HALL THRUSTER DIRECT-DRIVE ASSESSMENT AND DEMONSTRATION

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TESI DI LAUREA

IN

INGEGNERIA AEROSPAZIALE

INDIRIZZO SPAZIALE

HALL THRUSTER DIRECT-DRIVE

ASSESSMENT AND DEMONSTRATION

Alessio Pampaloni

Mirko Trisolini

Anno Accademico 2012-2013

T291

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UNIVERSITA’ DEGLI STUDI DI PISA

Tesi di Laurea in Ingegneria Aerospaziale

Indirizzo Spaziale

HALL THRUSTER DIRECT-DRIVE

ASSESSMENT AND DEMONSTRATION

Candidati

Alessio Pampaloni

Mirko Trisolini

Relatori

Prof. Mariano Andrenucci

Ing. Tommaso Misuri

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I

Abstract

This thesis presents a theoretical and experimental study of the Hall thruster Direct-Drive configuration: an innovative way to deliver power to electric thrusters and a candidate for future spacecraft’s propulsion system architecture. The direct connection between the solar array and the Hall thruster allows a drastic simplification of the Power Processing Unit (PPU) of the propulsion system. This has an immediate impact also on the thermal control system (TCS) which can be consequently lightened. Further mass benefits can be obtained in other subsystems of the spacecraft such as the electric power system.

The work is mainly divided in two parts. The first one assesses, in terms of mass reduction, the impact that the Direct-Drive configuration entails in the spacecraft subsystems; different kinds of space missions are considered, with different level of Hall Effect thruster power. Although the mass advantages that the Direct-Drive can afford are mission dependent, it has been proved that as the thruster and spacecraft power increase also the mass benefits become larger.

The second part of thesis concerns an experimental demonstration of a Direct-Drive system supplying the Alta’s HT-100, a low power Hall thruster. The test required the procurement of the solar panel and the design of an electrical filter. By means of simulations with Pspice software and experimental tests, a LC filter was developed and then arranged between the solar array and the thruster in order to dampen the current oscillations. This test successfully proved the correct ignition and the thruster operations up to 370 W of discharge power, representing in this way the first accomplished attempt in Europe of a Direct-Drive application.

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III

Sommario

Questa tesi si occupa dello studio teorico e sperimentale della configurazione

Direct-Drive per i motori ad effetto Hall: un modo innovativo di fornire potenza ai propulsori

elettrici e candidata per futuri sistemi propulsivi di veicoli spaziali. La connessione diretta tra i pannelli solari e il motore Hall permette una drastica semplificazione dell’unità di processo della potenza del sistema propulsivo. Questo ha un impatto immediato anche nel sistema di controllo termico che può essere di conseguenza alleggerito. Inoltre, ulteriori benefici di massa possono essere ottenuti in altre parti del veicolo spaziale come per esempio nel sistema elettrico di potenza.

Il lavoro è principalmente diviso in due parti. Nella prima parte si è valutato in termini di riduzione di massa l’impatto che la configurazione Direct-Drive comporta nei sistemi del veicolo spaziale; sono stati analizzati differenti tipi di missione con motori Hall con diversi livelli di potenza. Nonostante i vantaggi di massa dipendano dal tipo di missione, si è dimostrato che in generale all’aumentare della potenza del motore e del veicolo anche i benefici di massa crescono.

La seconda parte della tesi riguarda una dimostrazione sperimentale di un sistema

Direct-Drive che ha alimentato il propulsore HT-100, un motore di bassa potenza

sviluppato ad Alta. Il test ha richiesto l’approvvigionamento di pannelli solari e il progetto di un filtro elettrico. Mediante simulazioni con il software Pspice e test sperimentali, è stato sviluppato un filtro LC che poi è stato interposto tra il motore e il pannello solare per smorzare le oscillazioni di corrente. Questo test ha dimostrato con successo il corretto funzionamento del motore fino ad una potenza di scarica di 370 W, rappresentando in questo modo il primo tentativo riuscito in Europa di una dimostrazione Direct-Drive.

