Potential Propulsion System for Microsatellites

Potential Propulsion System for Microsatellites Presented by:

Tareq Bin Ali,#090324622 Dissertation Submitted In partial fulfillment of the: Master of Science (MSc) in Aerospace Engineering

Supervisor: Dr Kate Smith Lecturer in Aerospace Engineering School of Engineering and Materials Science Queen Mary, University of London London, United Kingdom Date: 27th August 2010 Word count: 12,944

I confirm that the contents of this report are entirely my own work and that nothing has been included from other sources without acknowledgement or reference.

Abstract Microsatellites have become increasingly popular with the advancement in microfabrication and computing technologies. In order to take advantage of this highly efficient, cost effective technology, on board micro-propulsion systems has to be developed. This research aims to identify the potential propulsion systems for the near future missions such as LISA, IXO and TPF. Due to the mission constraints, the on-board micropropulsion unit has to provide precise control and high accuracy. Both chemical and electrical propulsion systems have been studied. Colloid propulsion system has been found to be the most promising technology because of their miniature design and capability of providing thrusts in micronewton level. Finally, a low thrust lunar CUBESAT has been proposed. The feasibility study shows that colloid thrusters can be successfully implemented in such missions, thus opening the opportunity to apply electrospray in microsatellite applications.

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Acknowledgement I am really excited that I am writing this long awaited part of my dissertation. It has been a long journey, I suppose. I would like to take the opportunity to thank all those people who helped me to complete this journey. First of all, I would like to thank my project supervisor Dr Kate Smith, who was kind enough to agree to advise me on this project. Her immense knowledge in the colloid thruster technology, enthusiasm, promptness and the willingness to help have been a great help for me to complete this thesis. My words cannot express my gratitude and appreciation for all the support, guidance and time you provided. I would like to thank Professor Stark, Professor Vepa, Dr Duddeck and Professor Munjiza for making the lessons so attractive. I am also grateful to Dr M Hasan Shaheed for his invaluable advice during my post graduation in Queen Mary. A huge thanks to you for giving me the offer to work with you for my PhD. My heartfelt thanks to Alam, my roommate and also my best friend for years. Thank you for tolerating my temper for long seven years! Also I am grateful to my best mate Arif; without his help I could not possibly complete my studies. Baccha apu, you have inspired me to take this course of study. I wanted to be an Aerospace Engineer like you and here I am! Lets plan about that trip to moon! Reshma apu, you are the loveliest and simplest sister in the world. I cannot express how much I will miss you. Thank you Sona for all those messages wishing me luck. Osmosis worked! “ďakujem!” I would also like to thank my colleague Purushoth for his continuous support during the project. I made a habit of working with you! I will miss all those offline messages in gmail. You better keep them coming! Peter, I thank you for all those sneaky tea breaks, and particularly for all those random discussions about life, women, space and whatever we talked about! Last, but very far from least, a huge thanks to my great family. Ammu and Bapi, you are the greatest parents one could ever have. Thanks for always believing in me. I am greatly indebted to my younger brothers-Antu and Soumik for their profound love for me. You guys are the sweetest!

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Table of Contents Abstract ................................................................................................................................................... 2 Acknowledgement .................................................................................................................................. 3 List of Tables ........................................................................................................................................... 7 List of Figures .......................................................................................................................................... 9 List of Abbreviations ............................................................................................................................. 11 Nomenclature ....................................................................................................................................... 13 Chapter 1............................................................................................................................................... 15 1.1

Prospects of Microsatellites .................................................................................................. 15

1.2

Requirements of Onboard Propulsion .................................................................................. 16

1.2.1

Orbit Insertion ............................................................................................................... 16

1.2.2

Station keeping and Drag reduction ............................................................................. 17

1.2.3

De-commissioning ......................................................................................................... 18

1.2.4

Orbit Phasing ................................................................................................................. 19

1.2.5

Attitude Control ............................................................................................................ 19

1.3 1.3.1

Missions Requiring Micro- propulsion .................................................................................. 20 LISA .................................................................................................................................... 20

Mission Overview .......................................................................................................................... 20 Propulsive Requirements .............................................................................................................. 22 1.3.2

International X – ray Observatory (IXO) ........................................................................... 23

Overview ....................................................................................................................................... 23 Propulsive Requirements .............................................................................................................. 24 Page 4 of 92

1.3.3

Terrestrial Planet Finder Interferometer (TPF –I ) ............................................................ 25

Overview ....................................................................................................................................... 25 Scientific Requirements ................................................................................................................ 26 1.2.4

Summary ........................................................................................................................... 27

