Technische Universität Berlin(ISF-VAT) with the RPA found an average half-angle of divergence of...

Technische Universität Berlin

Institut für Luft- und Raumfahrttechnik

Fakultät für Verkehrs- und MaschinensystemeMarchstrasse 12-14

10587 Berlinhttp://www.ilr.tu-berlin.de

Master’s Thesis

Characterization of Micro Plasma ThrusterPlumes Using Retarding Potential Analyzer

Alexander Raoul Linossier

Matriculation Number: 38056913.05.2018

Primary Thesis AdvisorAssist. Prof. Igal Kronhaus

Faculty of Aerospace EngineeringTechnion - Israel Institute of Technology, Haifa, Israel

Secondary Thesis AdvisorDr.-Ing. Zizung Yoon

Asher Space Research InstituteTechnion City, Haifa, 32000

Israel

This dissertation was undertaken in the Aerospace Plasma Laboratory as part of theAsher Space Research Institute, Technion - Israel Institute of Technology, in Haifa, Israel.

I would like to thank Assist. Prof. Igal Kronhaus for his invaluable supervision anddirection in the lab at the Technion, as well as Dr. Matteo Laterza for his advice andassistance. Thanks must also go to Maxim Rubanovych for his assistance in setting upand operating the Hall thruster experiments, as well as my colleagues Anton, Gianpaolo,and Omri who made ASRI a welcome and fun place to be.I would also like to thank Dr.-Ing. Zizung Yoon, my supervisor at TU Berlin, and Dipl.-

Ing Cem Avsar and Olga Homokova, Master of Space Engineering Programme Managers,for their support and assistance over the last two years.Finally, toda raba, vielen Dank, and thank you to my parents, family and friends for

their never-ending support in Israel, Germany, and back home in Australia.

Declaration of Authorship

I hereby certify that this thesis has been composed by me and is based on my ownwork, unless stated otherwise. No other person’s work has been used without dueacknowledgement in this thesis. All references and verbatim extracts have beenquoted, and all sources of information, including graphs and data sets, have beenspecifically acknowledged.

..................................................................................................................................Place, Date Signature - Alexander Linossier

Agreement on Rights of Utilization

The Technische Universität Berlin, represented by the Chair of Space Tech-nology, may use the results of the thesis at hand in education and research. Itreceives exclusive rights of utilization as according to § 31 Abs. 2 Urheberrechts-gesetzt (Urhg). This right of utilization is unlimited and involves content of anykind (e.g. documentation, presentations, animations, photos, videos, equipment,parts, procedures, designs, drawings, software including source code and similar).

..................................................................................................................................Place, Date Signature - Professor Dr.-Ing. Klaus Brieß

Head of the Chair of Space Technology

Abstract

Plasma thrusters, such as the Hall Effect Thruster (HET), have been used in the pastfor both precise orbital and attitude control manoeuvres, as well as the primary propul-sion system for deep space missions. Rather than relying on chemical combustion toaccelerate a gaseous propellant, plasma thrusters utilize electrical discharges to produceand accelerate an ionized gas to generate thrust. Plasma thrusters typically have higherspecific impulses than chemical thrusters, and are therefore significantly more fuel effi-cient. In past years, there has been increasing interest in developing plasma thrusters fornanosatellites, also known as CubeSats.The plasma thruster plume is dependent on the operating behaviours and plasma gen-

eration conditions of the thruster. By analyzing the ion current density and energy distri-bution, the internal operation of the thruster is revealed, as well as the macro-performancecharacteristics. In addition, the thruster plume can interact with and damage spacecraftsurfaces, and so the distribution of the plume must be well characterized.The exhaust plume of two micro electric thrusters was analyzed using a retarding

potential analyzer (RPA) and biased planar probe via remotely operated mechanicalpositioning systems. Far field scans of the Inline-Screw-Feeding Vacuum Arc Thruster(ISF-VAT) with the RPA found an average half-angle of divergence of 54.7o, with amaximum total ion current of 0.87 A. The thrust calculated from RPA measurementswas 3.54 µN, in good agreement with the 4.44 µN measured by a thrust balance. Thesystem total efficiency was 3.16% at an input power of 1.5 W.Near field scans of the Narrow Channel Hall Thruster (NCHT) were conducted with

the biased planar probe. Significant asymmetry in the thruster plume was observed. Theplume is fully formed within 20 mm of the channel exit. Thrust was calculated from theplanar probe data as 0.53 mN to 2.90 mN, matching the thrust balance value of 0.94mN. Far field scans were performed with both the RPA and planar probe for calibrationpurposes, at input powers of 17.3 W and 25.5 W. The plume contains a main beambetween ±20o, with a half-angle of 71o. The 25.5 W mode produced more thrust at0.73 mN compared to 0.51 mN for the 17.3 W mode, with a total efficiency of 7%. Thegeometry of the plume was not affected by either input power or propellant flow rate.

Contents

List of Figures xi

List of Tables xv

Nomenclature xvii

1 Introduction 11.1 Fundamental Concepts of Plasma Plumes and Ion Diagnostics . . . . . . . 3

1.1.1 EP System Propellant Mass Consumption . . . . . . . . . . . . . . 31.1.2 Ion Exhaust Thrust Calculation . . . . . . . . . . . . . . . . . . . . 31.1.3 EP System Efficiencies . . . . . . . . . . . . . . . . . . . . . . . . . 41.1.4 Diagnostics of Ions in the Plume . . . . . . . . . . . . . . . . . . . 51.1.5 Ion Energy Distribution in the Plasma Plume . . . . . . . . . . . . 6

1.2 Electrostatic Probes for Plume Ion Diagnostics . . . . . . . . . . . . . . . 71.2.1 Biased Planar Probe . . . . . . . . . . . . . . . . . . . . . . . . . . 81.2.2 Retarding Potential Analyzer (RPA) . . . . . . . . . . . . . . . . . 8

1.3 Inline-Screw-Feeding Vacuum Arc Thruster (ISF-VAT) . . . . . . . . . . . 111.4 Narrow Channel Hall Thruster (NCHT) . . . . . . . . . . . . . . . . . . . 141.5 Purpose of Thesis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171.6 Scope of Thesis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171.7 Outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

2 Inline-Screw-Feeding Vacuum-Arc Thruster (ISF-VAT) Experimental Char-acterization 192.1 RPA Experimental Setup & Method . . . . . . . . . . . . . . . . . . . . . 19

2.1.1 Vacuum Facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192.1.2 Semion Single RPA . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

2.2 RPA Results and Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . 252.2.1 Ion Energy Distribution Function . . . . . . . . . . . . . . . . . . . 25

2.2.2 Ion Current and Half-Angle Plume Divergence . . . . . . . . . . . 282.2.3 Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 312.2.4 Thruster Specific Impulse and Efficiency . . . . . . . . . . . . . . . 31

2.3 Thrust Balance Experimental Setup & Method . . . . . . . . . . . . . . . 322.4 Thrust Balance Results and Discussion . . . . . . . . . . . . . . . . . . . . 32

3 Narrow Channel Hall Thruster (NCHT) Experimental Characterization 353.1 Near Field Experimental Setup & Method . . . . . . . . . . . . . . . . . . 35

3.1.1 Vacuum Facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353.1.2 Probe Setup, Data Acquisition & Measurement Procedure . . . . . 36

3.2 Near Field Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 413.2.1 Gas Distributor Modification . . . . . . . . . . . . . . . . . . . . . 413.2.2 Effect of Flow Rate . . . . . . . . . . . . . . . . . . . . . . . . . . . 433.2.3 Total Ion Current & Thrust Estimation . . . . . . . . . . . . . . . 45

3.3 Far Field Experimental Setup & Method . . . . . . . . . . . . . . . . . . . 463.4 Far Field Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

3.4.1 Measured Ion Energy Distribution Function . . . . . . . . . . . . . 503.4.2 RPA Measured Ion Current and Half-Angle Plume Divergence . . . 553.4.3 Thrust Measurements and Specific Impulse . . . . . . . . . . . . . 573.4.4 Thruster Efficiencies . . . . . . . . . . . . . . . . . . . . . . . . . . 573.4.5 Comparison to NASA-173Mv2 Hall Thruster . . . . . . . . . . . . 57

4 Conclusion 594.1 Summary and Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

Bibliography 61

Appendices 65

A. NCHT Magnetic Field Testing 67

B. NCHT Far Field Full Angle Charts 87

List of Figures

1.1 Example distribution of ion current density, with angle from thruster cen-treline on the horizontal axis. Note the ’wings’ at higher angles caused bycharge-exchange events. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

1.2 Example ion energy distribution showing main beam and charge exchangeion peaks. The most probable beam ion energy is marked by a red line,with the most probable charge exchange ion energy much lower at the greenline. The dashed black line demonstrates the full-width half maximum(FWHM) value. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

1.3 Schematic of a 4-grid Retarding Potential Analyzer (RPA) probe. Thecontrol unit can vary the biases applied to each of the grids. . . . . . . . . 9

1.4 Example I-V data from a scan of the NCHT at a radial distance of 70mm. Only positive positions are shown for clarity. Each line shows thecharacteristic initial plateau, S-curve degradation as the repelling bias isincreased, until no ion current is measured. At this point, no ions havehigh enough energy to overcome the repelling bias on G2. . . . . . . . . . 10

1.5 Schematic of the ISF-VAT (top), and the ISF-VAT firing (bottom) [1] . . 121.6 Schematic of a typical Hall thruster. . . . . . . . . . . . . . . . . . . . . . 141.7 Schematic cross-section of the NCHT (left) and an image of the NCHT

operating (right) [2]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

2.1 Schematic of the experimental setup for the smaller vacuum facility, withthe thruster mounted for plume measurements, and not on the thrustbalance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

2.2 CAD drawing of the far field RPA positioning mechanism for the VATscans. The VAT is shown in red, RPA in brown, and the Zaber T-LSM200B-SV2 200 mm linear stage in blue. The arm axle is locateddirectly underneath the tip of the VAT cathode. Moving the linear stagerotates the RPA around this point at a constant radius. . . . . . . . . . . 20

Characterization of Micro Plasma Thruster Plumes Using Retarding Potential Analyzer

2.3 Schematic of the measurement setup for the ISF-VAT far field scans. . . . 21

2.4 Image of the Semion RPA button probe used [3]. . . . . . . . . . . . . . . 21

2.5 Data from a single RPA scan taken at 0o. This scan is comprised ofa number of voltage sweeps over ≈ 2000 pulses and shows the averagecase. The red line shows the smoothed current trace using a 25-pointquadratic/cubic Savitzky-Golay smoothing function. . . . . . . . . . . . . 25

2.6 Smoothed IV traces at 0o over time. The same profile is visible for alltimes, which implies that the ion generation process remains the sameover time. As the arc progresses and the discharge current drops, themagnitude of the ion current decreases in proportion. . . . . . . . . . . . . 26

2.7 IEDFs derived from the IV traces in Fig. 2.6. While there is some varia-tion in the distribution, it remains relatively constant in shape and onlydegrades in magnitude over time. . . . . . . . . . . . . . . . . . . . . . . . 27

2.8 IEDFs at all measured angles. Note the shift in most probable ion energyfrom 42 eV to 34 eV at angles 45o and above. . . . . . . . . . . . . . . . . 28

2.9 Ion Current Density over time according to angle. The markers correspondto measurement points. The 25% error shown on the t = 20 µs case appliesto all measurements. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

2.10 Comparison of discharge and ion current temporal profiles. The dischargecurrent traces are from individual pulses and are intended to demonstratethe variation of the thruster. The ion current points are the average of ≈2000 pulses. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

2.11 Normalized cumulative integral of ion current from 0o to 90o. The red lineindicates the angle within which 95 % of the total ion current is contained. 30

2.12 Drift compensated thrust balance data. The data were adjusted so thatthe thruster ’off’ periods produced an average zero thrust. . . . . . . . . . 33

2.13 Thrust data sampled every 3000 pulses. Note the higher thrust producedduring the initial firing with applied graphite coating over the insulator. . 34

3.1 Schematic of the general experimental setup for the larger vacuum facility. 36

3.2 Photo of the near field scanning setup for the planar probe. The linearstages controlling axial and radial position are at the top, with the planarprobe holder extending down into the plume. The planar probe is locatedat the end of the thin ceramic tube at the end of the probe holder. . . . . 37

xii Master’s Thesis, ILR, TU Berlin, 2018


3.3 CAD drawing of the near field probe positioning mechanism. When onlythe planar probe is being used, the smaller linear stage and RPA holderare removed, and the planar probe holder is fixed to the stage arm. Thismaintains the probe facing parallel to the thruster axis. . . . . . . . . . . 38

3.4 Schematic of a generic biased planar probe . . . . . . . . . . . . . . . . . . 393.5 Circuit diagram showing the electrical connections between the experi-

mental equipment. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 403.6 Ion current density as measured in the near field of the unmodified NCHT.