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V

Acknowledgements

The work described in this thesis has been developed and funded by Alta S.p.A. a leading European small company in the aerospace propulsion sector. A special thank goes to the professor and president of Alta S.p.A. Mariano Andrenucci who proposed us the theme of the thesis, pushed and authorized us in planning the experimental phase of the work. We want also to sincerely thank the engineer Tommaso Misuri who has followed us during the whole development of the thesis, supporting us and coordinating the team of Alta’s engineers. An important role has been played also by the engineer Luca Pieri whose knowledge in the electric and electronic field has been fundamental.

We want also to acknowledge the engineer Riccardo Moreschi who helped us in the solar modules procurement. A thank must go to the engineer Cosimo Ducci who managed the Hall thruster during the experimental tests. Finally, we want to thank the engineers Riccardo Albertoni and Stefan Gregucci for their advices.

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IX

Abstract ... I

Sommario ... III

Acknowledgements ... V

List of Figures ... XV

List of Tables ... XIX

Nomenclature ... XXI

Acronyms ... XXV

Introduction ... 1

Part A

1. Hall Effect Thruster ... 7

1.1. Introduction ...7

1.2. Crossed-Field Structure and Hall Current ...9

1.3. Ionization length and scaling ...11

1.4. Potential and Current Distribution ...12

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Table of Contents

X

1.6. Thrust Efficiency ... 15

1.7. Hall Thruster Lifetime ... 17

1.8. Hall Thruster Oscillations ... 18

1.9. Hall Thruster Plume ... 19

1.9.1. Introduction ... 19

1.9.2. Spacecraft Interactions ... 20

1.9.3. Sputtering and Contamination ... 21

2. Power Processing Unit (PPU) Description ... 23

2.1. Anode Supply ... 23

2.1.1. EMI Filters... 25

2.2. Ignitor (or Keeper) Supply ... 25

2.3. Heater Supply ... 26

2.4. Electromagnet Supply ... 26

2.5. Fuel Flow Controller Supply ... 26

2.6. PPU Architecture ... 27

3. Direct-Drive Implementation... 31

3.1. Direct Drive Description ... 31

3.2. Direct Drive System Advantages ... 33

3.2.1. Direct Advantages ... 34

3.2.2. Indirect Advantages: High Voltage Bus ... 34

3.2.2.1. Alternative Solutions for Voltage Bus Regulation... 37

3.3. Direct Drive Implementation Issues ... 38

3.3.1. Plasma Phenomena ... 38

3.3.2. Ignition ... 42

3.3.3. Thruster Performance Regulation ... 42

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XI

4. Direct-Drive Implementation in Low-Power Thrusters ... 47

4.1. TacSat-2 Mission ...47

4.1.1. Mission Profile ...48

4.1.2. The Spacecraft ...48

4.1.3. Electric Propulsion System ...50

4.1.4. Direct-Drive Implementation Assessment in TacSat-2 Spacecraft ...51

4.1.4.1. Direct Advantages ...52

4.1.4.2. Indirect Advantages ...52

4.1.4.3. Mass Saving Result ...54

4.2. Meteosat-8 Mission ...55

4.2.4. Direct-Drive Implementation Assessment in Meteosat-8 Spacecraft ...58

5. Direct-Drive Implementation in Medium-Power Thrusters ... 63

5.1. STENTOR Mission ...63

5.1.3. Direct-Drive Implementation Assessment in STENTOR Spacecraft ...67

5.2. Smart–1 Mission ...71

5.2.4. Direct-Drive Implementation Assessment in Smart-1 Spacecraft ...74

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XII

5.2.4.2. Indirect Advantages ... 76

5.2.4.3. Mass Saving Result ... 77

6. Direct-Drive Implementation in High-Power Thrusters ... 79

6.1. AEHF Mission ... 79

6.1.1. Mission Profile ... 80

6.1.2. The Spacecraft ... 80

6.1.3. Electric Propulsion System ... 81