1.3

Aims of the Research Project ................................................................................................ 28

1.4

Methodology ........................................................................................................................ 29

Chapter 2............................................................................................................................................... 30 2.1

Available Propulsion Technologies ...................................................................................... 30

2.2

Chemical Propulsion Technology ......................................................................................... 33

2.2.1 2.3

Microsatellite Gas Propulsion System .......................................................................... 33 Electric micro propulsion systems ....................................................................................... 35

2.3.1

SSTL Low Power Resistojet ............................................................................................ 36

2.3.2

Ion thrusters .................................................................................................................. 38

2.4

Performance and Operating Characteristics of Electric Propulsion ..................................... 47

2.5

Feasible Propulsion for microsatellites ................................................................................. 49

Chapter 3............................................................................................................................................... 50 3.1

Physics of Colloid Propulsion ............................................................................................... 50

𝑬𝑴𝑰 − 𝑩𝑭𝟒 .................................................................................................................................. 50 𝑬𝑴𝑰 − 𝑰𝒎 .................................................................................................................................... 51 3.1.1

Surface Charge .............................................................................................................. 52

3.1.2

Taylor Cone: .................................................................................................................. 52

3.1.3

Starting Voltage: ........................................................................................................... 57 Page 5 of 92

3.2

Related work in the area of Electro-spray ............................................................................ 58

3.3 Developments in colloid propulsion ........................................................................................... 59 3.3.1 3.4

Colloid Propulsion Research at Queen Mary ................................................................ 66

Applicability of Colloid thrusters ........................................................................................... 70

Chapter 5............................................................................................................................................... 73 5.1

Proposed Lunar CUBESAT Mission ........................................................................................ 73

5.2

Power Requirement ............................................................................................................. 74

5.3

Communication ..................................................................................................................... 74

5.4

Attitude Control System........................................................................................................ 75

5.5

Total Mass budget ................................................................................................................ 76

5.6

Propulsive Requirement ....................................................................................................... 77

Chapter 6............................................................................................................................................... 81 Conclusions and future work ............................................................................................................ 81 Appendix ............................................................................................................................................... 82 A.

Link Budget................................................................................................................................ 82

References ............................................................................................................................................ 85

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List of Tables

Table 1.1 Classification of satellites ……………………………………………………… .16 Table 1.2 Colloid Micropropulsion Requirements ………………….

……………….23

Table 1.3 ∆ 𝑉 Budget for IXO ……………………………………………………………….25 Table 1.4 Propulsion requirements for a typical formation flying mission…………………..28 Table 2.1 Specifications of Xenon Gas Propulsion system….……………………………….35 Table 2.2 Specifications of Low power Resistojet thruster ………………………………….37 Table2.3 Materials used for Laboratory vs. material to be used for on flight model …. …. ..40 Table2.4 Performance and Operating Characteristics of Electric propulsion systems ……...48 Table 3.1 Specification of MAI colloid thruster developed at MAI ……………….………..60 Table 3.2 Thrust and Thrust variance vs. applied voltage…………………………….……..64 Table 3.3: Colloid thrusters developed to date ………………………………………………66 Table 3.4 Proposed thruster flight experiment …………………………..………………….71 Table 5.1 General Specifications of the spacecraft

……………………………………….74

Table 5.2 Power Requirements…………………………………………………………… .75

Table5.3 Transceiver Specification………………………………………………………….76 Table 5.4 Operating Characteristics of the reaction wheel

………………………………..76

Table 5.5 Spacecraft total mass budget ……………………………………………………...77 Page 7 of 92

Table5.6 Propulsive Requirements

………………………………………………………..79

Table 5.7 Fuel Tank Size Estimation ………………………………………………………..81

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List of Figures Figure1.1 Transferring the s/c into final orbit through an elliptical transfer orbit ……………17

Figure 1.2 Protected Region ……………………………………………... …………… … 18 Figure1.3 LISA orbits …………………………

…………… …………… ……………19

Figure 1.4 Earth Analog spectrum ………………. …………… …………… ……………. 26 Figure 2.1 Schematic of a rocket device …………… …………… …………… ………….30 Figure 2.2 SSTL Xenon Gas propulsion system

