Michele Coletti

PPT Development for Nanosatellite Applications: Experimental Results

In this paper, the design and performances of a pulsed plasma thruster (PPT) for nanosatellite applications will be presented. The breadboard model PPT presented in this paper will be a part of six PPTs propulsion system designed to provide attitude and translational control for a 20-kg nanosatellite for a total delta-V of 40 m/s. The thruster performances have been characterized in terms of electrical parameters, mass bit, impulse bit, and specific impulse. The thruster was found to be compliant with the mission requirement. Moreover, preliminary electromagnetic noise interference measurements have been performed. The spark plug discharge was found to be the main source of noise as already found by previous authors. From the collected data, it can be inferred that the noise is mostly radiated.

An impulsive thrust balance for applications of micro-pulsed plasma thrusters

This paper describes the design and testing of a direct torsional impulsive thrust balance. The design philosophy allows the balance to measure impulse bits (Ibit) in the range of 20?120??N?s typical of pulsed plasma thrusters (PPTs) for pico- and nano-satellites. The uncertainty in the Ibit measurement is quantified to be about 8.8%, smaller than the typical values of this kind of balance (between 12% and 15%). This has been possible due to an in-depth analysis of all the possible sources of disturbance, which allows the choice of the most suitable measurement and estimation methods to minimize the errors. The balance has successfully been used for testing two PPTs with different propellant feeding system, nominal energy, mass and delivered impulse bits.

Electric Propulsion Subsystem Architecture for an All-Electric Spacecraft

For many space missions, both a main propulsion subsystem and additional attitude control (AOCS) propulsion subsystem are required. These subsystems normally use different propellants, hence require separate tanks, different flow control units (FCU) and, in case of solar electric propulsion (SEP), separate power processing units (PPU). This leads to increases in total mass of the spacecraft and complexity while reducing system specific impulse. One possibility to alleviate this problem would be to develop a main and an AOCS propulsion technology which could be integrated, sharing some of the components required for their operation, hence reducing system mass. A spacecraft employing such combined technologies as part of an SEP system is referred to as an “All-electric-spacecraft” (Wells et al., 2006). In this chapter, the system design for an all-electric-spacecraft will be presented. A gridded ion engine (GIE) is proposed as a main propulsion subsystem with hollow cathode thrusters (HCT) considered for the AOCS propulsion subsystem. The mission considered during this study is the ESA European Student Moon Orbiter (ESMO), which the University of Southampton proposed to use SEP for both attitude control and main propulsion. During the ESMO phase-A study, a full design of the SEP subsystem was performed at QinetiQ as part of a wider study of the mission performed in conjunction with QinetiQ staff and funded by ESA. The output of this study will be here presented to explain the concept of the all-electric-spacecraft, its benefits, drawbacks and challenges.

Barium Oxide Depletion from Hollow-Cathode Inserts: Modeling and Comparison with Experiments

In this paper, the results of a barium oxide depletion model are compared with some experimental results. The model is used to simulate the T5 and T6 cathodes and the NSTAR discharge cathode. A comparison with the experimental data is performed. For the T5 and T6 cathodes, good qualitative agreement is found, but for theNSTAR cathode, the agreement is not as good. In both cases, the agreement is improved when the boundary conditions are modified to better reflect the experimental conditions. The model presented is the first three-dimensional axisymmetric insert model that includes the dependency of BaO depletion from both the impregnant chemistry and the diffusive motion inside the insert.

A micro PPT for Nano-satellite applications: Design and experimental results

One current trend in space technology is towards miniaturisation of spacecraft, from mini to micro, pushing towards nano-sats. A modular µPPT thruster system is being developed to be integrated and customized to accommodate different nano-satellites mission requirements. A study case scenario has been selected and will be used to demonstrate the potential application that the µPPT module could provide to a nano-satellite. A six µPPT propulsion system to perform station keeping and attitude control manoeuvres (ΔV=40m/s, Total Impulse=800Ns) is currently under development by Mars Space Ltd, the University of Southampton and Clyde Space Ltd under an ESA funded project. In this paper the design of the thruster will be presented together with experimental results.

Insert Temperature Measurements of a 180A Hollow Cathode for the HiPER Project

In the scope of the HiPER project a 180A cathode to be used as discharge cathode inside high power GIE or as neutralizer cathode for high power HETs has been designed and tested. The main goal of the testing was to measure the insert temperature distribution along the cathode insert to verify the cathode ability of providing the required lifetime of 17,000 h. In this paper the design will be quickly presented and the measurements reported. The measured data will be then analyzed to verify if the cathode is able to meet the required lifetime.

A Chemical Model for Barium Oxide Depletion from Hollow Cathode's Insert

In this paper a chemical model to predict barium oxide depletion from hollow cathode insert is developed. This model is based on the knowledge of the ternary diagram that exists in the BaO-CaO-Al2O3 system. This model takes also into account the diffusive motion of barium oxide inside the insert. A comparison between numerical and experimental data is made to determine the diffusion coefficient inside the insert. The diffusion coefficient presents an Arrhenius trend with activation energy similar to the one of barium oxide evaporation. A two dimensional model is used to demonstrate how temperature profile along the insert is a key parameter to calculate barium depletion.

Simple Thrust Formula for an MPD Thruster with Applied- Magnetic Field from Magnetic Stress Tensor

It has been previously shown that the application of an axial magnetic field to an MPD thruster increases the thrust and focuses the plasma jet. The goal of this paper is to obtain a simple and ready-to-use theory to predict the thrust generated by an applied-field MPD. To derive a simple thrust formula some assumptions have been made: the applied field is assumed to be purely axial inside the thruster and the total current, plasma temperature, mass flow rate and axial velocity at the exit section of anode to be not sensibly influenced by the application of the magnetic field. Using the magnetic stress tensor the forces acting on the plasma inside the thruster can be derived. Assuming that the detachment from the magnetic nozzle outside the MPD happens when the charged particle motions violate the adiabatic condition the thrust expression can be finally derived. Comparison with experimental data confirms these hypotheses and gives good agreement with theory.

A thrust balance for low power hollow cathode thrusters

A hanging thrust balance has been designed, manufactured and tested at the University of Southampton. The current design allows for direct steady thrust measurements ranging from 0.1 to 3 mN but this can be easily extended to measure thrust in a different range. Moreover the chosen balance design and the thrust measurement procedure allow for the cancellation of thermal drifts. The thrust balance was tested with a T6 hollow cathode thruster providing measurements with an uncertainty of about 9.7%. The thrust data were compared to those obtained with another direct thrust balance and they are in quantitative agreement?the maximum difference being only 6%.

The Applicability of Dual Stage Ion Optics to Ion Engines for High Power Missions

In this paper, the applicability of dual stage ion optics and in particular of the so-called dual stage ion engine to high power, high specific impulse missions will be evaluated. First, the performance limits of conventional two gridded ion engines (GIE) will be discussed and the advantages provided by dual stage ion engines reported. The limits of applicability of a dual stage ion engine will be analyzed analytically and the results confirmed numerically. The lifetime and performance of a three gridded dual stage ion engine (DS3G) will be numerically investigated and compared to those of a conventional GIE assessing for the first time in the open literature under what condition dual stage ion optics provide performance improvements over conventional GIEs and what is its impact on the thruster lifetime. Dual stage ion engines have been found to be capable of providing higher thrust density and longer lifetime with respect to conventional gridded ion engines.