6.1.4. Direct-Drive Implementation Assessment in AEHF Spacecraft... 83

6.1.4.1. Direct Advantages ... 84

6.1.4.2. Indirect Advantages ... 85

6.1.4.3. Dual Mode Thrusters vs Direct-Drive Configuration ... 87

6.1.4.4. Mass Saving Result ... 89

6.2. Deep Space Mission with High-Power HETs ... 90

7. Direct-Drive Implementation in Very High-Power Thrusters .... 95

7.1. Direct Advantages ... 96

7.2. Direct-Drive Indirect Implications ... 97

8. Direct-Drive Assessment Conclusions ... 101

Part B

9. Test Setup ... 111

9.1. Thruster ... 111

9.2. Solar Array ... 112

9.3. Test Facilities and Instrumentation ... 113

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XIII

9.3.2. Pumping System ... 114

9.3.3. Laboratory Power Supply ... 114

9.3.4. Instrumentation ... 114

10. Filter Design ... 117

10.1. Hall Thruster Modeling with Pspice ... 117

10.2. Solar Array Modeling with Pspice ... 120

10.3. Direct-Drive System Simulation for Filter Design ... 124

10.3.1. Limits of the Simulation ... 128

10.4. Filter Test with Laboratory Power Supply ... 130

11. Solar Array System ... 137

11.1. Solar Irradiance Prediction Model ... 137

11.2. Solar Array I-V Curve Plot... 140

11.3. Real Solar Array Performance ... 140

12. Direct-Drive Demonstration ... 145

12.1. Experimental Test ... 146

12.2. Test Sequence and Results ... 148

12.3. Ignition ... 151

12.4. Current Oscillations Filtering ... 153

13. Direct-Drive Demonstration Conclusions ... 157

13.1. Further Developments ... 158

Appendix A: Anode Supply Mass Assessment ... 161

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XIV

Appendix C: Indirect Mass Savings Esteem Procedure ... 169

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XV

List of Figures

Chapter 1: Hall Effect Thruster

Figure 1.1: Hall thruster cross-section schematic showing the crossed electric and magnetic fields, and the ion and electron paths.

Figure 1.2: Magnetic field lines in the channel region of the NASA-173Mv Hall thruster Figure 1.3: Typical Hall thruster radial magnetic field and axial electric field along the channel length

Figure 1.4: Schematic cross section of the plasma in the Hall thruster channel. Figure 1.5: Hall thruster electrical schematic and potential distribution

Figure 1.6: Electrical schematic for the currents flowing through the discharge plasma and power supply.

Figure 1.7: Measured evolution of the discharge current of the SPT-100.

Chapter 2: Power Processing Unit (PPU) Description

Figure 2.1: Anode supply I-V characteristic

Figure 2.2: 200 W Hall thruster discharge supply (buck derived PWM full-bridge) Figure 2.3: Ignitor pulse form

Figure 2.4: SPT cross section and PPU power supplies scheme Figure 2.5: SPT-100 PPU block diagram (SS/L and Fakel) Figure 2.6: Alcatel ETCA PPU for 1.6 kW SPT-100

Figure 2.7: PPU supplies sequence of operation (1.5 kW SPT-100)

Chapter 3: Direct-Drive Implementation

Figure 3.1: PPU and solar array output I-V characteristic

Figure 3.2: Standard and Direct Drive configuration for electric propulsion system Figure 3.3: Direct drive SPT system

Figure 3.4: Tapped solar array system

Figure 3.5: Current, voltage and power threshold for ISS array

Figure 3.6: High voltage solar array in direct drive configuration vs standard PPU system

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List of Figures

XVI driven HET

Figure 3.8: V-I characteristic of Express BPT-4000 and its solar arrays Figure 3.9: Trajectory analysis approaches for 10 direct drive 30 kW HETs

Chapter 4: Direct-Drive Implementation in Low-Power Thrusters

Figure 4.1: Artistic representation of the TacSat-2 spacecraft Figure 4.2: Busek BHT-200