…………… …………… ……………34

Figure 2.3 Low power Resistojet…………………………………………..………………...37 Figure2.4 MRIT size comparison ………………………………… …………… …………..39 Figure2.5 MRIT system diagram …………………... …………… …………… …………..39 Figure 2.6 Two-Dimensional beam current density profile …………………………………41 Figure 2.7 MRIT thrust vs. time …………………... …………… …………… …………...41 Figure2.8 MRIT mass efficiency vs. thrust at multiple propellant flow rates …..…………..42 Figure 2.9 RIT- μX elegant breadboard ……………………………………….…………….44 Figure2.10 RIT-μX Performance, Specific Impulse as function of total power and thrust level ………………… …………… …………… …………… ………..…… …….45 Figure 2.11 RIT- μX 50μN Thrust Stepping ………………… …………… …..…………..46 Figure 2.12 Thrust Stepping -wide range …………………. …………… …….…………...46 Figure 3.1: Schematic of a colloid thruster …………… …………… ……………………..51 Figure 3.2 : Charge Concentration change in Electric conductor …………… …………….52 Figure 3.3 Cone jet structre for Ethylene – Glycol ………………… …………… ………...53

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Figure3.4 Taylor cone geometry with an inner angle 𝛼. …………… …………… ………54 Figure3.5 Spherical Coordinate system …………… …………… …………… ………….55 Figure 3.6 Plot of Legendre polynomials ………………. …………… ……………………56 Figure 3.7 Prototype of the 100-nozzle thruster ………………… …………… …………..61 Figure 3.8 Testing arrangement of prototype …………… …………… …………………. 62 Figure 3.9 Capillary Geometry …………… ……………..……… …………… ………….67 Figure3.10 Schematic cross section of the colloidal thruster …………… ………. ………..68 Figure 3.11 Current vs. voltage curve- 25 𝜇𝑚 spacing …………………… ……………….69 Figure 3.12 Current vs. Voltage –extractor and emitter distance 25 𝜇𝑚 …. …………… …. 69

Figure 3.13 Hybrid Colloid Thruster …………… ………………..… …………… ………70

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List of Abbreviations ADCS

Attitude Determination and control subsystem

AU

Astronomical Units

EJSM

Europa Jupiter System Mission

ELITE

European Lisa Technology Experiment

𝐸𝑀𝐼 − 𝐵𝐹4

1 – ethyl – 3 – methylimidazolium tetrafluoroborate

𝐸𝑀𝐼 − 𝐼𝑚

1-ethyl-3-methyllimidazolium bis (triflouromethylsulfonyl)

ESA

European Space Agency

EX-5

Earth Science Experimental Mission 5

FCU

Flow Control Unit

FEEP

Field Emission Electric Propulsion

GSO

Geostationary Orbit

GSTP

Gaia Science Team Program

IADC

Inter – Agency Space Debris Coordination committee

𝐼𝑆𝑃

Specific Impulse

IXO

International X-Ray Observatory

JAXA

Japan Aerospace Exploration Agency

JPL

Jet Propulsion Laboratory

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LEO

Low Earth Orbit

LIRE

Laser Interplanetary Ranging Experiment

LISA

Laser Interferometer Space Antenna

LV

Launching Vehicle

MAXIM

Micro Arcsecond X-ray Imaging Mission

MEMS

Micro-electrical and Mechanical Systems

MRIT

Miniature Radio-Frequency Ion Thruster

NAI

Sodium Iodide

PM

Propulsion Module

PPT

Pulsed Plasma Thruster

PPU

Power Processing Unit

RFIT

Radio-Frequency Ion Thruster

s/c

Spacecraft

SMART

Small Missions for Advanced Research in Technology

SPECS

Sub-millimeter Probe of the Evolution of Cosmic Structure

SSTL

Surrey Satellite Technology Limited

ST-3

Space Technology 3

TPF-I

Terrestrial Planet Finder Interferometer

UHF

Ultra High Frequency Page 12 of 92

Nomenclature 𝑎

𝐴

Specific power of the power plant Surface Area

𝑐

Effective exhaust velocity

𝐶𝑟

Coefficient of solar radiation pressure

𝐸𝑛

𝐸𝑡𝑖𝑝 𝑓

𝑓𝑠𝑡

Electric field Electric potential at the tip of the conical surface Frequency

Surface tension of the liquid Gravitational acceleration

m

Mass

𝑚𝑝

𝑀0 𝑀𝑒 𝑃

𝑃𝑒

𝑃𝐸

𝑃𝑎 𝑃𝑟

Structural Mass Initial mass of the system Final mass of the system Kinetic power of jet Pressure in the exhaust area Electrical power Atmospheric pressure Momentum of the mass