The channel exits are shown in black on the left. Note the significantdisparity between the upper and lower channel exits. . . . . . . . . . . . . 41

3.7 Visual comparison of NCHT plumes before (left) and after (right) gasdistributor modification using the same camera settings and lighting con-ditions. The plasma probes scanned the plane perpendicular to the onepictured. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

3.8 A sample axial scan taken at 0 mm radial position. The raw data is inblue, with the 150 moving point average shown in red. . . . . . . . . . . . 43

3.9 Near field ICD for the four flow rates tested: 0.32 mg/s (top left), 0.34mg/s (top right), 0.40 mg/s (bottom left), 0.50 mg/s (bottom right). Thechannel exits are depicted in black on the left of each plot. . . . . . . . . . 44

3.10 High resolution scan of upper, high intensity channel exit . . . . . . . . . 453.11 High resolution scan of lower, low intensity channel exit . . . . . . . . . . 463.12 CAD drawing of the far field probe positioning mechanism used to sweep

the RPA through the plume. The RPA is shown in brown, and the pulleymass in black. The biased planar probe is mounted to the negative angleside of the RPA. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

3.13 Schematic showing the definitions of angles for the far field scans. Thethruster is in the same orientation as for the near field scans. . . . . . . . 48

3.14 Ion current measured by the planar probe along the centreline of thethruster, with an inverse square model fitted and extrapolated to 120 mm. 49

3.15 IV data from a single RPA scan taken at setpoint 2 and 0o. More datasamples can be averaged for each measurement point than in the time-resolved mode used for the ISF-VAT, improving the SNR. The 17-pointSavitzky-Golay smoothing function preserves the shape of the data well. . 51

3.16 IEDF smoothed with a 17-point Savitzky-Golay function calculated fromthe data in Fig. 3.15 with the thruster in setpoint 2 at 0o. . . . . . . . . . 51

Master’s Thesis, ILR, TU Berlin, 2018 xiii

3.17 IEDFs within the main beam with smoothing functions applied for setpoint 1. 523.18 IEDFs within the main beam with smoothing functions applied for setpoint 2. 523.19 Full Width Half Maximum for the IEDFs between ±40o and both thruster

setpoints. A beam is clearly visible between -20o and 20o. . . . . . . . . . 533.20 Ion current density angular distribution as measured by the RPA and

biased planar probe. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 543.21 Ion current density at all angles and both power modes. The main beam

is seen as the ICD peak between -20o and 20o. . . . . . . . . . . . . . . . . 543.22 Normalized integrated ion current for setpoint 2. The blue line indicates

the 95 % value, and the red line the half-angle. . . . . . . . . . . . . . . . 563.23 Normalized integrated ion current for setpoint 1. The blue line indicates

the 95 % value, and the red line the half-angle. . . . . . . . . . . . . . . . 56

List of Tables

2.1 Semion RPA voltage biases . . . . . . . . . . . . . . . . . . . . . . . . . . 22

3.1 NCHT operating conditions for near field scans . . . . . . . . . . . . . . . 413.2 Semion RPA voltage biases . . . . . . . . . . . . . . . . . . . . . . . . . . 493.3 NCHT operating conditions for far field scans . . . . . . . . . . . . . . . . 503.4 Total ion current measured in the far field . . . . . . . . . . . . . . . . . . 553.5 Total ion current measured in the far field [4] . . . . . . . . . . . . . . . . 58


xvi Master’s Thesis, ILR, TU Berlin, 2018

Nomenclature

α Charge state ratio correctionA Retarding Potential Analyzer (RPA) total orifice area [m2]Ap Planar probe area [m2]B Magnetic field vector [T]B Magnetic field strength [T]e Unit charge 1.60 × 10−19 [C]E Electric field vector [V/m]f(V ) Ion Energy Distribution Function (IEDF)γ Plume divergence thrust correctiong Gravitational acceleration 9.81 [ms−2]I+ Singly charge ion current [A]I++ Doubly charge ion current [A]Ib Beam current [I]Ii,ax Ion current parallel to thruster axis [A]Iic RPA collector current [A]Ii,tot Total ion current [A]Ip Planar probe current [A]Isp Specific impulse [s]Impp Average impulse per pulse [Ns]Imptot Total impulse [Ns]ji,c Ion current density at RPA [A/m2]ji,sat Ion current density at planar probe saturation [A/m2]Mi Ion mass [kg]mi Ion mass flow rate [kgs−1]mp Propellant mass [kg]mp Propellant mass flow rate [kgs−1]ms System dry mass [kg]ηe System electrical efficiency


ηm Mass utilization efficiencyηT Total system efficiencyni Ion number density [m−3]Np Number of pulsesφ Angle from thruster main axis [o]Pin Input power [W]Pjet Jet power [W]qi Ion charge staterL Larmour radius [m]r1 Inner radius scan edge [m]r2 Outer radius scan edge [m]θ Half-angle of plume divergence [o]t Time [s]T Thrust [N]Tg RPA grid transmission factor∆v Delta-v [ms−1]v⊥ Particle velocity perpendicular to magnetic field [ms−1]Vb Beam voltage [V]vex Exhaust velocity [ms−1]vi Ion velocity [ms−1]Z Charge state

ADC Analog-Digital ConverterAPB Amplified Piezoelectric BrakeAPL Aerospace Plasma LaboratoryASRI Asher Space Research InstituteCAD Computer Aided DesignCEX Charge ExchangeDAQ Data AcquisitionEP Electric PropulsionFEMM Finite Element Method MagneticsFWHM Full-Width Half MaximumHET Hall Effect ThrusterICD Ion Current Density

xviii Master’s Thesis, ILR, TU Berlin, 2018


IEDF Ion Energy Distrbution FunctionIGBT Insulated Gate Bipolar TransistorISF-VAT Inline-Screw-Feeding Vacuum Arc ThrusterISS International Space StationLEO Low Earth OrbitMFC Mass Flow ControllerNCHT Narrow Channel Hall ThrusterPPU Power Processing UnitRF Radio-frequencyRPA Retarding Potential AnalyzerSCU System Control UnitSPT Stationary Plasma ThrusterTAL Thruster with Anode LayerTTL Transistor-transistor LogicVAT Vacuum Arc Thruster

Master’s Thesis, ILR, TU Berlin, 2018 xix

1 Introduction

In the last two decades, there has been a distinct trend of miniaturization in the spacesegment. This has driven down launch and program costs, and allowed more players, suchas universities and research organizations, to enter the space industry [5]. The reductionin the cost and size of each spacecraft has enabled large constellations consisting ofhundreds or thousands of micro- or nanosatellites in Low Earth Orbit (LEO) [6]. Theseconstellations offer the possibility of real time global coverage and data collection. Apropulsion system, however, is required in order to maintain these constellations [6].Miniaturizing the current generation of propulsion systems is a significant challenge.

Electric propulsion (EP) systems are preferable to chemical propulsion systems for useonboard nanosatellites due to their lower propellant consumption. Another advantageof EP systems is that they typically utilize inert and non-hazardous propellants such asXenon or metal.

All electric propulsion systems require electrical power generation and conditioning.An EP system selection is based not only on performance, but also on the availablepower and incremental mass of the power source [7, 8]. Small satellites are particularlylimited in their power generation abilities as they rely exclusively on solar panels, oftenbody-mounted, and hence require low power EP systems. For example, the typical powergenerated by a 1U CubeSat is 2 W, and up to 8 W for a 3U satellite [9]. Dependingon the power processing unit (PPU) and battery specifications, higher power may beavailable for short periods.

Many EP devices utilize a state of matter known as plasma, i.e. ionized gas, consistingof positively charged ions and negatively charged electrons. There is, on average acrossthe plasma volume, an equal amount of positive and negative charges, a condition knownas quasi-neutrality. Due to the difference in charge and mass, these particles can beselectively affected with the application of electric and magnetic fields. Within a plasma,there may be multiple species of ions and electrons that have varying charge states orenergy distributions. A plasma plume is the region downstream of an operating EPdevice containing accelerated ions and electrons under quasi-neutrality. The electrons


may be carried together with the ions or be supplied by an external source [10].There are several types of electric propulsion devices, divided into three major classes

[10,11]:

1. Electrothermal: Propellant is heated using electrical methods, either throughresistive heating or arcing, producing expansion and ejection from the thruster.Electrothermal thrusters do not necessarily produce plasma, and may simply heata propellant such as a liquid to produce higher exhaust velocities without inducingcombustion.

2. Electrostatic: Propellant is ionized using various methods, producing a plasma.The ion species within the plasma are then accelerated using an electric field.Electrostatic thrusters separate the ions and electrons before acceleration, and arenot considered plasma thrusters. Ejected ions must therefore be neutralized toprevent charging of the spacecraft.

3. Electromagnetic: Propellant is ionized using a variety of methods, usually in-volving a magnetic field to control electron movement. Either the entire plasmais accelerated out of the thruster using induced magnetic fields, or ions are ac-celerated using a static electric field, requiring neutralization as for electrostaticthrusters. In contrast to electrostatic thrusters, electromagnetic thrusters operatequasi-neutrally, avoiding limitations from space charge densities.

Within each class, thrusters can be further classified according to the propellant ioniza-tion method and discharge current regime. The term ’plasma thruster’ refers to systemsthat operate with quasi-neutrality maintained throughout [10].Until recently, EP system development focused on high-power systems of 1 kW and

above. With the development of CubeSats and nanosatellites in general, a new area ofresearch has emerged for thrusters with power consumption of less than 100 W, and aslow as 1 W. These thrusters can operate at power levels suitable for nanosatellites, albeitat significantly lower total efficiency [12].Several designs of miniaturized plasma thrusters are studied in the Aerospace Plasma

Laboratory at the Asher Space Research Institute, Technion, in Haifa, Israel. Thiswork focuses on the study of two devices. The first is the Inline-Screw-Feeding VacuumArc Thruster (ISF-VAT), designed for CubeSat applications down to 1U in size, andconsuming less than 1 W total power [1]. Four ISF-VATs will be demonstrated on-boardthe 2U DriveSat satellite for attitude control and apogee raising manoeuvres [13]. The

2 Master’s Thesis, ILR, TU Berlin, 2018


second is the Narrow Channel Hall Effect Thruster (NCHT), a sub-50 W Hall thrusterwith novel channel design developed for CubeSats larger than 3U [2].

1.1 Fundamental Concepts of Plasma Plumes and Ion

Diagnostics

1.1.1 EP System Propellant Mass Consumption

The ideal rocket equation is:

∆v = Ispg lnmp +ms

ms(1.1)

where g is the acceleration due to gravity (≈ 9.81ms−2), mp is the mass of propellantexpelled, and ms is the dry mass of the spacecraft. The main advantage of EP systemsover chemical is revealed in the exponential relationship between the Isp and mp for arequired ∆v. Rearranging Eq. 1.1 gives:

mp = ms(e∆vIspg − 1) (1.2)

It can be seen from Eq. 1.2 that increasing the specific impulse by an order of magnitudehas a dramatic effect on the required mass of propellant to provide the same ∆v.

1.1.2 Ion Exhaust Thrust Calculation

In most EP systems, thrust comes from the acceleration of ions. Thrust is calculated asthe momentum of the particles exhausted over time (following [10]):

T = mpvex (1.3)

where mp is the propellant mass flow rate, defined as the particle number flow ratemultiplied by the particle mass, and vex is the particle exhaust velocity. Given the massof ions is much larger than that of electrons, and the exhaust velocity of accelerated ionsgreatly exceeds that of any neutral particles leaving the thruster, Eq. 1.3 becomes:

T ≈ mivi (1.4)

Master’s Thesis, ILR, TU Berlin, 2018 3


where the subscript i denotes the ion species. The mass flow rate of ions is proportionalto the ion beam current, Ib, with average charge state, Z, by:

mi =IbM

Ze(1.5)

The ion velocity can be determined from the acceleration voltage, Vb, as:

vi =

√2ZeVbM

(1.6)

Determining the angular distribution of ion energy and current is required to accuratelycalculate the thrust. Substituting Eqs. 1.5 and 1.6 into Eq. 1.4, the thrust can becalculated from the measured ion current in the exhaust by:

T = Ib

√2MVbZe

(1.7)

1.1.3 EP System Efficiencies

The total system efficiency of an EP system relates the kinetic energy of the exhaust tothe input power of the system (i.e., the useful output of thrust and momentum to therequired input of electrical power). The kinetic energy of the exhaust, known as the jetpower, is:

Pjet =1

2mpv

2ex =

1

2miv

2i (1.8)

Relating this back to the thrust using Eq. 1.4:

Pjet =T 2

2mi(1.9)

The total system efficiency is then:

ηT =PjetPin

=T 2

2miPin. (1.10)

Here we assume that the thrust is generated by accelerated ions, and a negligible amountby exhausting neutrals and electrons.In this case, the best performance will be obtained when 100 % of the propellant is

singly ionized. The mass utilization efficiency is calculated as:

ηm =mi

mp=IbM

emp(1.11)



assuming only singly-charged ions are present.

Finally, there is also the electrical efficiency of the EP system. For singly-charged ions,the system electrical efficiency is:

ηe =PoutPin

=IbVbPin

(1.12)

1.1.4 Diagnostics of Ions in the Plume

Plume ion diagnostics enable evaluation of the thruster performance as well as obtainsome insight regarding the internal processes in the thruster. Plume characterizationcan also be used to consider damaging plume interaction with spacecraft componentsand structures [14,15].