Figure 4.3: Artistic representation of the Meteosat 8 spacecraft

Chapter 5: Direct-Drive Implementation in Medium-Power Thrusters

Figure 5.1: Artistic impression of STENTOR spacecraft Figure 5.2: STENTOR TCS fluid loop assembly

Figure 5.3: STENTOR Thruster Module: SPT-100 and PPS-1350 Figure 5.4: Plasma propulsion system architecture

Figure 5.5: Artistic representation of the SMART-1 Figure 5.6: EPS functional diagram

Chapter 6: Direct-Drive Implementation in High-Power Thrusters

Figure 6.1: Artistic picture of AEHF spacecraft Figure 6.2: Electric propulsion system schematic Figure 6.3: BPT-4000 Hall thruster

Figure 6.4: PPU block diagram

Figure 6.5: Comparison between HET system (with direct drive and conventional PPU) and available baseline Dawn payload mass

Chapter 8: Direct-Drive Assessment Conclusions

Figure 8.1: Direct Drive mass benefits

Figure 8.2: Mass saving in one PPU due to the implementation of a Direct Drive configuration

Figure 8.3: Percentage of indirect advantages over the total mass saving Figure 8.4: S/C percentage dry mass saving due to direct drive implementation

Chapter 9: Test Setup

Figure 9.1: HT-100 firing at Alta IV-4 test facility (left) and HT-100 mounted on ALTA thrust balance (right)

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List of Figures

XVII Figure 9.2: ALTA’s IV-4 vacuum facility Figure 9.3: The pyranometer (LP PYRA 02)

Chapter 10: Filter Design

Figure 10.1: Hall thruster schematization with Pspice

Figure 10.2: HT-100 current oscillation at anode voltage of 300 V and anode flow rate of 1 mg/s (scale 100mv/A)

Figure 10.3: HT-100 simulated current oscillation

Figure 10.4: Solar array circuit schematization with Pspice Figure 10.5: Nominal solar panel I-V curve simulated with Pspice

Figure 10.6: Filter architecture. Yellow circles represent the location where measurements were taken.

Figure 10.7: Direct-Drive electrical circuit (Pspice) Figure 10.8: Double-log plot of the simulation results Figure 10.9: Filter unit

Figure 10.10: Bode plot of the filter transfer function (neglecting parasitic characteristics of the circuit)

Figure 10.11: Bode plot of the filter transfer function (including parasitic characteristics of the circuit)

Figure 10.12: Filter test with laboratory power supply - electrical circuit Figure 10.13: Thruster current oscillations (scale 100 mV/A)

Figure 10.14: Power supply current oscillations (scale 100 mV/A)

Figure 10.15: Current oscillations (yellow) and its FFT (red) on the thruster side Figure 10.16: Current oscillations (yellow) and its FFT (red) and at the power supply Figure 10.17: Ignition transient. Thruster current peak (scale: 10 mV/A)

Figure 10.18: Ignition transient. Current peak on power supply side (scale: 100 mV/A) Figure 10.19: Ignition transient. Thruster current peak (scale: 10 mV/A)

Figure 10.20: Ignition transient. Thruster current peak (scale: 10 mV/A). Filter removed

Chapter 11: Solar array system

Figure 11.1: Solar array orientation angles

Figure 11.2: Real solar panel I-V curve in standard condition (1000 W/m2_{, 25°C)}

Figure 11.3: Real solar panel power-voltage curve in standard condition (1000 W/m2_,

25°C)

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List of Figures

XVIII Figure 12.1: Alta’s Direct-Drive solar array test-bed

Figure 12.2: Direct-Drive system electrical scheme. Green circles represent locations of current measurements

Figure 12.3: Test sequence of the HT-100 Direct-Drive operation

Figure 12.4.: Current-Voltage curve during test runtime and thruster operative conditions