Q

volumetric flow rate

r

radius

𝑅𝑐

Principal surface radius of the

𝑡𝑝

Time of operation or propulsive

T

Thrust

Mass flow rate

𝑚𝑝𝑎𝑦 Payload mass

𝑚0

𝑚𝑆

Propellant mass

Varying system

𝑔

𝑚̇

𝑚𝑝

Initial mass of the spacecraft Propellant mass

𝑚𝑝𝑜𝑤𝑒𝑟 Power plant mass

curvature

time

∆𝑉

Change of velocity

𝑣1

Initial velocity

𝑣2

Final orbit velocity

𝑣𝑐

𝑣𝑒

Characteristic speed Exhaust velocity relative to the vehicle

V

Voltage

𝑉𝑔

Gravitational speed loss

𝜏𝑏 𝜂

𝜌𝑠

𝜆

𝜀0 𝛾

Burn Time Thruster Efficiency Charge per unit area Wavelength Permittivity in vacuum Surface tension of the liquid

∆ 𝐻𝐺𝐸𝑂 Height above GEO for safe Decommissioning

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Chapter 1 1.1

Prospects of Microsatellites The idea of microsatellites is not a very recent one. Due to the limited lifting

capability of launch vehicles, spacecrafts were traditionally lighter and smaller. Vanguard 1 (85 kg), Explorer 1 (15 kg) and the first earth orbiting satellite Sputnik-1 (85 kg) provide good examples of this. With the advancement of aerospace technology, launching vehicles with heavy lifting capability have been developed (Ariane 5, Proton). During the 1980s and 1990s there was a trend of sending big satellites in space. It became a symbol of superiority of the then superpowers. However, advances in micro-fabrication and computing technologies have changed the scenario. Micro-electrical and Mechanical Systems (MEMS) have a great potential in implementing the ideas of micro and pico satellites. Recent advancement in the electronics industries has increased the functionality of the smaller spacecraft. Moreover, the budget constraints and recent government policies have resulted in a trend of decreasing satellite mass. The reduced cost of production and placing them into the orbit played a vital role in the recent advancement of microsatellites. Microsatellites make it possible to distribute various functionalities of a single spacecraft to a number of smaller satellites. This minimizes the risk and improves the reliability of the system and, at the same time, reduces the individual satellite production cost. The reduction of complexity of individual spacecraft enables rapid prototyping and also lowers the development life-cycle (Khayms, 2000).

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Table 1.1 presents the classification of satellites according to their wet mass (mass including fuel): Table 1.1 Classification of satellites Mass Category >1000 kg Large Satellites 500-1000 kg Medium Satellites 100-500 kg Mini Satellites 10-100 kg Micro Satellites 1-10 kg Nano Satellites 0.1-1 kg Femto Satellites

1.2

Requirements of Onboard Propulsion On board propulsion system is essential for various corrective manoeuvres like orbit

insertion, phasing, station keeping, drag reduction and decommissioning of spacecraft. Missions like Laser Interferometer Space Antenna (LISA) and International X-Ray Observatory (IXO) require precise attitude control of the spacecraft because of the subtlety of formation flying. 1.2.1

Orbit Insertion

The satellite is normally launched with a launching vehicle (for example, ARIANE, VEGA, SOYUZ etc). The spacecraft can be placed into the final orbit or it can be placed into a parking orbit. In the later case the onboard propulsion system has to be used to place the s/c Page 16 of 92

into the desired orbit. If the satellite (Figure 1.1) has a velocity of 𝑣1 in the initial orbit and if

the velocity in the final orbit (which is inclined to the plane by 𝑖°) then the thruster has to provide a ∆𝑉 change which is given by the following expression: ∆𝑉 = 𝑣12 + 𝑣22 − 2 𝑣1 𝑣2 cos 𝑖

1.1

∆V1

Transfer orbit apogee ∆V2

Inclined final Orbit

Figure1.1 Transferring the s/c into final orbit through an elliptical transfer orbit.

Even if the launching vehicle delivers the s/c into its final orbit there may be some error in the inclination or may be the velocity will not be in the required level. In that case the thruster has to provide the necessary ∆𝑉 (which can be obtained from equation 1.1). 1.2.2

Station keeping and Drag reduction

The satellites in the geostationary orbit (GSO) or in low earth orbits (LEO) are exposed to various perturbations like atmospheric drag, oblateness of earth and gravitational forces from sun and moon. The gravitational pull by the sun and the moon increases the satellite inclination about 1° per year. So to maintain the desired inclination with the Page 17 of 92

equatorial plane, north - south station keeping manoeuvre takes place about fortnightly. The satellite has to perform another type of corrective manoeuvre using its onboard propulsion which is known as east – west station keeping. This perturbation occurs due to the oblateness of earth and the direction of perigee rotates around the orbit. Atmospheric drag causes the orbit gradually decays to result into a re-entry. 1.2.3

De-commissioning

Once the mission lifetime is over, necessity may rise to decommission the spacecraft especially if it is a part of a constellation (e.g. GPS). Satellite decommissioning has two phases. In phase one the satellite altitude is raised as high as possible. Secondly, the satellite subsystems are reconfigured (Venting the pressure vessels, discharge of electrical energy or dumping any source of kinetic energy) to minimize the collision effect with micrometeorites or any other object.