Figure 1.1: Example distribution of ion current density, with angle from thruster cen-treline on the horizontal axis. Note the ’wings’ at higher angles caused bycharge-exchange events.

Fig. 1.1 shows an example angular distribution of ion current density. The total ioncurrent emitted by the thruster, Ii,tot is calculated by considering the plume to be a radialexpansion from the centre of the thruster emission plane. Assuming an axisymmetricplume, the integral at a specific time simplifies to:

0.95 =1

Ii,tot

∫ θ

02πr2sin(φ)ICD(φ)dφ (1.13)



where r is the radius of the hemisphere, φ is the azimuthal angle, and ICD is the ioncurrent density as a function of angle. The half-angle, θ, is defined here as the angle thatcontains 95 % of the total ion current [4, 16].A thrust correction term is typically introduced that takes both the plume divergence

and ion charge states into account:

γ = αcosθ (1.14)

where α is related to the charge state ratio. Considering single and doubly charged ionsthen:

α =I+ + 1√

2I++

I+ + I++(1.15)

with I+ and I++ the single and doubly charged ion currents respectively. Applying thesecorrections, Eq. 1.10 becomes:

ηT =γ2ηmIi,totVb

Pin(1.16)

where Ii,tot is the total ion current in all directions, not only the axial component.

1.1.5 Ion Energy Distribution in the Plasma Plume

The Ion Energy Distribution Function (IEDF) describes the spread of ion energies, andhence velocities, present in the exhaust plume. The relationship between ion energy andvelocity is based on the kinetic energy of the particle and described by Eq. 1.6.Particular ion generation and acceleration processes within a thruster will produce

ions with specific energies. For example, assuming a negligible velocity immediatelyafter ionization, ions accelerated by an ion thruster should have energies (in eV) closeto the potential across the acceleration grids [10]. In the case of a Hall thruster, wherethe ionization and acceleration zones overlap, ions may be generated near the start ofthe acceleration zone, experiencing a large acceleration voltage, or near the exit of thethruster channel, experiencing a smaller acceleration voltage. The ions will also undergocollisions with electrons and neutral particles within the acceleration and plume regions.This will lead to a wider spread in ion energies [10,17].Ions impacting with neutral particles in the exhaust, vacuum chamber, or with thruster

components, may produce low-energy charge-exchange ions [17]. Within the plume,the measured CEX ions originate mostly due to interactions with neutral background



particles remaining in the vacuum facility [18]. This effect is expected to increase withdistance from the thruster, as ions have an increased chance to collide with backgroundparticles.

Figure 1.2: Example ion energy distribution showing main beam and charge exchangeion peaks. The most probable beam ion energy is marked by a red line, withthe most probable charge exchange ion energy much lower at the green line.The dashed black line demonstrates the full-width half maximum (FWHM)value.

An illustrative example of a plume IEDF is shown in Fig. 1.2. The full-width halfmaximum (FWHM), a measure of spread, and the most probable ion energy have beenmarked on the plot. A low energy peak associated with charge exchange ions has alsobeen highlighted.

1.2 Electrostatic Probes for Plume Ion Diagnostics

Plume ion characterization has been performed on electric propulsion systems of all typesand power levels [4,16,17,19–22]. Most studies of the plume have focused on measuringthe plume ion divergence, ion energy, and thrust, using electrostatic probes such as theLangmuir probe and Retarding Potential Analyzer (RPA). Thrust has been measureddirectly by mounting the thruster on a thrust balance. These parameters were then usedto infer the conditions inside the thruster. By determining the ion energy distribution,it is possible to determine how ions are generated [18,23].



1.2.1 Biased Planar Probe

The planar probe is a type of Langmuir probe that presents a relatively large collectionarea to the thruster and ion flow, and was used to measure the ion current density [1,24].The probe has a simple construction of a conductive disc with a voltage bias applied. Ifthe plasma potential is applied to the probe, all charged particles in the quasi-neutralplume will be collected, and zero current registered. A positive bias will repel ions and anegative bias will repel electrons. At some negative bias, all electrons are repelled, andonly ions collected. As the bias becomes more negative, effects from the expanding sheathresult in more ions being collected. The sheath expansion effects can be characterizedby sweeping the applied voltage bias. Sweeping the probe bias produces three distinctregions in the voltage-current trace; ion saturation for sufficiently negative bias, electronsaturation for sufficiently positive bias, and an electron retarding region in between. Forion current analyses, the probe should be biased in the ion saturation region [24].The ion current density is calculated in Eq. 1.17 as the measured probe current

divided by the probe collection area, Ap, assuming the probe bias places it within theion saturation region:

Ip/Ap = ji,sat (1.17)

1.2.2 Retarding Potential Analyzer (RPA)

The Retarding Potential Analyzer (RPA) is a diagnostic device that measures the ionenergy distribution within a plasma plume. Its operating principle is similar to that of thebiased planar probe described in 1.2.1, in that electrons are repelled and ions absorbedthrough the application of a voltage bias. RPAs typically consist of either three or fourgrids in front of a collector plate, as illustrated by Fig. 1.3 [21]:

1. Floating aperture grid: This grid is floating or grounded to the probe body, andis used to reduce the sampling orifice diameter to below the plasma’s Debye length.

2. Electron repelling grid: A negative bias applied to this grid prevents electronsfrom entering the RPA. It performs the same role as the negative bias applied tothe planar Langmuir probe, and must also be sufficiently negative relative to theplasma potential to repel nearly all electrons.

3. Ion retarding grid: A positive bias sweep applied to this grid will selectivelyrepel ions, relative to the plasma potential. Only ions with energies greater than



the applied bias will be able to pass through. Importantly, this grid differentiatesions based on their energy per unit charge, and not total kinetic energy. That is,a Xe+ ion with 50 eV energy will register the same as a Xe+2 ion with 100 eVenergy. Species dependent effects cannot therefore be resolved without additionalinformation regarding the ratio of species in the plasma plume.

4. Secondary electron repelling grid (optional): Ions impacting the collector platewith sufficient energy may result in the emission of a secondary electron. The sec-ondary electron repelling grid is biased negative relative to the collector, preventingsecondary electrons from escaping the RPA and increasing the apparent ion current.Instead, they return to the collector.

Figure 1.3: Schematic of a 4-grid Retarding Potential Analyzer (RPA) probe. The controlunit can vary the biases applied to each of the grids.

The collector plate is biased slightly negatively relative to the probe body or plasmapotential if the probe body is floating to attract the remaining ions. A typical voltage-current curve is shown in Fig. 1.4, with the ’voltage’ referring to the applied bias on theion retarding grid. At low retarding voltages, all or nearly all ions arrive at the collector,producing the initial current plateau. As the retarding voltage is increased, fewer andfewer ions are able to overcome the potential barrier, until the collected ion current drops



to 0 A.

Figure 1.4: Example I-V data from a scan of the NCHT at a radial distance of 70 mm.Only positive positions are shown for clarity. Each line shows the charac-teristic initial plateau, S-curve degradation as the repelling bias is increased,until no ion current is measured. At this point, no ions have high enoughenergy to overcome the repelling bias on G2.

Following [23], the Ion Energy Distribution Function (IEDF), f(V ), is related to themeasured current-voltage curve by:

dI

dV= −q

2i e

2niAcMi

f(V ) (1.18)

Taking the derivative of the measured I-V curve then provides, in arbitrary units,the ion energy distribution. If the information is available, the IEDF for each speciescan be calculated by taking the proportional value of the IEDF magnitude for a givenspecies, and multiplying the ion energy axis by the specie’s charge state, qi [25]. Themost probable ion energy is the energy at which the peak of the derived curve occurs,and indicates the typical velocity of an accelerated ion.



1.3 Inline-Screw-Feeding Vacuum Arc Thruster (ISF-VAT)

Vacuum Arc Thrusters (VAT) are miniature, low power pulsed plasma electrothermalthrusters. Their small form factor and low power operation make them ideal for station-keeping, constellation maintenance and attitude control applications onboard nanosatel-lites [7,11,26]. VATs provide a solution to the problem of safe, robust and inert propulsionmethods for nanosatellites, which have tight restrictions on mass, physical dimensions,and available power. With power requirements less than 1 W, VATs can be operatedonboard 1U CubeSats with only body-mounted solar panels for up to 15 minutes perorbit [27].

VATs consist of four major components:

• Cathode (also serves as the propellant)

• Anode

• Insulator with an applied conducting layer

• Driving Electronics

A large potential on the order of 1 kV is generated by the driving electronics across theelectrodes. The high potential difference leads to a breakdown of the conducting layer onthe insulator, generating an initial localized cloud of plasma. As these breakdown prod-ucts expand, they provide a conducting pathway between the electrodes under vacuumconditions, resulting in the formation of an arc in less than 1 µs. The arc occurs betweenthe anode and a small region on the cathode [28].

At the cathode surface, the arc current is concentrated into a region on the order of 10µm in diameter, known as a cathode spot. The concentration of the discharge energy in asmall area causes rapid heating of the cathode surface material, resulting in disassociationof the material and generation of quasi-neutral plasma [29, 30]. This process is fast andis considered to be explosive; the expansion of the plasma away from the surface ofthe cathode generates the thrust [30]. The application of an external magnetic field todirect the plasma products (i.e., a magnetic nozzle) can produce higher thrust at thecost of increased power consumption by collimating the exhausted ions [28, 31, 32]. Theexplosive process also results in the ejection of solid cathode material particles, known asmacroparticles, although their relatively low velocity compared to the plasma productsmeans a negligible contribution to total thrust produced. The magnitude of the ion



current is affected by the discharge current of the arc, however the shape of the IEDFwill remain unaffected [19].

As the cathode is consumed with each pulse, it recedes with respect to the anode andinsulator. The energy of the arc is reduced as the gap between the electrodes increases.Eventually, the cathode will be too far inside the insulator for an arc to form. There mayalso be an offset in the construction or assembly of the system, leading to preferentialconsumption of the cathode on one side. Advancing the cathode to compensate willtherefore extend the thruster life [11].

Figure 1.5: Schematic of the ISF-VAT (top), and the ISF-VAT firing (bottom) [1]

The Aerospace Plasma Laboratory at the Asher Space Research Institute has devel-oped a novel VAT with an automatic propellant feeding system to extend the operational



lifetime of the thruster (Fig. 1.5. The Inline-Screw-Feeding Vacuum Arc Thruster (ISF-VAT) has dimensions of 15 x 15 x 65 mm3 and a mass of 60 g, with the Power ProcessingUnit (PPU) adding 100 g to the total system mass. The propellant used for the experi-ments in this thesis is a 0.7 mm diameter pure titanium rod located in the centre of thethruster, mounted within the feeding mechanism. Other metallic materials, includingcopper, tungsten, and steel have also been used as propellant.

An inductive energy storage PPU employing an Insulated Gate Bipolar Transistor(IGBT) to interrupt current flow during charging, generates 1 kV potential between theanode and cathode across the insulator. Initially, the conducting coating on the insulatoris an applied layer of graphite, with the graphite replaced by deposited cathodic materialduring subsequent firings, reaching a steady-state condition after a number of pulses.Depending on the initial charging parameters, the inductor in the PPU will discharge overapproximately 200 µs, with the pulse characterized by the peak discharge current, lengthof pulse, and total pulse energy. After the discharge, the IGBT switches the chargingcircuit on, and the inductor is recharged for the next pulse. For space applications, thethruster is typically operated with a pulse frequency of 20 to 30 Hz, although higherfrequencies are possible. While each pulse will remain largely the same, a higher pulsefrequency will increase the average thrust produced by the thruster, and hence also reducethe operational lifetime as limited by the amount of propellant available.

Figure 1.5 shows the screw based feeding mechanism driven by a torsional spring. Therotational motion is controlled by an amplified piezoelectric brake (APB), and monitoredvia an optical encoder with 4 pulses per revolution. To compensate for the erosion of thecathode, the ISF-VAT uses torque generated by the torsion spring to rotate the cathodewithin a screw thread, both advancing the cathode tip and rotating it relative to theanode. This not only maintains a near-optimal cathode tip position, but also reduces theeffect of azimuthal discharge direction bias due to asymmetry in the thruster manufactureor assembly. Both of these greatly extend the life and performance of the ISF-VAT, withthe erosion of the insulator becoming the limiting factor on the thruster lifetime [1].

The ISF-VAT can operate with input voltages as low as 4 V, with a step-up DC/DCconverter delivering 15 - 35 V to the inductor charging circuit. Typical discharge currentsare on the order of 45 A. The propellant rod is fed in a discrete manner after a specifiednumber of pulses have been completed, followed by a cooldown period of a few minutes.The APB is then activated to advance the propellant rod by a pre-calculated amount,typically 1/8 of a turn per firing period. Higher discharge power will result in an increasedconsumption rate of the cathode rod, and a lower number of pulses before needing to



advance the cathode. The simple mechanical action with few moving parts of the ISF-VAT in particular is robust and reliable. Tests of the ISF-VAT have completed in excessof 700,000 pulses with a single cathode [1]. The ISF-VAT plume is investigated in thiswork.