Figure 12.5: Power-Voltage curve during test runtime and thruster operative conditions Figure 12.6: Ignition transient during Soft Start on thruster side

Figure 12.7: Ignition transient during Soft Start on panel side Figure 12.8: Ignition transient during Hard Start on thruster side Figure 12.9: Ignition transient during Hard Start on panel side

Figure 12.10: RMS value of current oscillations on thruster side and panel side Figure 12.11: FFT of current signal on thruster side

Figure 12.12: FFT of current signal on panel side

Appendix C: Indirect Mass Saving Esteem Procedure

Figure C.1: Curve fitting obtained from data in [53]

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XIX

List of Tables

Chapter 2: Power Processing Unit (PPU) Description

Table 2.1: Qualified PPU data

Chapter 3: Direct-Drive Implementation

Table 3.1: Design impact of bus voltage on Electric Power Subsystem components mass

Chapter 4: Direct-Drive Implementation in Low-Power Thrusters

Table 4.1: BHT-200 main characteristics

Table 4.2: Main TacSat-2 mission characteristics

Table 4.3: Mass savings for the high voltage bus configuration Table 4.4: Meteosat 8 spacecraft characteristics

Table 4.5: Mass savings for the high voltage bus configuration of Meteosat 8

Chapter 5: Direct-Drive Implementation in Medium-Power Thrusters

Table 5.1: STENTOR spacecraft characteristics

Table 5.2: Mass saving due to direct drive implementation in STENTOR spacecraft Table 5.3: SMART-1 spacecraft characteristics

Table 5.4: Mass reduction breakdown for the high voltage configuration

Chapter 6: Direct-Drive Implementation in High-Power Thrusters

Table 6.1: BPT-4000 performance

Table 6.2: AEHF spacecraft characteristics Table 6.3: AEHF spacecraft mass distribution

Table 6.4: Mass saving due to direct drive implementation in AEHF spacecraft Table 6.5: BepiColombo mass budget for three different propulsion systems

Chapter 7: Direct-Drive Implementation in Very High-Power Thrusters

Table 7.1: High Power Hall Thruster performances

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List of Tables

XX

Table 9 .1: HT-100 performance and characteristics

Table 9.2: Schuco solar module main characteristics at an irradiance of 1000W/m2, an

air mass of AM 1.5 and a cell temperature of 25° C.

Table 9.3: Solar Panel nominal characteristics at an irradiance of 1000W/m2, an air mass of AM 1.5 and a cell temperature of 25° C.

Chapter 10: Filter Design

Table 10.1: Solar panel performance at an irradiance of 1000 W/m2, an air mass (AM)

of 1.5 and a temperature of 25°

Table 10.2: Pspice simulation results for different capacitance and inductance values

Chapter 11: Solar array system

Table 11.1: Solar module (Schuco MPE 90 AL 01) temperature coefficients

Chapter 12: Direct-Drive Demonstration

Table 12.1: Direct-Drive test data

Appendix A: Anode Supply Mass Assessment

Table A.1: Chopper voltage-efficiency correlation Table A.2: Transformer step ratio-frequency guidelines

Appendix C: Indirect Mass Savings Esteem Procedure

Table C.1: Impact of the bus voltage level on the power subsystem components Table C.2: Values of the constant c as a function of the system components