Figure 1.2 Protected Region (Inter – Agency Space Debris Coordination committee, 2002)

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Inter – Agency Space Debris Coordination committee (IADC) provided an equation to determine the height above GEO arc that the satellite should be raised during the decommissioning process for minimum risk of collision (Hope, 2007): ∆ 𝐻𝐺𝐸𝑂 (𝑘𝑚) = 235 + 1000 × 𝐶𝑟 ×

𝐴

𝑀

1.2

Where, 235 = 𝑡ℎ𝑒 𝑠𝑢𝑚 𝑜𝑓 𝑡ℎ𝑒 𝑢𝑝𝑝𝑒𝑟 𝑝𝑟𝑜𝑡𝑒𝑐𝑡𝑒𝑑 𝑟𝑒𝑔𝑖𝑜𝑛 𝑜𝑓 𝐺𝐸𝑂 𝑎𝑛𝑑 𝑡ℎ𝑒 𝑚𝑎𝑥𝑖𝑚𝑢𝑚 𝑑𝑒𝑠𝑐𝑒𝑛𝑡 𝑓𝑟𝑜𝑚 𝑙𝑢𝑛𝑖 𝑠𝑜𝑙𝑎𝑟 𝑝𝑢𝑟𝑡𝑢𝑟𝑏𝑎𝑡𝑖𝑜𝑛𝑠 (Figure1.2)

𝐶𝑟 = 𝑐𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡 𝑜𝑓 𝑠𝑜𝑙𝑎𝑟 𝑟𝑎𝑑𝑖𝑎𝑡𝑖𝑜𝑛 𝑝𝑟𝑒𝑠𝑠𝑢𝑟𝑒 (𝑡𝑦𝑝𝑖𝑐𝑎𝑙𝑙𝑦 1 ≤ 𝐶𝑟 ≤ 2)

1.2.4

Orbit Phasing

𝑚2 𝐴 = 𝐴𝑠𝑝𝑒𝑐𝑡 𝑎𝑟𝑒𝑎 𝑡𝑜 𝑑𝑟𝑦 𝑚𝑎𝑠𝑠 𝑟𝑎𝑡𝑖𝑜𝑛 ( ) 𝑘𝑔 𝑀

When the satellite has to intercept any other target objet of interest, both the s/c and the target must be in the same rendezvous point at a given time. The phasing manoeuvre involves a 2-impulse Hohmann transfer to bring the satellite out and to place it on the same orbit but at a different point. This can be used to place a satellite into a new position in its previous orbit. For example, a communications satellite in GEO can use Phasing manoeuvre to gain a different altitude which enables it to cover a new area. 1.2.5

Attitude Control

Attitude Determination and control subsystem (ADCS) is an important part of the spacecraft. ADCS maintains the desired orientation of the s/c by cancelling the external perturbations. Although magnetic torquers are used most of the times to cancel the torque produced and to stabilize the spacecraft, if the mission profile requires (e.g. LISA) precise control then micro-newton thruster has to be used.

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1.3

Missions Requiring Micro- propulsion

1.3.1 LISA Mission Overview

LISA is the second space mission of the European Space Agency’s Small Missions for Advanced Research in Technology (SMART) programme. In 1998 the mission was proposed as European Lisa Technology Experiment (ELITE). The proposed mission was to launch a satellite in geostationary orbit. The goal of the mission was to achieve a differential acceleration of 10−14 𝑚𝑠 −2 /√𝐻𝑧 in the frequency range between 1 – 100 𝑚𝐻𝑍. With the announcement of SMART programme, the proposal was refined to launch two spacecrafts

called LISA and DARWIN. However, the DARWIN Pathfinder was cancelled after an initial feasibility study. LISA Pathfinder will carry a European Space Agency (ESA) built Technology package and a NASA built Technology Package. The LISA mission is designed to observe and study the gravitational waves from different gravitational wave sources, such as massive black hole binaries, intermediate- mass black holes, stellar – mass compact objects, close binaries of stellar – mass compact objects; over the frequencies from 0.03 milliHertz to 0.1 Hertz (NASA, 2010). It is not possible to carry out this measurement in this frequency band on earth due to ground motion and time variations in gravity from mass motions on the earth. The satellite is scheduled to launch in 2011 from an ESA VEGA launcher form the French Guyana (Kourou) facility. The launcher will place the spacecraft into a low earth parking orbit with a semi major axis of 1820 km and the inclination of 5.3°.

