indicate that ion-wall neutralizations remove almost 13% of the total ions produced. Reduced wall losses could significantly improve the thruster performance without increased discharge power or propellant consumption. The thruster performance, plume ion characteristics, and internal plasma contours are experimentally measured. The plume and internal plasma measurements are important to determine the cause of the performance changes. The thruster is tested in three conditions: no electrode bias, low bias (10 V), and high bias (30 V). The performance measurements show the electrodes do indeed improve the thrust and thrust-to-power ratio, the latter only at the low bias level. Adding bias increases the ion density and decreases the plume angle compared to the no bias case. The plume measurements indicate that the performance improvements at low bias are due to increased ion number density as opposed to increased ion energy. The increased ion density is attributed to reduced wall losses, not increased ionization. The in-channel measurements support this due to little change in the acceleration potential or the electron temperature. At the high bias level, a drop in thrust-to-power ratio is seen, even though a larger increase in thrust is observed. This is due to increased power draw by the electrodes. Plume measurements reveal the increased thrust is due to ion acceleration. The internal measurements show increased acceleration potential and electron energy which can lead to increased ionization. At the high bias condition, the electrodes become the dominant positive terminal in the thruster circuit. This causes the increased ion acceleration and the creation of domed potential contours that conform to the near-wall cusp-magnetic fields. The domed contours produce focused electric fields, which cause the decreased wall losses and plume angle.