1.4 Narrow Channel Hall Thruster (NCHT)

Figure 1.6: Schematic of a typical Hall thruster.

Hall thrusters are the most successful electric propulsion devices to date, with hundredslaunched over the last 50 years [4]. Figure 1.6 shows the basic structure of a Hall thruster.A neutral gas, typically Xenon, is fed into an annular channel via the anode [10]. Acathode external to the channel, either side or center mounted, supplies the thrusterwith an electron source. The shape of the channel and placement of the cathode andanode produce an electric field aligned with the thruster axis of symmetry. This electricfield is directed so that electrons are drawn into the channel towards the anode. Magneticcoils in the body of the thruster establish a radial magnetic field perpendicular to theelectric field within the channel.



Electrons entering the channel experience an E×B drift, rotating around the radialmagnetic field lines rather than propagating directly to the anode. The radius of gyrationis defined by [10]:

rL =Mv⊥|q|B

(1.19)

where rL is the radius of gyration, or Larmor radius, M is the mass of the particle,v⊥ is the particle velocity perpendicular to the magnetic field, q is the charge on theparticle, and B is the strength of the magnetic field. Through a combination of collisionalprocesses, the electrons make their way to the anode. Xenon atoms are ionized byelectron collision to produce plasma, and maximizing electron dwell time within thechannel increases the degree of ionization. The electric field distribution within thechannel is determined by the reduced mobility of the electrons. By preventing the freeflow of electrons, the electric field is established along the thruster axis between theelectrodes [10].

Once the ions are generated, they are then accelerated by the electric field. The ionsalso experience gyration caused by the magnetic field, however due to their much largermass than electrons, the radius of gyration is much longer than the channel length.

The ions exiting the thruster provide the thrust. In Hall thrusters, the force transferpath is from the ions to the electric field established by the electrons; the electrons thentransfer the force via the magnetic field, to the magnetic coils, and finally to the thrusterstructure and body of the spacecraft.

The plasma generation and acceleration regions of a Hall thruster are mixed, and bothoperate under a condition of quasi-neutrality. This means they are not restricted byspace charge density limits and hence have a high thrust density [33]. Hall thrusters stillrequire a cathode to provide neutralizing electrons.

There are two main types of Hall thrusters. The first, Stationary Plasma Thrusters(SPT), have an insulator layer on the channel walls, and a channel that is long comparedto its width. The dielectric walls generate low energy secondary electrons emissions whenions and electrons collide with the walls. These low energy electrons serve to reduce theelectron temperature in the plasma, reducing electron mobility further and providing alengthened and more gradual acceleration region to the ions [10,34].

In comparison, Thruster with Anode Layer (TAL) type Hall thrusters have metallicwalls biased at cathode potential. The cathode bias repels electrons to reduce collisionswith the walls in the ionization region, maintaining a higher electron temperature. This



causes the acceleration and ionization regions to collapse to a thin layer near to theanode, giving rise to the name [35]. The channel can therefore be significantly shortened.

Figure 1.7: Schematic cross-section of the NCHT (left) and an image of the NCHT op-erating (right) [2].

When scaling down the Hall thruster to low and very low power (<100 W), their per-formance is considerably reduced. Reducing the input power results in a lower ionizationefficiency, reducing system specific impulse and total efficiency [36]. In addition, thrusterssuffer from difficulties establishing adequately strong magnetic fields of the correct topol-ogy, resulting in increased particle collisions with the channel walls as the magnetic fielddoes not properly constrain the particles [4,20]. More collisions increase the wear of thechannel walls, decreasing system life, efficiency and specific impulse [34].

The Narrow Channel Hall Effect Thruster (NCHT) was designed to overcome theselimitations. Figure 1.7 shows the unique structure of the NCHT channel. With a channelcentre radius of 30 mm and channel width of only 1 mm, the NCHT has a channel widthto radius ratio of 1/30. This is in comparison to other Hall thrusters, where the ratio isaround 1/6 [2]. Neutral Xenon gas is fed in through the metallic gas distributor whichalso acts as the anode. The entire metallic structure of the thruster forms part of themagnetic circuit, with the sharp tips of the poles at the channel exit concentrating themagnetic field and producing a high local field strength. With conductive channel walls,the NCHT is closer to the TAL in operation. The slanted walls of the plasma chamberprovide a small volume for plasma generation, increasing the ionization efficiency evenat low power.

The NCHT is able to operate at less than 30 W total input power, and as low as 15W, producing ≈1 mN thrust. It requires only two power supplies; one for the anode, andone for the magnetic coil, and is a simple mechanical design with no moving parts. The



outside of the thruster is coated with a ceramic insulator material to isolate the metallicstructure from attracting electrons and ions.While lifetime testing has yet to be performed for the NCHT, the thruster has operated

for more than 90 hours without failure [2]. The low power consumption and form factoris ideal for small satellites 3U in size and larger, and has not been matched to date ( [20]tested thrusters consuming at a minimum 50 W, although at much higher efficiencies).A hot wire cathode is used for the experiments in this thesis, however future iterationsof the thruster will employ a hollow cathode.

1.5 Purpose of Thesis

Both the ISF-VAT and NCHT are new types of EP devices and were not yet charac-terized in full. This thesis aims to perform the first analysis of the ion content of theplume of both the ISF-VAT and NCHT using a Retarding Potential Analyzer (RPA) inconjunction with a planar probe. While plume characterization has been conducted onlarger thrusters, this is the first application of these plasma measurement techniques tothese thrusters. Both thrusters are too small for detailed internal investigations, and sothe ion content of the plume is used to characterize system performance and collect in-formation on the system operation. The results of this thesis are to be used for a range offuture development activities, including thruster design revisions, satellite system design,and thruster numerical modelling verification.

1.6 Scope of Thesis

This thesis focuses on the experimental characterization of plasma thruster exhaustplumes, specifically the plume ion content such as the angular distribution and ion energydistribution. While this data will be used to further the development of these thrusters,this thesis will not consider the design or operational changes beyond some basic recom-mendations. For the NCHT, an investigation was also conducted into the magnetic fieldtopology as described in 4.1.

The primary objective is characterization of the thruster plumes for:

• Ion current density angular distribution

• Ion energy distribution



1.7 Outline

This thesis is separated into 3 chapters.

Chapter 2 concerns the far field characterization of the ISF-VAT. First, the experi-mental setup, including Semion RPA, is presented. The data collected by the RPA arethen analyzed and discussed.

Chapter 3 focuses on the analysis of the NCHT. The experimental setup, includingbiased planar probe, is described. Results and analysis of near field scans using a planarprobe and far field scans using the RPA alongside the biased planar probe are presentedand discussed.

Chapter 4 summarizes the conclusions of the experimental work.


2 Inline-Screw-Feeding Vacuum-ArcThruster (ISF-VAT) ExperimentalCharacterization

This chapter describes the experimental analysis of the plume from the ISF-VAT usingan RPA probe. The experiment equipment and methodology are described, followed bythe results. The results are then analyzed and discussed.

2.1 RPA Experimental Setup & Method

2.1.1 Vacuum Facility

The Asher Space Research Institute (ASRI) has two vacuum chamber facilities for per-forming thruster tests under space environmental conditions. The smaller facility (Fig.2.1), in the Aerospace Plasma Laboratory (APL), is a cylindrical vacuum chamber 1.2m long and 0.6 m in diameter (≈0.34 m3 internal volume). It is equipped with a PfeifferVacuum HiPace®700 turbopump capable of 685 L/s of N2, backed by a Pfeiffer VacuumDuo 20 M dual-stage rotary vane pump capable of 6.7 L/s of air. This pumping systemis capable of maintaining a chamber pressure below 1.0 × 10−6 mbar during operationof the ISF-VAT. The chamber has pass-throughs allowing for communication to devicesunder vacuum and control of actuators for probe positioning, as well as 6 window portsfor optical observations. Chamber pressure is monitored by a Pfeiffer Vacuuum CompactFullRange™Pirani and cold cathode-type combination gauge, with a measurable range of5× 10−8 to 1000 mbar at ±30 % accuracy.A Zaber T-LSM200B-SV2 200 mm linear stage controlled by a LabView program gave

an RPA angular position range of -5o to 90o, where 0o is defined as the main thrusteraxis. This was achieved using the mechanism shown in Fig. 2.2. The angular accuracywas better than 1o with >99 % reproducibility. The RPA probe was positioned 153 mmfrom the front surface of the ISF-VAT anode (Fig. 2.3. The entire positioning assembly,


Figure 2.1: Schematic of the experimental setup for the smaller vacuum facility, with thethruster mounted for plume measurements, and not on the thrust balance.

Figure 2.2: CAD drawing of the far field RPA positioning mechanism for the VAT scans.The VAT is shown in red, RPA in brown, and the Zaber T-LSM200B-SV2200 mm linear stage in blue. The arm axle is located directly underneath thetip of the VAT cathode. Moving the linear stage rotates the RPA around thispoint at a constant radius.



and RPA body, was mounted and grounded to an internal tray.

Figure 2.3: Schematic of the measurement setup for the ISF-VAT far field scans.

2.1.2 Semion Single RPA

Figure 2.4: Image of the Semion RPA button probe used [3].

For this work, a commercial RPA (Semion Single, Fig. 2.4) was purchased fromImpedans Ltd. The Semion RPA uses a 3 grid configuration, without the secondaryelectron repelling grid (see Section 1.2.2). Each of the grids has a transmission of 50 %,with the front orifice plate containing 37 holes of 800 µm diameter arranged in a 10.4mm wide hexagon. The factor relating measured ion current Iic to ion current density



jic is:

jic =IicATg

[A/m2] (2.1)

where A is the area of the orifice plate open to the plasma:

A = 37× π × 0.00082

4= 1.8598× 10−5m2 (2.2)

and Tg is the total transmission of the grids:

Tg = 0.5× 0.5× 0.5 = 0.125 (2.3)

giving a total factor of:

1

ATg= 4.3015× 105 m−2 (2.4)

The Semion system is operated via a control unit (SCU) external to the vacuumchamber and proprietary software. The SCU is connected to the probe via a vacuumfeed-through mounted on the vacuum chamber that contains RF filters (These filtersare included as the system can be operated for RF generated plasma systems in timeresolved mode). The RF filters were not required for the scans performed in this thesis,however they did affect the scanning performance. A settling delay must be specified inthe software between voltage steps in the G2 voltage sweep. It was found that reducingthis delay below a certain duration affected results as the RF filter circuits, containinglarge capacitors, did not reach a steady state before scanning began. Increasing thesettling delays to 1 s for the first step and 40 ms for the subsequent steps was sufficientto prevent this issue.

Table 2.1: Semion RPA voltage biases

G1 Collector G2 Sweep Start G2 Sweep End G2 Sweep StepV V V V V

-60.0 -80.0 0.0 80.0 1.0

Table 2.1 lists the biases applied to the RPA grids and collector. The SCU contains twomeasurement channels; a high resolution, 10 µA channel, and a lower resolution, 60 µAchannel, however in the configuration for analyzing the VAT, the 60 µA channel was not



available. Both channels have an absolute measurement error of 20 %, with a precisionbetter than 1 %. The 10 µA channel was used for all measurements. The Semion systemis designed to work with both pulsed and steady-state plasma systems, and performsthousands of measurements to calculate an average result. The RPA collects plasmaacross the 10.4 mm wide hexagonal orifice area, equivalent in this arrangement to anangular region of:

arctan

(10.4

153

)= 4o (2.5)

This results in both spatial and temporal averaging of the measurement data. TheRPA offers two scanning modes; time-averaged and time-resolved. Time-averaged is forsteady-state systems, and time-resolved for pulsed or RF systems where the plasma at theRPA varies over time. The temporal averaging can be adjusted using three parameters,plus extra temporal resolution parameters in time-resolved mode:

• Integration Time: Number of voltage and current samples taken at each voltagestep and averaged. This is the fastest method to average results, as it requires theminimum number of settling delays. This was increased to the maximum of 4000.

• Sweep Average: Number of complete sweeps to perform, with the samples av-eraged across all of them. This is the slowest method to average results, with thefull set of settling delays being applied for each sweep. The value was set to theminimum, 1.

• Number of scans to run: Number of complete scans to perform, with the resultsthen averaged together. This is as slow as increasing the Sweep Average, howevercan produce individual scan files if required. The value was adjusted according tothe individual thruster to ensure adequate Signal-to-Noise Ratio (SNR).

• Time-Resolved Only Scan Parameters:

– Resolution: Size of each time step bin. Reducing this requires more scans tobe taken to reduce noise, as the data are only averaged within each bin, andmultiple sweeps are needed at each time step. 20 µs was used, limited by thepulse period of the thruster.

– Acquisition Start: Time after trigger signal is received to begin scans. Set to0 µs.



– Acquisition End: Time after trigger signal is received to stop scans. Set to400 µs.

– Average 1 µs ON/OFF: Option to either collect once per µs or average overeach µs and save the average value. This was set to ON.