A detachable thruster module will be used to perform a number of apogee raising manoeuvres

to place the satellite in a transfer orbit towards L1 (the first Earth – Sun Lagrange point). LISA will enter the final Lissajous orbit around L1 using the onboard micropropulsion system (McNamara, 2009).

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LISA mission will consist of three spacecraft which will orbit in a near-equilateral triangular formation. LISA interferometer will have an arm length of five million kilometres. So the strain sensitivity will be improved by the amplification effect of low frequency gravitational waves (0.1 mHz – 1 Hz) due to longer arms (Shaddock, 2008).

Figure1.3 LISA orbits (Danzmann et al., 2007) The constellation will lie in a plane which is inclined to ecliptic by an angle of 60° so

that the spacecrafts’ relative has a period of one year and it will be trailing the earth by 20° (constrained by the launch vehicle capability) (Reichbac, 2001). Figure 1.3 depicts the

orbit of LISA. Three spacecraft are denoted in the constellation by dots. Ecliptic is the thick line in the snapshot. The circle running through same dot is the desired orbit of each spacecraft. Page 21 of 92

Gravitational waves are studied based on the effect on motion of an object. The object masses which need to be measured are known as “proof masses”. The displacements of the proof masses are measured by laser interferometer. LISA will have two proof masses in each of the spacecraft. The acceleration noise of the proof masses must be approximately 10−15

𝑚/𝑠 −2 √𝐻𝑧

. To achieve this goal, proof masses are shielded by the spacecraft from solar

wind and solar radiation. In order to detect gravitational waves, each spacecraft emits two phase-locked laser beams simultaneously in the direction of the other two spacecraft. So the spacecraft in the vertices can track each other. The two lasers with different frequencies will produce a beat note as because of interference. A phase shift of one cycle is produced in the beat not if the path length changes by one optical wavelength. So the phase of the beat note indicates any change in displacement. The range of beat note frequency for LISA is 2 𝑀𝐻𝑧 − 20 𝑀𝐻𝑧

(Shaddock, 2008).

Propulsive Requirements

The spacecraft is launched with an ESA VEGA launcher. The Launching Vehicle (LV) will place the spacecraft into a low earth parking orbit with a semi major axis of 1820 km and an inclination of 5.3°. The micro-thruster unit has to support the spacecraft bus and payloads during ground operations. During launch, the Propulsion Module (PM) acts as the

primary load path for the spacecraft. The micropropulsion unit has to provide the required ∆ 𝑉 to place the spacecraft in the desired science orbit from its initial parking orbit. The

duration requirement of orbit transfer is 15 months. The payload has to be delivered to the operational orbit within this time limit. Moreover, the PM should be able to alter the attitude and orbit of the spacecraft throughout the transfer period. The total ∆ 𝑉 budget for the mission is 1139 𝑚/𝑠 (NASA Report, 2009).

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Table1.2 presents the main performance criteria of the colloid thruster onboard LISA. The maximum thrust (30𝜇𝑁) is calculated from the ∆ 𝑉 required to cancel solar radiation. Thrust resolution has to be less than ≤ 0.1 𝜇𝑁 and the thrust noise must be ≤ 0.1 𝜇𝑁⁄√𝐻𝑧 for effective study of gravitational waves. The specific impulse needed is ≥ 150 𝑠𝑒𝑐. Table 1.2 Colloid Micropropulsion Requirements (McNamara et al., 2009) Propulsion Parameter

Colloid Thruster Requirement 5 − 30𝜇𝑁

Thrust Range

≤ 0.1 𝜇𝑁

Thrust Resolution

≤ 0.1 𝜇𝑁⁄√𝐻𝑧

Thrust Noise

≤ 100 𝑠𝑒𝑐

Thrust Response Time

≥ 150 𝑠𝑒𝑐

Specific Impulse

25 𝑊

Cluster power consumption(@30𝜇𝑁)

14.6 𝑘𝑔

Cluster Mass

90 𝑑𝑎𝑦𝑠

Lifetime (Thruster ON)