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XXI

Nomenclature

A = Area, m2 AM = Available Module B = Magnetic Field, T C = Capacitance, F

CSE = Chopper Stage Efficiency

CSM = Chopper Stage Mass, kg

CSPO = Chopper Stage Power Input, kW

CSVI = Chopper Stage Voltage Input, V

E = Electric Field, N/C

e = Elementary Charge, C

FSE = FilterStage Efficiency

FSM = Filter Stage Mass, kg

FSPO = Filter Stage Power, kW

FSRF = Filter Stage Ripple Factor

FSVO = Filter Stage Voltage Input, V

f = Frequency, 1/s

fI = Current Loss Coefficient

fV = Voltage Loss Coefficient

G = Irradiance, W/m2

g0 = Gravitational Acceleration, m/s2

I = Current, A

Isp = Specific Impulse, s

ITSE = Inverter Transformer Stage Efficiency

ITSM = Inverter Transformer Stage Mass, kg

ITSPO = Inverter Transformer Stage Power, kW

ITSVI = Inverter Transformer Stage Voltage Input, V

ITSVO = Inverter Transformer Stage Voltage Output, V

L = Characteristic Lenght, m

L = Inductance, H

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Nomenclature

XXII

Li = Ionization Mean Free Path, m

l = Magnetized Plasma Thickness, m

m = Electron Mass, kg

m = Mass Flow Rate, kg/s

M = Mass, kg n = Number Density, 1/m3 N = Number of Sample Nv = Avogadro’s number P = Power, W R = Resistance, Ω R = Erosion Rate, m/s RM = Required Module

RSE = Rectifier Stage Efficiency

RSM = Rectifier Stage Mass, kg

RSPO = Rectifier Stage Power, kW

RSVI = Rectifier Stage Voltage Input, V

rL = Larmor Radius, m

SF = Switching Frequency, 1/s

s.e. = Specific Energy Density, Whr/kg

t = Time, s

T = Temperature, °K

T = Thrust, N

T = Period, s

Tc = Solar Cell Temperature, °K

V = Voltage or Potential, V

v = Velocity, m/s

ve = Exhaust Velocity, m/s

W = Atomic Weight, kg

w = Plasma Width, m

Y = Sputtering Yield of the Material

β = Hall Parameter

βIpp = Temperature Coefficient for Current at Peak Power Point, 1/°C

βIsc = Temperature Coefficient for Current at Short Circuit Condition,1/°C

βVpp = Temperature Coefficient for Voltage at Peak Power Point, 1/°C

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Nomenclature

XXIII

Γ = Flux, 1 / (m2_s)

ε = Loss factor

εi = Ionization Loss factor

ΔV = Delta-V, m/s

η = Efficiency

ηi = Current Efficiency

ν = Collision Frequency, 1/s

ρ = Density, kg/m3

σ = Collision Cross Section, m2

ωC = Cyclotron Frequency, 1/s

Subscripts and Superscripts

a = Anode

A = Acceleration

A.C. = Attitude Control

av = Avarage C = Coupling Cg = Cathode-to-Ground b = Beam d = Discharge D = Daylight e = Electron

ea = Electron Arriving at Anode

ec = Electron Arriving at Cathode

E = Eclipse Ei = Ionization Energy eV = Electron-Volt H = Hall H.V = High Voltage i = Ion

ia = Ion Arriving at Anode

ic = Ion Arriving at Cathode

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Nomenclature XXIV m = Propellant Usage n = Neutron OC = Open Circuit p,P = Propellant PP = Peak Power r = Radial Direction

RMS = Root Mean Square

s = Series sa = Solar Array sav = Saved sh = Shunt SC = Short Circuit T = Thrust TE = Transmission t =Thruster th = Thermal v = Spread Velocity w = Wall z = Axial Direction β = Focusing ε = Acceleration

In the case of identical symbols the context of their mention will make the meaning completely clear.