A Tektronix TDS1154 Oscilloscope and Pearson coil were used to capture the dischargecurrent profile of individual pulses, with an IDS UI-5240CP-C-HQ 1.3MP networkedcamera with 50 mm focal length lens capturing images of the discharges perpendicularto the anode face. Both the oscilloscope and camera were triggered by a TTL pulsegenerated by the Arduino Nano control board. The positioning system, oscilloscope,and camera were controlled by a LabView program, with the RPA controlled from theproprietary Semion software. The experimental setup is shown in Fig. 2.1.

RPA scans were performed every 15o between 90o and 15o, and every 5o between15o and -5o. Each RPA scan required approximately 2000 pulses to complete at 30 Hzfrequency and 20 µs temporal resolution. The time of flight of the ions to the RPA wascalculated as ≈ 10 µs. Times specified hereafter are in µs after the trigger pulse wasreceived. The thruster performance degrades over time as the cathode is consumed andrecedes into the anode, requiring the cathode position to be reset at the start of eachscan. Hence, the thruster was allowed to complete a full 9000 pulse firing cycle withcooldown period for each scan, resetting the cathode to the same position before thenext scan would be taken over the first 2000 pulses of the next 9000 pulse firing period.

The LabView software captured the discharge current, voltage and trigger signal, aswell as an image from the camera, approximately every 20 seconds during thruster oper-ation. It is important to recognize that the RPA results are averaged over the thousandsof pulses each scan required, while the data from the oscilloscope and camera are forindividual pulses. All data were processed in MATLAB.

There are a number of sources of error in the experimental setup. The angular posi-tional accuracy was ±1o, resulting in a worst case absolute error of 3 %. Radial positionof the RPA was measured to within 1 mm accuracy, producing a potential 1.3 % variationin the measured ion current. The Semion system has an absolute measurement error of20 %. As the Semion system averaged ≈ 2000 pulses, the error due to thruster variationis negligible. A further 1 % error was introduced by the numerical integration process,giving a total error on the ion current measurements of 24.3 % and on the ICD of 25.3%. The current measurements at varying angles have a smaller relative error of 5.3 %due to the precision in the calibration of the Semion voltage and current measurement



systems.

2.2 RPA Results and Discussion

2.2.1 Ion Energy Distribution Function

An example RPA scan taken at 0o and 20 µs is shown in Fig. 2.5. The current-voltagecurve follows the expected profile, with an initial plateau, followed by a slope to 0 A asthe retarding voltage increases. A 25-point quadratic/cubic Savitzky-Golay smoothingfunction was fitted to the raw IV data and then derived to calculate the first IEDF [22,37].Derivation of the raw IV data resulted in exceedingly noisy data due to the increase innoise via the derivation process. The most probable ion energy is 42 eV.

Figure 2.5: Data from a single RPA scan taken at 0o. This scan is comprised of a numberof voltage sweeps over ≈ 2000 pulses and shows the average case. The red lineshows the smoothed current trace using a 25-point quadratic/cubic Savitzky-Golay smoothing function.

As the discharge progresses, the discharge current steadily drops, and the amount ofplasma produced decreases in proportion. In Fig. 2.6, the degradation of the plume at



0o over time can be seen, with each I-V profile captured 20 µs apart. By 200 µs, thepulse is completed and no ions are collected by the RPA. Figure 2.7 shows the smoothedIEDF at each time step until 200 µs. The distribution remains relatively constant inshape over time, and only reduces in magnitude. There is, however, a variation in thepeak ion energy from 42 eV to 28 eV. This is likely due to a combination of variance inthe thruster discharge behaviour and noise in the RPA signal, which can be seen in theraw data in Fig. 2.5.

Figure 2.6: Smoothed IV traces at 0o over time. The same profile is visible for all times,which implies that the ion generation process remains the same over time.As the arc progresses and the discharge current drops, the magnitude of theion current decreases in proportion.

At t = 20 µs, the IEDF at each sampled angle has been plotted in Fig. 2.8. Forthe angles -5o to 30o, the most probable ion energy remains constant at 42 eV. Above30o, however, the most probable energy shifts downwards to 34 eV. This is evidenceof a higher-energy beam with a half angle between 30o and 45o. With a higher ioncurrent and higher ion energies, this beam contributes a large proportion of the thrust.The ions in the expanding plasma are affected by the electron pressure; in particular,the distribution of electrons generates an accelerating electric field [38]. Close to the



Figure 2.7: IEDFs derived from the IV traces in Fig. 2.6. While there is some variationin the distribution, it remains relatively constant in shape and only degradesin magnitude over time.



cathode surface, the plasma particles move directly outward from the cathode spot, untilthey begin to expand into the vacuum after a short distance [30]. Higher energy ionswill deflect less than slower ions under the same accelerating field, resulting in a higherenergy beam closer to the thruster axis.

Figure 2.8: IEDFs at all measured angles. Note the shift in most probable ion energyfrom 42 eV to 34 eV at angles 45o and above.

2.2.2 Ion Current and Half-Angle Plume Divergence

Using Eq. 1.17, the ICD was calculated over time and according to angle from thrustercentreline (Fig. 2.9). As expected, the ion current density decreases over time as thepulse progresses and discharge current degrades. It is interesting to note that the ICDdecreases with increasing angle. The plasma is therefore not expanding uniformly, givinga somewhat directional plume.

The ion current density according to angle is integrated over the hemispherical surfacedefined by 0o to 90o at 153 mm from the thruster anode front face using Eq. 1.13.Numerically, the integral was performed using the trapezoidal method at each timestep,giving a maximum total ion current of 0.87 A at t = 20 µs. The average discharge current



Figure 2.9: Ion Current Density over time according to angle. The markers correspondto measurement points. The 25% error shown on the t = 20 µs case appliesto all measurements.

at t = 20 µs was 11.5 A, giving an ion fraction of the discharge current fi = Ii,tot/Id =

0.076. This is within the expected range of 5-10 % [1].Figure 2.10 shows the degradation of the ion current over time compared to the range

of discharge currents measured. All currents were normalized to the value at t = 0 µs. Itcan be seen that the plume ion current follows the same profile over time as the dischargecurrent. This is further evidence that the ions measured by the RPA are generated almostexclusively by the arcing discharge process.Equation 1.13 was used to calculate the half-angle of plume divergence. Figure 2.11

displays the cumulative value of the integral over the angle 0o to 90o at t = 20 µs,normalized to the total current.The plume is only well defined at the first measurement time. The calculated half-angledivergence was 64.6o at t = 20 µs. In terms of spacecraft protection and shielding ofcritical components, it is important to note that there is a non-negligible ion current atall angles up to 90o. No measurements were taken at angles greater than 90o, howeverdue to observed deposition on the anode surface it is assumed that the ion current dropssharply to zero above 90o.



Figure 2.10: Comparison of discharge and ion current temporal profiles. The dischargecurrent traces are from individual pulses and are intended to demonstratethe variation of the thruster. The ion current points are the average of ≈2000 pulses.

Figure 2.11: Normalized cumulative integral of ion current from 0o to 90o. The red lineindicates the angle within which 95 % of the total ion current is contained.



2.2.3 Thrust

The thrust is directly proportional to the ion current and ion velocity according to Eqs.1.6 and 1.7:

T =Ii,axMivi

Ze[N] (2.6)

where T is the instantaneous thrust and Ii,ax is the axial component of the ion current(i.e. the component that contributes to the thrust, calculated by adding a cosθ term tothe integral in Eq. 1.13). The use of Ii,ax accounts for the plume divergence componentof γ, and α is negligible relative to the measurement error.

The ion velocity is calculated from the average most probable ion velocity across allangles and all timesteps using Eq. 1.6. The average ion energy was 33.6 eV, giving avelocity of 16.6 km/s, in good agreement with the literature [39].

With an average charge state of Z = 2.04, the calculated instantaneous thrust at t =20 µs was 1.46 mN [39]. For pulsed thrusters, the impulse bit is more important thanthe instantaneous thrust value. The impulse bit from a single pulse is:

Impp =

∫ tp

0T (t)dt = 0.12µNs (2.7)

where tp is the pulse duration, in this case 230 µs. This gives an average thrust of 3.54µN at a pulse frequency of 30 Hz.

2.2.4 Thruster Specific Impulse and Efficiency

During the test run, 700,000 pulses were fired. This consumed a total of 12 mg of cathoderod [1]. The total impulse delivered by the 12 mg of cathode was:

Imptot = NpImpp = 0.084Ns (2.8)

The specific impulse is therefore:

Isp =Imptot∆mg0

= 714s (2.9)

The total impulse delivered is also used to calculate the thruster efficiency:

ηT =Imp2tot

2∆mNpεp= 3.16% (2.10)



where εp is the average discharge energy per pulse in Joules (≈ 13 mJ). εp is calculatedas the average integral of the discharge current-time curve multiplied by the averagedischarge voltage of 28 V.

2.3 Thrust Balance Experimental Setup & Method

Once the RPA scans were completed, the RPA and positioning mechanism were removed,and the thruster was mounted onto a FOTEC µN Thrust Balance. This thrust balancecan measure up to 6 mN thrust with an accuracy of ±0.075 µN. The entire thrusterassembly is mounted horizontally on one end of a 70 cm long arm fixed in the center viaa 0.096 Nm/rad rotational spring. The horizontal deflection of the arm is measured bya high precision optical reflection-based transducer, with an absolute error of 1.5 %, anda floor noise level of 0.15 µN. The thrust balance also includes 6 ultra-low friction liquidmetal baths to allow electrical signals and power to be passed to the thruster withoutaffecting thrust measurements. When operating in vacuum, an eddy current brake using aneodymium permanent magnet is used to dampen long period oscillations of the system.An electrostatic force actuator is used to counteract the force of the thruster using aclosed loop control system, maintaining the relative position of the arm. This has theadvantage of not requiring accurate knowledge of the spring constant, which is affectedby loading conditions, ambient temperature, and total displacement. This also preventsindividual pulses from being resolved, however, and so the measured thrust is an averagevalue over the sample period of the thrust balance, approximately 0.5 s. The thrust wasmeasured continuously over multiple firing and cooldown periods.

2.4 Thrust Balance Results and Discussion

In Fig. 2.12, a sample of the thrust balance measurements over the first 6 firing periodsis shown. The raw data were manually adjusted to compensate for long-period drift inthe thrust balance measurements by applying multi-period linear corrections. The datawere then offset to produce zero average thrust during periods when the thruster wasnot firing. This contributes an uncertainty to the thrust values of ±0.25 µN in additionto the ±0.07 µN noise from the thrust balance, totalling ±0.32 µN uncertainty.There are three phenomenon characteristic of the ISF-VAT visible in Fig. 2.12:

• Thrust Variation: As the discharge profile is dependent on the condition ofdeposited material between the cathode and anode, condition of the electrodes,



Figure 2.12: Drift compensated thrust balance data. The data were adjusted so that thethruster ’off’ periods produced an average zero thrust.

and local spot erosion of the cathode tip, as well as how these all change as thedischarge progresses, there is a relatively large variation in the thrust produced.This presents as the noise within each firing period.

• High Initial Thrust: As an adequate coating of deposited material on the insu-lator is only established after a number of pulses, a conducting layer of graphite isinitially applied to the insulator face. This produces a higher discharge current asit is more conductive than the deposited material, and gives the higher thrust seenat the beginning of the first firing period.

• Thrust Degradation: Within each firing period, there is a general trend of re-ducing thrust with increasing number of pulses. As previously mentioned, this isdue to the consumption of the cathode material and recession into the insulator.

The thrust was sampled approximately every 3000 pulses in Fig. 2.13. The average ofthese samples was 4.4 µN, which, with the exception of the initial firing with graphitecoating, remained relatively constant across all 6 firing period. This indicates that thefeeding mechanism is working as intended.



Figure 2.13: Thrust data sampled every 3000 pulses. Note the higher thrust producedduring the initial firing with applied graphite coating over the insulator.

Comparing the thrust balance results and thrust calculated from the RPA data, we seegood agreement between the values of 4.4 µN and 3.54 µN respectively. The differencecan be attributed to a number of factors. First, the measurements of each were taken overdifferent sets of firing periods. Second, the RPA only measures the ions exhausted fromthe thruster, and does not register electrons or neutral particles. Finally, experimentalerror introduces uncertainty to both values. These measurements validate the use ofthe RPA to characterize the performance of plasma thrusters in addition to the iondistribution.


3 Narrow Channel Hall Thruster(NCHT) ExperimentalCharacterization

This section describes the experimental characterization of the NCHT plume using aplanar probe and the RPA. Both the near and far field plume were characterized inseparate experiments. Only the planar probe could be used for the near field scans,as the large Semion RPA probe adversely disrupts the plume. After modifying the gasdistributor, near field scans were then repeated, and the far field scans were taken. Thespecific experimental setup for both sets of scans is described, followed by an analysis ofthe results.The magnetic field internal and external to the NCHT channel was measured to vali-

date modelling in the Finite Element Method Magnetics (FEMM) open-source softwareprogram and inform design changes to the thruster, described in 4.1.