300 𝑁𝑠

Total Impulse

1.3.2 International X – ray Observatory (IXO) Overview

The IXO mission is a collaborative mission by ESA, NASA and Japan Aerospace Exploration Agency (JAXA). It is the merger between two previously proposed missions called XEUS (The X-Ray Evolving Universe Spectroscopy Mission) and Constellation-X. The mission is aimed to study the evolution of universe. The influence of black holes in the formation and expansion of galaxy will be studied. The behaviour of matters under strong

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gravity and at a high density will be observed. IXO will also investigate various astrophysical phenomena such as cosmic rays of supernova and planet and star formation (Rando, 2010). Propulsive Requirements

The IXO will be launched with the Ariane 5 ECA or Atlas V 551 launcher. Total launch mass is expected to be around 6500 kg. The final operating orbit will be around L2, the second Lagrangian point of the Sun - Earth system. The main advantage of an L2 halo orbit is its thermally stable environment and a good visibility of sky. The satellite will be injected to the transfer orbit towards L2 and after getting the tracking data (two days after injection) a corrective manoeuvre will be performed to minimize the launcher dispersion error. Another corrective manoeuvre will take place after 10 days of injection. The telescope will be deployed after the second corrective manoeuvre which will enable the commissioning phase of the satellite. The spacecraft will arrive in its final halo orbit approximately after 100 days after launch. Station keeping will be performed monthly which will require a total ∆𝑉

of 2 𝑚⁄𝑠 /𝑦𝑒𝑎𝑟. The mission is expected to have a lifetime of 10 years requiring a total ∆𝑉

of approximately 120 𝑚⁄𝑠 (Rando, 2010). Table 1.3 presents ∆𝑉 requirements for various manoeuvres during the mission lifetime of 10 years.

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Table 1.3 ∆ 𝑉 Budget for IXO (ESA, JAXA report, 2008) Perigee velocity correction

∆𝑉 requirement (𝑚𝑠 −1 )

Launcher dispersion correction

34.0

Transfer

3.0

Orbit insertion

0.0

Station keeping (10 years)

20.0

de-orbit

0.0

Wheel off-loading correction

7.7

Total

85.0

Margin (5%)

4.63

Thruster mounting

11.7

Total ∆ 𝑉 Budget

109.4

Manoeuvre

28.0

1.3.3 Terrestrial Planet Finder Interferometer (TPF –I ) Overview

The Terrestrial Planet Finder Interferometer is a mission to identify habitable planets like Earth around nearest stars. The anticipated launch is between 2012 –2015. TPF-I will study the planets using space-borne telescopes which will be more effective with a high resolution interferometer. The mission aims to study the planets outside our solar system in several ways. The formation of planets, their properties, evolving disks of newly forming stars and the possibility of existence of life will be studied during the mission. TPF-I will perform this task by identifying and analyzing the molecular lines from thermally emitted and

Page 25 of 92

reflected light from distant planets. It is expected that TPF-I will be able to analyze planets within 5 Astronomical Units (AU) of nearby stars and which are also within the 10 –𝜇𝑚 view of the interferometer. TPF-I will also look for various gaseous elements in the mid-IR which indicate biological existence. Figure 1.4 shows that the 𝑂3 band, 𝐶𝑂2 band and 𝐻2 𝑂 band are

the three strongest bands in the earth-analog spectrum. All these can be detected with the high

spectral resolution of TPF-I. The mission will be designed with a lifetime of 5 years, but can be extended to a 10 years programme depending on the quality of the data obtained from the mission. The satellite will be launched using an Ariane 5 ECA or equivalent launcher. The final orbit will be an L2 Halo orbit (Lawson, 2007).

Figure 1.4 Earth analog spectrum (Lawson, 2007) Scientific Requirements

The primary goals of TPF-I in Pre Phase A is to demonstrate mid-infrared nulling and to demonstrate reliability for formation flying. The level of nulling required by the mission is at a level of 1 × 10−5 . As every telescope will have delay line of several tens of centimetres Page 26 of 92

or may be in the range of a meter, the relative range control between spacecraft has be ≤ 5 𝑐𝑚. TPF-I must be able to determine relative positioning to ±1 𝑐𝑚. Relative velocity

and attitude information must be respectively within ±1 𝑚𝑚⁄𝑠 and ±1 𝑎𝑟𝑐𝑚𝑖𝑛 (Reichbach, 2001). As the mission is in pre development stage, most of the mission requirements are still

being researched. However, it is clear from the type and functionality of the mission that TPF-I will need very precise and accurate level of thrust to ensure an accurate position control to carry out the mission goals. The thrusters on board TPF-I satellites have to be able to provide the required ∆𝑉 for a mission lifetime of five years (or possibly ten years).

Moreover, the thrust resolution and thrust level has to be within the range of micro – milli newton.