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XXV

Acronyms

ABM Analogue Behavior Modeling

AC Alternating Current

ACS Attitude Control System

BOL Beginning Of Life

DC Direct Current

DOD Depth of Discharge

DET Direct Energy Transfer

DDU Direct-Drive Unit

EM Engineering Model

EMI Electromagnetic Interference

EOL End of Life

EOR Electric Orbit Raising

EPS Electric Propulsion System

FFT Fast Fourier Transform

FU Filter Unit

GEO Geostationary Earth Orbit

GTO Geostationary Transfer Orbit

GPS Global Positioning System

GSD Ground Sample Distance

HKM Heater Kathode Magnet

HET Hall Effect Thruster

LAE Liquid Apogee Engine

LEO Low Earth Orbit

NEA Near Earth Asteroid

NSSK North-South Station Keeping

PMAD Power Management and Distribution

PPU Power Processing Unit

PV Photovoltaic

PWL Piecewise Linear

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Acronyms

XXVI

S/C Spacecraft

SPT Stationary Plasma Thruster

RMS Root Mean Square

TAL Thruster with Anode Layer

TCS Thermal Control System

TSU Thruster Switching Unit

TWTA Travelling Wave Tube Amplifier

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1

Introduction

Since the early sixties, when the Hall Effect thruster started to be developed both in United States and Soviet Union, it was already clear that this kind of engines, and in general all the electric thrusters, could have ensured large propellant savings with respect to the chemical thrusters. In fact, since the atoms accelerated through the chamber reach exhaust velocity about one order of magnitude higher than in the case of chemical thrusters, this leads to a reduction of the propellant needed for the same velocity change (to the detriment of a longer thrust period). After the launch in 1971 of the first Soviet Hall thruster (SPT-50) the interest and the use of this type of devices increased, first in Russia and then in the Western world since the nineties.

The discharge voltage required by the Hall thruster is in the order of 300 V, but the energy produced by the solar array is generally characterized by a lower voltage (e.g. 28 V, 50 V). Therefore, a Power Processing Unit (PPU) is interposed between the thruster and the solar array in order to step up the voltage and to manage the thruster functions. The removal of the DC-DC converter (anode supply) dedicated to the thruster discharge involves a significant simplification of the PPU. Thus, the Direct-Drive configuration proposes to directly connect the solar array, arranged for producing power at 300 V, to the thruster. Even if there are some technological issues to overcome (e.g. avoid arcing in the solar array), this simplification involves direct benefits not only in the PPU but also in its Thermal Control System (TCS). Indirect advantages can also be exploited by the implementation of high voltage bus, in fact this configuration allows to improve the electrical power system efficiency and thus to reduce its overall mass.

The idea of the Direct-Drive was born in the early seventies in NASA laboratories [58] in relation to the ion thrusters. “For large electrical loads such as ion thruster and high-power

radio frequency amplifiers, the necessary power processors are heavy, complex, and expensive to design and build and are a substantial burden on the spacecraft thermal control system” stated Gooder in 1977. As said, the opportunity to simplify and then

lighten the power processor of the electric propulsion system was and is still now a very attractive opportunity for space missions.

The Comet Halley rendezvous which would be occurred in 1986 pushed the engineers to study a feasible mission strategy with solar electric propulsion; the NASA’s

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30-cm-Introduction

2

diameter gridded ion thruster, directly supplied by a solar array, was a real candidate for such a mission. The task revealed too much demanding for the state-of-the-art technology, however the first experimental tests of a Direct-Drive configuration were fulfilled. The ion thruster was correctly operated at a discharge supply up to 1 A at a beam voltage of 1100 V, the solar array was placed indoor and illuminated with a set of lamps. Gooder highlighted that the tests were successfully accomplished and no issues were detected in thruster operations; in parallel, some studied involving plasma interactions in high voltage solar array were carried out.

After the desertion of the comet Halley mission, in the 1980s the interest in Direct-Drive waned, also because the possibility to develop reliable space solar array, capable to produce more than 1000 V for ion thruster, seemed to be very remote.

But in the first years of the nineties, after the confirmation of the Russian Hall thruster performance, the attention toward the Direct-Drive revived. As already outlined, the Hall thruster needs a discharge voltage of about 300 V; this entails the use of a much more manageable solar array than in the case of the ion thruster.

First in 1997, Hamley et al. [25] tested the 4.5 kW T-160 Hall thruster directly driven by a 1 kW linear concentrator terrestrial solar array placed outdoor. The thruster, firing up to 1 kW at 200 V and 780 W at 300 V, correctly operated during start up and steady state phases.

In 2001, NASA undertook the Direct-Drive Hall Effect Thruster (D2HET) program whose scope was to understand and overcome the issues related to the Direct-Drive implementation. System studies were carried out and the interactions in high voltage solar array immersed in a plasma environment were investigated.