3.1 Near Field Experimental Setup & Method

3.1.1 Vacuum Facility

The larger vacuum facility (Fig. 3.1) is cylindrical, 2.7 m long and 1.2 m in diameter(≈3.05 m3 internal volume). It is equipped with 3 Sumitomo CP-22 Cryopumps and1 Oerlikon Leybold Vacuum 251/D65B Three-Stage RUTA Pump System. The RUTApump is used as a forevacuum pump, with a maximum pumping speed of 58 L/s. The CP-22 pumps have a total maximum pumping speed of 12,500 L/s for Xenon, and can obtaina minimum chamber pressure below 5.5 × 10−8 mbar. During thruster operation, thechamber pressure was maintained below 2.0× 10−5 mbar. Vacuum pressure is measuredby an Oerlikon Leybold Vacuum Ionivac ITR90 Pirani and Bayard Alpert hot cathode-type combination gauge with 5×10−10 to 1000 mbar pressure range and ±15 % accuracy.Xenon gas flow to the HET is controlled by a MKS Mass-Flo®M100B flow controller


Figure 3.1: Schematic of the general experimental setup for the larger vacuum facility.

with 0 to 196 mg/s flow range. The xenon is supplied from a single gas cylinder with1000 L volume of Xenon at 50 bar, via a pressure regulator that reduces the inlet pressureat the flow controller to 1.5 bar. Inside the chamber is a custom Platar Ltd. SEMS-200thrust balance with a suspended pendulum feedback mechanism capable of measuringthrust levels between 0.5 mN and 200 mN with measurement noise of 0.05 mN. TheNCHT was mounted directly onto the thrust balance for all measurements.

3.1.2 Probe Setup, Data Acquisition & Measurement Procedure

The near field scanning mechanism is shown in Figs. 3.3 & 3.2. The axial linear stagewas a 125 mm precision stage, and the radial linear stage was a 200 mm precision stage,both with Empire Magnetics Inc. VS-U22 Vacuum Stepper Motor, driven by a SchneiderElectric MForce MicroDrive MFI3CRD17N4 programmable motion controller. The nearfield scans were fully automated using a LabView program, with the maximum possiblescan area a rectangle, 125 mm in the axial direction and 200 mm in the radial direction.A single horizontal plane was scanned, with the cathode positioned approximately at thethruster’s 12 o’clock. Scans were performed of the entire plume from 0.5 mm to 50.5 mmaxially, and -25 mm to 25 mm radially. Close-up scans of the channel exits were alsotaken from 0.5 mm to 20.5 mm axially, and ±20 mm to ±10 mm radially.



Figure 3.2: Photo of the near field scanning setup for the planar probe. The linear stagescontrolling axial and radial position are at the top, with the planar probeholder extending down into the plume. The planar probe is located at theend of the thin ceramic tube at the end of the probe holder.



Figure 3.3: CAD drawing of the near field probe positioning mechanism. When onlythe planar probe is being used, the smaller linear stage and RPA holderare removed, and the planar probe holder is fixed to the stage arm. Thismaintains the probe facing parallel to the thruster axis.



Figure 3.4: Schematic of a generic biased planar probe

Only the biased planar probe was used for the near field scans. The Semion RPAprobe presents an area equivalent to the total area of the thruster, and thus disturbs theplume to the point of preventing the thruster from operating. Even the much smallerplanar probe disrupted the thruster operation at low flow rates (Fig. 3.9). Additionally,the larger collection area of the Semion RPA does not allow for adequate resolution ofthe narrow channel. A schematic of the biased planar probe used is shown in Fig. 3.4.It was mounted parallel to the thruster main axis. The total probe area was calculatedas:

Ap =π

4(0.00132 − 0.00052) = 1.13× 10−6m2 (3.1)

Both the probe bias control and current measurements were managed by a LabViewprogram via a National Instruments USB-6216 BNC multi-function I/O Data Acquisition(DAQ) device. The USB-6216 communicated with low-voltage signals (maximum ±10 V)to a Trek Model 2205 Power Amplifier. The amplifier has a gain of 50 V/V on the HighVoltage input line, a High Voltage Monitor signal output stepped down by the same gain(theoretically equivalent to the High Voltage input signal for steady-state operation),and a Current Monitor line that outputs 0.1 V per 1 mA of current measured on theHigh Voltage connection. One of the analog outputs of the USB-6216 was connectedto the High Voltage input line of the power amplifier to control the planar probe bias.Both the High Voltage Monitor and Current Monitor were measured via two analoginputs on the USB-6216. All connections were standard BNC cables with no externalconditioning circuits. Bias sweeps from -100 V to 60 V were performed to obtain theminimum negative bias voltage within the ion saturation region. Subsequent scans useda constant -60 V bias. The circuit of the setup is shown in Fig. 3.5.



Figure 3.5: Circuit diagram showing the electrical connections between the experimentalequipment.

The position of the axial stage, probe current, and probe applied potential were cap-tured at high resolution in the axial direction, with a spatial resolution of less than 0.001mm, using the USB-6216. Synchronized sample clocks were implemented based on thestep pulse from the MForce MicroDrive controller to ensure concurrent measurementswere taken. For the full plume scans, a radial resolution of 1 mm was used. For theclose-up scans, a radial resolution of 0.2 mm was used to resolve the narrow 1 mm chan-nel. Scans were repeated under a number of different thruster operating conditions, listedin Table 3.1. The thrust was measured by the thrust balance inside the vacuum chamber.The near field positioning system introduces a negligible error due to its high precision.

A 1 % error results from the biased planar probe area measurement. The noise inthe current signal from the high voltage amplifier is high at ±16 %, however the 150moving point average smoothing mitigates this. With the Semion system error, the totalmeasurement error in the near field results is 21 %.



Table 3.1: NCHT operating conditions for near field scans

Scan Mass Flow Rate Power Thrust Isp Anode Efficiencymg/s W mN s %

1 0.40 23.3 1.04 265.0 5.812 0.50 27.6 1.20 243.5 5.173 0.34 23.9 0.91 269.6 5.044 0.32 38.7 1.13 353.5 5.04

5 & 6 0.35 24.4 0.94 270.8 5.12

3.2 Near Field Results

3.2.1 Gas Distributor Modification

Figure 3.6: Ion current density as measured in the near field of the unmodified NCHT.The channel exits are shown in black on the left. Note the significant disparitybetween the upper and lower channel exits.

Figure 3.6 shows the ion current density distribution within the near field before thegas distributor was modified. There is a clear discrepancy between the upper and lowerchannel exits, with the ICD at the upper exit 2.4 times that at the lower exit. As thecathode was located approximately in the center of the scan, it would not have such a



strong asymmetric effect on the plume in the scanning plane. Therefore, irregularities inthe gas distributor ring were suspected. Bubble tests of the gas distributor ring in wateralso demonstrated that the gas preferentially exited one side of the ring. This resultsin a higher local density of neutral particles in the channel, producing more plasma andions than in other channel locations.

It is noted that the thruster operation was unstable at flow rates at or below 0.32mg/s.

Figure 3.7: Visual comparison of NCHT plumes before (left) and after (right) gas dis-tributor modification using the same camera settings and lighting conditions.The plasma probes scanned the plane perpendicular to the one pictured.

The gas distributor was modified to more evenly distribute the Xenon propellant, andthe near field scans were re-performed. A simple 150-point moving average was appliedto the current data in the axial direction to reduce noise in the signal (Fig. 3.8). Visually,the plume still exhibits asymmetry in the vertical plane, with the higher intensity locatedtowards the edge of the thruster closest to the cathode (Fig. 3.7). The planar probe data



Figure 3.8: A sample axial scan taken at 0 mm radial position. The raw data is in blue,with the 150 moving point average shown in red.

confirmed that the modifications were ineffective, with the high intensity peak ICD now4.1 times the peak ICD near the lower channel exit.

3.2.2 Effect of Flow Rate

The flow rate of Xenon propellant was adjusted, and stable operating conditions foundfor the thruster giving similar thrust and efficiency (Table 3.1). Four flow rates weretested; 0.32 mg/s, 0.34 mg/s, 0.40 mg/s and 0.50 mg/s. The results are shown in Fig.3.9. It can be seen that the ion current density increases with increasing flow rate, as ahigher neutral particle density within the ionization region increases the likelihood of anionizing collision. The anomalies between 0 mm and 5 mm radial distance in the 0.32mg/s case were caused by unstable thruster operation due to probe interference. At flowrates below 0.32 mg/s, the thruster begins to pulse on and off, which can be seen in thealternating regions of high and low ICD intensities. We note that for the higher massflow rate regimes, the planar probe only induced slight (less than 0.5 V) increases in thedischarge voltage, while no change was observed in the discharge current.Although the intensity changes with flow rate, the geometry of the plume does not.



Figure 3.9: Near field ICD for the four flow rates tested: 0.32 mg/s (top left), 0.34 mg/s(top right), 0.40 mg/s (bottom left), 0.50 mg/s (bottom right). The channelexits are depicted in black on the left of each plot.



Each case, with the exception of the 0.32 mg/s case due to the obfuscation of the pulsing,exhibits a central beam characteristic of Hall thrusters [16, 24, 40]. The beam beginsforming approximately 4 mm from the channel exit, and is fully formed by 15 mm. Byapproximately 20 mm, the plume is considered to be fully formed in the axial direction.Radially, the plume is fully formed by 25 mm. Far field measurements can therefore betaken only at distances exceeding 25 mm, beyond which the thruster can be treated asa point source as the effects of the annular channel are negligible.

3.2.3 Total Ion Current & Thrust Estimation

High resolution scans (Scans 5 & 6) were taken of each channel (Figs. 3.10 & 3.11).Under the assumption of axisymmetric operation, the low and high intensity channelsexits represent low and high side cases, giving a range of estimated ion current and thrust.

Figure 3.10: High resolution scan of upper, high intensity channel exit

The ion current was calculated by numerically integrating the ion current density inthe radial direction over the scanned range (±10 mm to ±20 mm) at the 0.5 mm axial



Figure 3.11: High resolution scan of lower, low intensity channel exit

position:

Ii,tot = 2π

∫ r2

r1

ICD(r)rdr (3.2)

where r1 represents the inner radial scan edge, and r2 the outer radial scan edge.From this value of Ii,tot, the thrust is calculated using Eq. 1.4, assuming an average

ion energy of 50 eV as measured by previous RPA experiments between 70 mm and 90mm.The low side total ion current was 0.045 A giving an estimated thrust of 0.53 mN.

The high side total ion current was 0.251 A giving an estimated thrust of 2.90 mN. Themeasured thrust from Table 3.1 was 0.94 mN, comfortably within the estimated thrustrange. The assumptions made regarding the full capture of the ion flux are thereforevalid.

3.3 Far Field Experimental Setup & Method

The same vacuum facility and data acquisition setup as the near field scans (Section 3.1.2)were used for the far field scans, with the addition of the RPA. The far field positioning



Figure 3.12: CAD drawing of the far field probe positioning mechanism used to sweepthe RPA through the plume. The RPA is shown in brown, and the pulleymass in black. The biased planar probe is mounted to the negative angleside of the RPA.



mechanism shown in Figs. 3.12 was mounted to the chamber bulkhead, using the samelinear stage setup as the near field scans. A probe holder held both the Semion RPA(Section 2.1.2) and the biased planar probe with a 20o angular separation (Fig. 3.13, andwas maintained facing the thruster by a pulley system. At a radial distance of 90 mm,the Semion RPA collected plasma across a 6.5o angular region. The angular accuracywas ±1o with >99 % reproducibility. This mechanism allows the RPA to be positionedfrom 70 mm to 90 mm radial distance through the full plume region from -90o to 90o.Due to the mounting offset, the biased planar probe scanned from -110o to 70o.

Figure 3.13: Schematic showing the definitions of angles for the far field scans. Thethruster is in the same orientation as for the near field scans.

The radial distance of the far field probes is limited in each direction by the capabilitiesof the RPA and the geometry of the plume. The centreline current measured by the planarprobe from near field Scan 1 has been plotted in Fig. 3.14, from the 25 mm, fully formedplume distance. It is expected that the ICD will decay by the inverse square law, and soa x−2 power fit has been applied. Extrapolating this model provides an estimate of themaximum current, and hence ICD, that would be encountered by the RPA at varyingradial distances.For the 10 µA and 60 µA Semion measurement channels, with a total orifice area of

1.86 ×10−5 m2 and grid transmission of 0.125, the currents limits are equivalent to amaximum allowable ICD of 4.3 A/m2 and 25.8 A/m2 respectively. Multiplying these by



Figure 3.14: Ion current measured by the planar probe along the centreline of thethruster, with an inverse square model fitted and extrapolated to 120 mm.

the area of the planar probe provides current limits in terms of the current measuredby the planar probe. The 10 µA channel provides finer resolution of measured current,although a the cost of increased measurement noise, and is preferred over the 60 µAchannel. The maximum allowable planar probe current for the 10 µA channel is 4.87µA, equivalent to a radial distance of 90 mm (Fig. 3.14). The far field scans weretherefore conducted at 90 mm radial distance to maximize the SNR without saturatingthe Semion SCU. The Semion RPA voltage bias settings are listed in Table 3.2, and scanswere completed in time-averaged mode. As the Semion software does not allow externalautomation, a LabView program was used to position the arm while RPA scans werecontrolled manually.

