Electrospray thrusters are unique in their ability to emit both positive and negative beams, each contributing similarly to thrust. Spacecraft charging could be prevented without a dedicated neutralizer if currents of simultaneously emitted opposing polarity beams were matched, despite a lack of implicit coupling between them. Both an active current-balance control circuit and a passive self-balancing configuration are discussed and evaluated experimentally. Tests are performed with a highly ionic electrospray source which includes a pair of machined porous glass electrospray emitter arrays in a single holder. In the passive configuration, unbalanced charge collection on a common floating extractor yields rapid charging toward artificially imposed 200 V limits without means for charge suppression, showing the unsuitability of that architecture. Active balancing of emitted currents was fully effective at charge neutralization during stable periods of emission yet insufficient during potential alternations. Methods to improve this approach are presented, including demonstrating beam current modulation at up to 450 Hz, Direct thrust measurements when simultaneously emitting opposing polarity beams of the ionic liquid 1-ethyl-3-methylimidazolium tetrauoroborate are presented. Direct thrust measurements up to 35 mu N were in excellent agreement with commanded levels and computations based on measured currents (up to 190 mu A) and voltages (up to 2100 V).

A performance model of a radio frequency plasma bridge neutralizer was developed to calculate the electrical parameters and optimize the neutralizer design. Minimization of power losses and gas consumption, and a maximization of the neutralizer lifetime and the reliability of the system are requirements of all electric propulsion concepts and strongly determine their future application. The requirements of the neutralizer depend on mission profiles.

Recently, we have set up an Advanced Electric Propulsion Diagnostic (AEPD) platform [1], which allows for the in-situ measurement of a comprehensive set of thruster performance parameters. The platform utilizes a five-axis-movement system for precise positioning of the thruster with respect to the diagnostic heads. In the first setup (AEPD1) an energy-selective mass spectrometer (ESMS) and a miniaturized Faraday probe for ion beam characterization, a telemicroscope and a triangular laser head for measuring the erosion of mechanical parts, and a pyrometer for surface temperature measurements were integrated. The capabilities of the AEPD1 platform were demonstrated with two electric propulsion thrusters, a gridded ion thruster RIT 22 (Airbus Defence & Space, Germany, [13]) and a Hall effect thruster SPT 100D EM1 (EDB Fakel, Russia, [1], [4]), in two different vacuum facilities.

We present an advanced diagnostic system for in situ characterization of electric propulsion thrusters and ion beam sources. The system uses a high-precision five-axis positioning system with a modular setup and the following diagnostic tools: a telemicroscopy head for optical imaging, a triangular laser head for surface profile scanning, a pyrometer for temperature scanning, a Faraday probe for current density mapping, and an energy-selective mass spectrometer for beam characterization (energy and mass distribution, composition). The capabilities of our diagnostic system are demonstrated with a Hall effect thruster SPT-100D EM1.

It is known that the presence of residual gases during on-ground testing of electric propulsion thrusters can significantly influence the plasma parameters and the quantity of charge-exchange ions obtained in the backflow of the ion source. In the frame of an ESA study dedicated to the "Assessment of Interactions between Spacecraft and Electric Propulsion Systems" (AISEPS) a miniaturized μN-RIT thruster (RIT-4), developed by Giessen University, was tested in the Corona vacuum facility at the ESA Propulsion Laboratory (EPL), ESA-ESTEC. A single filament neutralizer was used for beam neutralization. The main ion beam and backflow properties of the thruster were investigated by means of Faraday probes (FP) and retarding potential analyzers (RPA). The Cathode Reference Potential (CRP) was also investigated. The background pressure in the vacuum facility was increased to assess the influence of the xenon residual neutral density on the plume of the thruster. Different electrical coupling configurations between the thruster and the ground were also studied. The μN-RIT was operated at three different thrust levels ranging from 100 to 500 μN while the neutralizer was operated with constant heating voltage during the entire test campaign to allow emission up to 12 mA depending on the electrical grounding configuration. The background pressure was increased in the main vessel of the Corona facility by injecting an auxiliary xenon flow ranging from 10 to 50 sccm. The influence of the neutralizer was clearly observed on the RPA measurements in the main ion beam. The divergence was also clearly correlated to the grounding configuration of the thruster. However, even if the xenon background pressure increased the backflow ion current it did not have a clear influence on the divergence of the thruster.

In recent years, there has been a large increase in the number of small satellites being designed, built and launched. Due to resource constraints, these spacecraft have not generally included any propulsion capability, and this has severely limited mission capabilities and lifetime. To enhance their performances, next generation of small spacecraft will require extremely miniaturized, highly integrated propulsion systems capable to meet stringent mass, volume and power constraints. Two of the most promising technologies to achieve these goals are Electric Propulsion (EP) systems and Micro Electro Mechanical Systems (MEMS). The study identifies a wide variety of mission scenarios, satellites and EP technologies which could benefit from the use of MEMS leading to a selection of one propulsion technology that seems the most promising: the colloid thruster propulsion system. For this technology the requirements are identified and then a preliminary MEMS based EP system design is established. Modularity is very important to enable the same design to be used over and over again and critical subsystem units, such as the high voltage power and control electronics, have also to be developed for the integration in the modular concepts. Two basic design concepts have been investigated to cover the wide range of applications and missions scenario stated within this study and they are presented in this paper.