In 2009, the first test using triple junction solar cells is documented. In this case Brandhorst et al. [8] employed a 1.2 kW stretched-lens concentrator solar array to supply the Russian 1.3 kW T-100 Hall thruster. The results proved that the thruster successfully worked up to 600 W and 550 V.

The current interest in high-power solar electric propulsion for human missions toward near-Earth asteroids [9] makes the Direct-Drive concept a natural option for such missions. In this context, NASA decided (in 2011) to set up at JPL an 11 kW solar array test-bed constituted by mono-crystalline silicon cells. A dedicated power control station ensures flexibility of the array performance by varying the series-parallel configuration of the solar panels. In 2012, at JPL, the most extensive and detailed Direct-Drive experimental investigation [58] was carried out in order to understand the several issues identified in previous tests. The Hall thruster utilized was the American H6, capable to

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Introduction

3

reach a discharge power up to 12 kW. The experimental campaign involved the examination of thruster operations in various portions of the solar array I-V curve, different procedures of start-up and shut-down were analyzed and then, also how filter capacitance affects the system oscillations was studied. The thruster correctly operated up to 10 kW and a lot of useful answers came out from these tests.

Arising from the above literature, this thesis has the scope to assess in terms of mass reduction the Direct-Drive implementation impact in spacecraft systems and then, to experimentally demonstrate that the Alta’s HT-100 low power Hall Effect thruster can be effectively supplied by a solar array (in particular a set of thin-film amorphous silicon solar panels capable to provide up to 370 W).

The first part of the thesis is related to the assessment of the Direct-Drive effects on the spacecraft systems; after a brief description of Hall thruster principles of operation (chapter 1), it is illustrated the standard Power Processing Unit (PPU) architecture (chapter 2) and in particular the discharge supply which is its most significant element in terms of mass, inefficiency and complexity. Then, the chapter 3 explains in detail what is and which advantages can be exploited by the Direct-Drive implementation; the mass advantages are subdivided in direct and indirect referring to the ones attainable as a direct consequence of the Direct-Drive and the ones that can be exploited by the adoption of solutions strictly related to the Direct-Drive (e.g. the high voltage bus); in this chapter, also the issues related to such a configuration are qualitatively analyzed. The following chapters (4, 5, 6 and 7) quantitatively assess the mass benefits that the Direct-Drive would affordif implemented in existing (or existed) missions propelled by respectively low (less than 1 kW), medium (1-2 kW) , high (in the order of 5 kW) and very high-power Hall thrusters (more than 10 kW). Here, a remark must be made: the Hall thrusters’ subdivision in terms of power level depends on the current state-of-the-art. In the medium-brief term, this classification will be probably obsolete and the development of very high-power thruster (only prototypes to date) will have an influence also in the subdivision. For instance, nowadays a 5 kW thruster belongs to the class of high-power, in fact there are only one operative mission (AEHF) propelled by Hall thrusters in this range of power. In the next years, this thruster will be probably classified as medium-power thruster and, as the 10 kW (or more) thruster will become more common, they will belong to the high-power category. The chapter 8 shows the conclusions inferred by the analysis of the previous chapters.

The second part of the thesis is related to the Direct-Drive demonstration for the Alta’s HT-100 low-power thruster. The list of the devices, facilities and instrumentation employed

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Introduction

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in the experimental test is displayed in the chapter 9. Then, in the chapter 10, the steps toward the filter design are shown: circuit simulations with Pspice software and HT-100 tests with laboratory power supply were carried out in order to achieve satisfactory filter performance. The chapter 11 describes the effort in understanding and evaluating the influence of several factors such as the solar irradiance, cells temperature, and losses, involved in the solar array output performance. The chapter 12 illustrates the procedure and the results of the experimental test which has successfully demonstrate the HT-100 Direct-Drive operations. Finally the chapter 13 explains the conclusions derived from the experience of this work.