Table 3.2: Semion RPA voltage biases

G1 Collector G2 Sweep Start G2 Sweep End G2 Sweep StepV V V V V

-80.0 -100.0 -40.0 150.0 1.0

Measurements were taken under the conditions listed in Table 3.3, with the thrusteroperating in a higher power 25.5 W mode and lower power 17.3 W mode at constant



flow rate. The positioning system introduces an error of 4.5 % in addition to the 20 %absolute error of the Semion system, giving a total measurement error for the far fieldscans of 24.5 %.

Table 3.3: NCHT operating conditions for far field scans

Setpoint Mass Flow Rate Anode Voltage Power Thrust Isp Anode Efficiencymg/s V W mN s %

1 0.50 85.0 25.5 1.16 235.3 5.232 0.50 75.0 17.3 0.91 185.4 4.80

3.4 Far Field Results

3.4.1 Measured Ion Energy Distribution Function

An example RPA scan taken in setpoint 2 at 0o is shown in Fig. 3.15. The cathodefloating potential with respect to the ground was measured to be +6 V. The RPA datawas corrected accordingly. The data show the characteristic slope expected from RPAscans. As the NCHT operates in a steady-state, the Semion system is able to averagemore samples per scan as opposed to the time-resolved mode, reducing scan noise. A17-point quadratic/cubic Savitzky-Golay smoothing function was applied to the IV trace,and was also used to calculate the IEDF (Fig. 3.16).

The central section of the plume has a higher ion density than the extremities at agiven radial distance. Figs. 3.17 & 3.18 show the IEDFs within 20o of the thrustercentreline. These angles had considerably higher ion current measured by the RPA thanwider angles, indicating that the main beam ions are contained within ±20o. This alignswith the near field scans where a directional beam is clearly present (Fig. 3.6). Theasymmetry of the NCHT observed in the near field scans is also visible in the far fieldresults, although it is not as pronounced. It is interesting to note that for angles above20o, the asymmetry of the plume is not as apparent, and in fact reverses, with the positiveangles recording higher ion currents. Considering the peak 0o current measured for bothpower modes, setpoint 2 has a peak current of 9.6 µA, compared to 14.2 µA for setpoint1.



Figure 3.15: IV data from a single RPA scan taken at setpoint 2 and 0o. More datasamples can be averaged for each measurement point than in the time-resolved mode used for the ISF-VAT, improving the SNR. The 17-pointSavitzky-Golay smoothing function preserves the shape of the data well.

Figure 3.16: IEDF smoothed with a 17-point Savitzky-Golay function calculated fromthe data in Fig. 3.15 with the thruster in setpoint 2 at 0o.



Figure 3.17: IEDFs within the main beam with smoothing functions applied for setpoint1.

Figure 3.18: IEDFs within the main beam with smoothing functions applied for setpoint2.



This gives a reduction in peak current of:

1− 9.6µA

14.2µA= 32.3% (3.3)

This matches well with the reduction in input power:

1− 17.3W

25.5W= 32.2% (3.4)

The most probable ion energy was 56.5 eV for setpoint 2 compared to 71.9 eV forsetpoint 1. The first ionization potential of Xenon is 12.13 eV [41]. Therefore, accordingto the respective discharge voltages, the energy per ion/electron creation for setpoint 1and 2, are 13.1 eV and 18.5 eV respectively, within the expected range. The IV tracesand IEDFs for all angles and both setpoints can be found in Appendix B.

The FWHM of the IEDFs was also calculated for each setpoint, shown in Fig. 3.19.Both thruster setpoints exhibit the same profile, with a reduced FWHM value at theextremities and between -20o and 20o. The main beam therefore has a smaller spread ofion energies and a half-angle of 20o.

Figure 3.19: Full Width Half Maximum for the IEDFs between ±40o and both thrustersetpoints. A beam is clearly visible between -20o and 20o.



Figure 3.20: Ion current density angular distribution as measured by the RPA and biasedplanar probe.

Figure 3.21: Ion current density at all angles and both power modes. The main beam isseen as the ICD peak between -20o and 20o.



3.4.2 RPA Measured Ion Current and Half-Angle Plume Divergence

The ion current density angular distribution was calculated from both the RPA data andplanar probe (Fig. 3.20). Both probes recorded a similar maximum, and show the sameangular profile until ≈ ±40o, where the RPA exhibits ’wings’. Previous studies have alsoobserved these wings, attributing them to CEX ions [4]. It is unclear why the wings wereonly observed in the RPA data and not also by the planar probe. Figure 3.21 showsthe ICD angular distribution for both power modes, with the main beam clearly visible.The ion current density drops sharply outside of the main beam, and gradually reduceswith increasing angle. Again, the asymmetry of the thruster is still visible, with negativeangles showing slightly higher ion current density. The near field scans revealed a centralspike that forms slightly closer to the low intensity channel exit. This manifests itself inFig. 3.21 as a higher ICD at 10o than -10o, despite positive angles measuring lower formost other angles.Performance metrics were calculated for the two power modes from RPA data and are

listed in Table 3.4.

Table 3.4: Total ion current measured in the far field

Setpoint Ii,tot θ T ηm ηT vi

A o mN % % m/s

1 0.105 70.8 0.95 31.9 7.2 91132 0.079 70.6 0.67 23.8 6.9 10281

The normalized integrated ion current by angle is shown in Figs. 3.22 & 3.23. Thehalf-angle of divergence remains relatively constant at ≈71o, indicating that the plumegeometry is not affected by the input power. Despite the main beam being much morefocused with a half-angle of 20o, higher angles still exhibit non-negligible ion current.This gives the plume a much higher half-angle of divergence.



Figure 3.22: Normalized integrated ion current for setpoint 2. The blue line indicates the95 % value, and the red line the half-angle.

Figure 3.23: Normalized integrated ion current for setpoint 1. The blue line indicates the95 % value, and the red line the half-angle.



3.4.3 Thrust Measurements and Specific Impulse

The thrust was calculated using the IEDF to determine the relative current contributionof ions at each energy level. The thrust for both power modes was in good agreementwith the thrust balance measurements in Table 3.3. The ion velocities in Table 3.4 weretaken from the most probable RPA ion energy data. The effective specific impulse listedin Table 3.3, calculated from input and thrust balance measurements, includes systemefficiencies and is hence much lower.

3.4.4 Thruster Efficiencies

From the measured total ion current, the mass utilization efficiency, ηm can be found.For setpoint 1, the mass utilization efficiency was 31.9 %, compared to 23.8 % for setpoint2. As expected, a lower input power ionizes a lower proportion of the propellant. Thereduction in ηm was:

1− 23.8

31.9= 25.4%, (3.5)

approximately equal to the reduction in input power accounting for the minimum Xenonionization energy overhead and measurement error. The RPA total system efficiencies are≈40 % higher than the values calculated from input and thrust balance measurements inTable 3.3, close to the upper limits of measurement error. The assumption of azimuthallyconstant total ion current may be incorrect, contributing to the difference. Without fullycharacterizing the azimuthal asymmetry of the thruster, it is impossible to know exactlywhat the total ion current is. If the value measured in the single horizontal plane islower than the true value, ηT would increase according to Eq. 1.16. Thruster efficienciesof 15 % to 35 % were reported for similarly sized Hall thrusters operating with anodepowers of between 70 W and 200 W [20]. The asymmetry of the thruster likely reducesefficiency somewhat, and so it may be possible with future improvements to raise thetotal efficiency to similar lower end values, at less than one third the input power.

3.4.5 Comparison to NASA-173Mv2 Hall Thruster

In order to put the capabilities of the NCHT in context, the results are briefly comparedwith a larger Hall thruster that has also been well characterized. The NASA-173Mv2Hall thruster is an SPT-type thruster designed for input powers of 1.2 kW to 5.2 kW,



with an outer channel diameter of 173 mm. It uses external and internal trim coils toinfluence the shape of the magnetic field within and external to the channel [4].The performance metrics compared are listed in Table 3.5.

Table 3.5: Total ion current measured in the far field [4]

Thruster Power mp T Isp θ ηT Thrust-Power RatioW mg/s mN s o % mN/kW

NCHT 25.5 0.5 1.16 235.3 71.0 5.23 45.5NASA-173Mv2 3920 5.0 140 2700 40.0 50.0 35.7

The NASA-173Mv2 consumes 150 times the power to operate, producing only 120 timesthe thrust for the operating mode described in Table 3.5. This results in the NCHTproducing a thrust-power ratio of 45.5 mN/kW to the NASA-173Mv2’s 35.7 mN/kW.For nanosatellites, maximizing the thrust-power ratio is vital in reducing the powergeneration requirements for the satellite.Although it produces a higher thrust per unit power, the NCHT suffers from compar-

atively low total efficiency and specific impulse. This is partly due to the less focusedplume, with a half-angle divergence 31o larger than the NASA-173Mv2. Additionally, theNCHT is currently operated with a relatively high mass flow rate (10 % of the NASA-173Mv2 with <1 % the input power). This restricts propellant flow and utilizationefficiency within the channel, reducing total efficiency and hence Isp.


4 Conclusion

4.1 Summary and Conclusions

Experimental investigation of the far field ion plume of two miniature plasma thrusterswas carried out for the first time, and the performance of these devices was characterized.A commercial RPA (Semion Single) was used to analyze the ion content of the plumeand determine the total ion current, ion angular distribution, ion energy distribution,and derive the thrust, Isp, and total efficiency. Thrust was also measured using thrustbalances. Finally, The NCHT near field was scanned using a biased planar probe.

The following performance parameters were measured for the ISF-VAT:

1. The most probable ion energy was 42 eV.

2. The far field plume was found to have a half-angle of divergence of 64.6o soon afterthe arc formation.

3. The maximum total ion current measured was 0.87 A at t = 20 µs after arc ignition.

4. An average thrust of 3.54 µN was calculated from the RPA data, in agreement withthe thrust balance measured value of 4.44 µN.

5. The maximum instantaneous thrust from RPA data was 1.46 mN.

6. The Isp was calculated from RPA data as 714 s.

7. System total efficiency was found to be 3.16%.

The following performance characteristics were derived from near field scans of theNCHT:

1. The NCHT plume is fully formed in the first 20 mm from the channel exit plane.


2. The flow rate of Xenon propellant affects the magnitude of the ion current densityin the plume, but does not affect the geometry.

3. High resolution scans were performed of the channel exit regions, providing a totalion current between 0.045 A and 0.251 A.

4. The thrust was calculated from the total ion current as between 0.53 mN and 2.90mN, with the thrust stand measurement in the middle at 0.94 mN.

The following performance metrics were derived from far field scans of the NCHT at17.3 W and 25.5 W power modes:

5. The plume contains a main beam with a half-angle of 20o.

6. The plume has a half-angle of divergence of 71o. This angle was consistent for bothpower modes tested, indicating that the geometry of the plume is not affected bythe thruster input power.

7. The most probable ion energy was 56.5 eV for setpoint 1, with an anode voltageand current of 85 V and 0.3 A. For setpoint 2, it was 71.9 eV, with an anode voltageand current of 75 V and 0.23 A. Both values are lower than their respective anodevoltage by approximately the ionization energy for singly charged Xenon ions in aground or excited state.

8. The thrust measured for the 17.3 W and 25.5 W modes was 0.67 mN and 0.95 mNrespectively. This is in agreement with the thrust balance values of 1.16 mN and0.91 mN.

9. The Isp calculated from accelerated ion energies was 929 s in the 17.3 W modeand 1048 s in the 25.5 W mode. These are significantly higher than the effectiveIsp calculated from thrust balance and input measurements as they do not includesystem efficiency.

10. The total system efficiency was found from RPA data to be ≈7% for both setpoints,roughly 40% higher than the value calculate from input and thrust balance data.This is within the measurement error.

The near field scans of the NCHT revealed asymmetric operation due to azimuthalirregularities in the gas distribution. Initial attempts to remediate the issue were unsuc-cessful.


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Appendices

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A. NCHT Magnetic Field Testing

The profile of the magnetic field within the ionization chamber and channel of the NCHThas an influence on the efficiency of the thruster [4]. Near to the channel, electron motionis affected by the magnetic field, which in turn alters the plume profile and ion and plasmaenergy distributions [32]. Increasing the ratio between the magnetic field strength in theradial and axial directions increases thruster efficiency by focusing the plume in the axialdirection [4] [40]. Varying the profile of the magnetic field near to the channel exitresults in a change in the plume divergence angle; as this is usually achieved by usingmultiple trim coils, and the NCHT contains only a single coil, the influence on the plumedivergence could not be measured directly [40].

It is also important that the peak of the magnetic field be located at or near to thechannel exit. Locating the peak downstream of the exit results in decreased wear on thethruster and hence extends the lifetime; conversely, this increases the plume divergenceand reduces the ionization efficiency of the thruster [40]. Given that the low power andmass flow rate of the NCHT already reduces ionization efficiency, maintaining a peakmagnetic field near to the channel is desirable [36].

Non-axisymmetric magnetic field distributions can also lead to differential thrust, affectgas flow within the ionization region, and increase wear on the channel non-uniformly[40]. Non-uniform performance of the thruster introduces uncertainty during operation,requiring a higher pointing accuracy and accurate attitude control from the spacecraft.Magnetic field measurements were therefore taken to determine the magnetic field withinthe channel and near field region. These measurements are used to verify the design toensure the thruster is operating near maximum efficiency and with a suitably focusedplume.

In addition, the magnetic scans were compared to modelling in the FEMM softwareprogram. The magnetic profile of components within the thruster is complex, and theFEMM software allows quick calculation of the magnetic field strength and profile. Aninitial verification was completed using a simple 50 mm diameter test cylinder of SAE1020steel with a 88 turn coil wrapped around roughly the bottom half, before scanning the


thruster itself. The thruster was not operated during testing as it was performed outsideof a vacuum chamber, with power supplied to the solenoid only.

A.1 Experimental Setup and Method

Sampling Equipment

A F.W. Bell 5180 Gauss/Tesla Meter was connected to a National Instruments USB-6216ADC for measuring the magnetic field. Data collection was automated using a LabViewprogram. Background measurements taken with zero current through the test cylindercoil showed a drift of only 2.5×10−5 mT/s over the sampling period of approximately360 s (Fig. A.1).

Figure A.1: Zero coil current sampling data taken for the test cylinder. The measure-ments were taken twice, with both data sets showing a near-zero drift (redlines).



A.1.1 Magnetic Probes

Three F.W. Bell Hall effect probes were used for the test cylinder case to assess theirrelative accuracy: an axial probe, large transverse probe, and small transverse probe(Fig. A.2). Hall effect sensors rely on the interaction between a perpendicular magneticfield and applied current to produce an output voltage variance proportional to thestrength of the perpendicular magnetic field [42]. The probes must therefore be orientedappropriately relative to the local magnetic field. All three probes were connected to thesampling system and used to measure a 25 mm × 25 mm region from the centre of thetest cylinder to the outer edge, in both radial and axial directions. It was assumed thatthere was zero azimuthal component to the magnetic field. The total magnitude of themagnetic field, calculated using Equation A.1, was then compared to the FEMM model,shown in Fig. A.3.

|B| =√B2a +B2

r (A.1)

where Ba is the axial magnetic field strength and Br is the radial magnetic field strength.

Figure A.2: The three magnetic probes: large transverse (top), small transverse (middle),and axial (bottom)

Both the axial and small transverse probes match well to the FEMM values. In par-ticular, the axial probe shows good agreement, with the small transverse probe having ahigher Signal-Noise-Ratio (SNR). The large transverse probe produced a similar profileat a lower magnitude and was therefore excluded from future scans. The axial probe,while the most accurate, could not fit within the narrow (1 mm wide) channel of the



Figure A.3: Comparison between the three magnetic probes. The FEMM model resultsare shown by the solid black line. At every radial position, the Large Trans-verse Probe undermeasures the magnetic field strength.



thruster. The small transverse probe was hence used for all scans, both internal andexternal to the channel, for consistency.

A.1.2 Positioning Devices

Two Zaber T-LSM025B-SV2 linear vacuum stages were used to position the probe in theradial and axial directions for all scans (Fig. A.4). These stages have an accuracy of 15µm with a reproducibility of >99%. The thruster was mounted on a 3-axis manual stagefor precise positioning with an accuracy of ±0.05 mm in all three axes.

Figure A.4: Image of the test cylinder scan experimental setup, showing the two linearstages mounted perpendicular to each other, with the axial magnetic probepositioned for radial scans.

For the scans of the NCHT, a servo motor was included to provide azimuthal po-sitioning for measuring the axisymmetry of the magnetic field. A Leadshine HBS507Servo Drive drove a 573HBM20-1000 Easy Servo Motor with built-in rotary encoder.The HBS507 was controlled by a LabView program via the USB-6216 digital outputs.The default setup of the HBS507 and servo motor gave a resolution of 4000 steps perrevolution. A bracket system, shown in Fig. A.5, was designed to mount the thruster andservo motor to the 3-axis manual stage for fine positioning control. The manual stagewas used after every movement of the servo motor to realign the probe within the channeland ensure the probe was positioned correctly before sampling began. This removed asignificant portion of alignment error from manufacturing tolerances.



Figure A.5: Image of azimuthal measurement experimental setup. The thruster ismounted on a plate attached to the shaft of the servo motor, itself mountedin a bracket attaching it to the 3-axis manual stage.



A.2 Finite Element Method Magnetics (FEMM)

FEMM is an open-source 2D magnetic modelling software program that uses the finiteelement method to numerically calculate the steady-state magnetic field within a region ofinterest. It includes a basic drawing suite with the ability to include rudimentary shapessuch as lines, squares, and arc segments. Real-life devices must therefore be modelledusing simplified geometry. Models can be run in either axisymmetric or infinite-depth2D mode. A far-field boundary condition can be applied to the analysis region, allowingcalculations to run as if there were no boundary. The software uses a built-in meshgeneration script that does not allow significant modification of the meshing parameters,so no mesh optimization analysis was conducted.FEMM contains built-in models for a range of magnetic materials. It was observed

during analysis of the NCHT that the results were highly dependent on the B-H curve(Hysteresis curve) used for the material. Peak values of the magnetic field could vary byas much as 25% by using a B-H curve for the same material from a different source (Fig.A.6). Importantly, the test cylinder case was not significantly affected by altering theB-H curve, and so the dependence is geometry specific. For the most accurate results,hysteresis testing should be conducted on a sample of the specific material used.

Figure A.6: Results from two FEMM models run with different B-H curves for the samematerial. The shape of the distribution and location of the peak are thesame for both models, however the magnitude is highly dependent on theB-H curve used.



A.2.1 Coil Area Calculation

FEMM does not offer an easy method of modelling a wrapped coil. Instead, a rectangularregion is denoted as containing a coil, with number of turns, wire diameter, and coilcurrent specified. The following method was used to calculate the height of the coilrectangle based on the number of layers and wire radius (Fig. A.7):

Figure A.7: Schematic of stacked coil geometry

h = r cos(30o) = 0.866r

x = 2r + 2h(n+ 1)

∴ x = (0.268 + 1.732n)r (A.2)

A.2.2 Simplified Geometrical Model

Magnetically unresponsive components were removed from the geometry, and replaced byregions of air. The thruster was also assumed to be axisymmetric. Figure A.8 shows thesimplified rotational cross-section used in the FEMM modelling. This was surrounded bya 200 mm radius semi-circular region of air with an infinite boundary condition applied.



The floating line through the centre of the channel represents the scanning path.

A.3 Results and Discussion

A.3.1 Vertical Channel Measurements

The radial magnetic strength was measured every 0.1 mm from 50.4 mm to 60 mm onthe centreline of the channel, at coil currents of 0.25 A, 0.50 A, 0.70 A, 1.00 A and 1.40A. After initial power up, the coil current was allowed to settle for 5 to 10 mins as the coilheated and changed its resistance. Only radial measurements were taken, based on theassumption that azimuthal field strength would be zero, and the probe design prohibitingaccess for axial measurements.

Data was extracted from the FEMM model results using a custom LUA script thatoutputs the data at specified sample points to a text file. With the exception of the 0.25A case, the FEMM model generally overpredicts the magnitude of the magnetic field.The 1.4 A case shows good agreement with the measured data. Of particular interestis the location of the peak magnetic field strength. For maximum efficiency, this shouldbe located at the channel exit for the reasons described in Section 4.1 [40]. Previousmodelling and testing had indicated that the peak was located inside the channel exit,slowing generated ions before they exited the thruster and reducing efficiency. It can beseen in Fig. A.10 that the tested improved design locates the magnetic field peak withinthe channel exit. Although there is a discrepancy between the magnitude of the FEMMpredictions and measured data, the shape of the magnetic field is in good agreement inevery case, with less than 0.1 mm variance in the axial location of the peak field strengthbetween models and measured data.

A.3.2 2D External Field Measurements

Measurements were then taken of the external magnetic field. A 25 × 25 mm radialslice from the centreline of the thruster across the channel exit was scanned at the sameset of coil currents as for the channel measurements. Figure A.11 shows an example ofthe results from the 0.5 A case. Figures A.12 - A.14 show the results for various axialdistances at 0.50 A, 0.70 A and 1.00 A coil current.

The correlation between the FEMM model and measured data is difficult to assessgiven the non-linearity of the model and distribution in both dimensions. To providea rudimentary gauge of correlation, the measured data were calculated as a weighted



Figure A.8: FEMM geometrical model with simplified NCHT geometry. The origin islocated at the bottom left node, on the axis of symmetry.



Figure A.9: Small transverse probe located within the thruster channel. The probe waspositioned manually before each measurement.

Figure A.10: Total magnetic field strength as measured and calculated in FEMM for allcoil currents. The black box represents the gas distributor, and the verticaldashed line locates the channel exit.



average percentage of the FEMM model according to:

madj =MM

(A.3)

C =

(1

n

∑n

BnMn ·madj,n

)−1(A.4)

where M is the 2D matrix of FEMM model values, M is the average FEMM value overthe sample region, n is the number of sampling points, Bn is the value of the measureddata at point n, and Mn is the value of the FEMM model at point n. This correlationcalculation requires the FEMM model and experimental sample points to be aligned.The ratio of FEMM value to measured value is adjusted according to the magnitude ofthe FEMM model at that point relative to the average FEMM value. This mitigatesthe influence of background noise or bias at low magnetic field intensities in the measuredata, which can be seen in Fig. A.11. A value of 1 indicates a perfect correlation. TableA.1 lists the correlations for the various coil currents. The 0.50 A case has the closestcorrelation and 0.25 A the worst. As for the internal channel scans, all FEMM modelshad a similar shape of magnetic field as the measured data.

Table A.1: Correlation between FEMMmodel and measured data at varying coil currents

Coil Current (A) Correlation, C

0.25 0.6370.50 0.9290.70 1.2071.00 1.1191.40 1.125

A.3.3 Azimuthal Channel Measurements at 1.00 A

In order to determine the azimuthal symmetry of the thruster, vertical channel scanswere taken every 10o around the full thruster channel. Figure A.15 displays all of thescans against the axisymmetric FEMM model. The FEMM model overestimates themagnetic field strength in every case, in particular close to the magnetic field peak at



Figure A.11: Surface plot of FEMM model and measured data for the 0.5 A case. Thepeak aligns with the thruster channel. Note the offset at lower magneticfield strengths due to measurement noise.



Figure A.12: Total normalized magnetic field strength as measured and calculated inFEMM for the 0.50 A case (C = 0.929)



Figure A.15: Magnetic field strength across the channel measured every 10o comparedwith the 1.0 A FEMM model output

the channel exit. The measured data themselves show a variation in peak field strengthof 22%. While close, the axial location of the peaks also vary by up to 0.4 mm.

In Fig. A.16, the peak values have been plotted against azimuthal position, and asinusoidal fit has been applied. There is a clear sinusoidal trend in the data, implyinga dependence on azimuthal position with a maximum at one position, and minimum180o opposite. The magnetic field strength at the centreline of the channel depends onthe channel geometry; bringing the tips of the poles closer together (i.e., narrowing thechannel) produces a higher peak magnetic field strength. The sinusoidal variation can beexplained by an offset in the alignment of the inner and outer poles, shown exaggerated inFig. A.17. It was verified that the operation and position of the servo motor has no effecton the magnitude of the magnetic field measured by the probe. Based on this testing,a new thruster was manufactured to tighter tolerances and with a new pole alignmentmechanism.

The axial location of the peak field strength should not be affected by thee width ofthe channel, however a rough sinusoidal variation was also observed, as seen in Fig. A.18.The source of this may be a slight tilting of the inner pole relative to the outer pole;however, it could also arise from from the positioning method used. The zero positionof the probe was set by placing the probe flush with the top of the gas distributor topsurface using the 3-axis manual stage. It was assumed that the gas distributor top surfacewas level with respect to the poles; if it is not, this would account for the variation in



Figure A.16: Variation in peak magnetic field strength by azimuthal position

Figure A.17: Schematic of offset poles showing exaggerated channel width variation



peak location.

Figure A.18: Variation in axial location of peak of magnetic field by azimuthal position


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B. NCHT Far Field Full Angle Charts

Figure B.19: IV RPA data with smoothing functions applied for setpoint 1. Positiveangles (left) and negative angles (right) are plotted separately for clarity.


Figure B.20: IV RPA data with smoothing functions applied for setpoint 2. Positiveangles (left) and negative angles (right) are plotted separately for clarity.

Figure B.21: IEDFs with smoothing functions applied for setpoint 1. Positive angles(left) and negative angles (right) are plotted separately for clarity.



Figure B.22: IEDFs with smoothing functions applied for setpoint 2. Positive angles(left) and negative angles (right) are plotted separately for clarity.


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