1.2.4 Summary A reliable and high performance miniature propulsion subsystem is the prerequisite for the near future missions like LISA, IXO, Europa Jupiter System Mission (EJSM). Missions involving high precision formation flying (e.g. LISA) demand a propulsion system with very high accuracy and very low noise to thrust ratio. Table 1.4 presents typical mission requirements for near future missions. From the data represented on the table it is clear that the propulsion system should be capable of providing thrust in the range of micro Newton to milli-Newton. Thrust resolution also becomes a decisive factor in micropropulsion system design. Thrust resolution is the smallest increment of thrust that can be commanded by the control system of the thruster. Near future missions require this thrust resolution to be less than 0.5 μN. Thrust noise is expected to be 1.65μN/√Hz up to 100μN/√Hz. Moreover, the 𝐼𝑆𝑃

of the system should be higher than 1500s. Lifetime of the propulsion system has to be very high ( ~ 21900 hours). It is quite obvious from all these mission requirements that the propulsions systems that are available at present (e.g. Cold gas thruster and Resistojet thruster) are unsuitable for small, micro and pico satellite missions. Page 27 of 92

Table 1.4 Propulsion requirements for a typical formation flying mission (Collingwood et al., 2009).

1.3

Thrust range (fine)

1 – 150μN

Thrust range (coarse)

150μN to >1mN

Thrust resolution

90s @12μN

Total impulse

40kNs

Beam divergence

× 4

System Volume

7.42 litres

Life duration

>7 years

2.3

Up to 48 sec

Electric micro propulsion systems Unlike chemical propulsion systems, electric thrusters have different

operating principle. They are limited by power. The exhaust velocity and specific impulse is directly related to the power supplied to the device. So it is very important for the overall design to ensure large amount of power supply without the demand of huge power supplies. However, for most of the state of the art electrical propulsion devices offset the propellant mass saving due to higher exhaust velocity by the enormous power supply. The performance of an electrical propulsion device can be analyzed in terms of mass and power. Let, 𝑚0 =Initial mass of the spacecraft 𝑚𝑝 = propellant mass

𝑚𝑝𝑎𝑦 = payload mass and,

𝑚𝑝𝑜𝑤𝑒𝑟 = power plant mass.

Page 35 of 92

So the initial mass can be expressed as: 𝑚0 = 𝑚𝑝 + 𝑚𝑝𝑎𝑦 + 𝑚𝑝𝑜𝑤𝑒𝑟

2.14

Specific power of the power plant, the ratio of the electrical power 𝑃𝑒 to the mass of the power plant 𝑚𝑝𝑜𝑤𝑒𝑟 , is defined as: 𝛼 =

𝑃𝑒

2.15

𝑚𝑝𝑎𝑦

If the efficiency of the thruster is 𝜂 then the electrical power input is: 𝑃𝑒 = 𝛼 𝑚𝑝𝑜𝑤𝑒𝑟𝑝𝑙𝑎𝑛𝑡 =

1 𝑚̇𝑣 2 2

𝜂

=

𝑚𝑝 𝑣 2 2 𝜂 𝑡𝑝

2.16

where, 𝑡𝑝 is the time of operation or propulsive time. Equation13, 14, 15 can be used to obtain the following relation for the payload mass fraction: 𝑚0

=

𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑

𝑒 ∆𝑢⁄𝑣

�𝑒∆𝑢⁄𝑣 − 1�𝑣2

1−

2.17

2 𝛼 𝜂 𝑡𝑝

Characteristic speed 𝑣𝑐 is given by:

𝑣𝑐 = �2 𝛼 𝜂 𝑡𝑝

2.18

For a given payload fraction (

𝑚0

𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑

) and characteristic speed (𝑣𝑐 ), an optimum

range of specific impulse can be obtained which can be used for an optimum propulsion system design. 2.3.1

SSTL Low Power Resistojet

SSTL low power Resistojet is designed for applications like orbit correction and station keeping of small satellites. It can be used as an augmentation to a compressed gas or liquefied gas thruster to improve the specific impulse. Specific impulse of the thruster varies Page 36 of 92

depending on the type of propellant, power, firing time and the level of thrust. Table 4 shows that the typical ISP for a Xenon propulsion system is 55 sec while the performance improves (99 sec) with the use of Nitrogen. The wound heater coils inside the thrust chamber (figure 3) heat up the propellant up to 500℃. The thruster needs a power supply of 50 W at 28 Vdc. It can be operated from the bus voltage.

Figure 2.3 Low power Resistojet (SSTL subsystem Datasheet, 2010) Table 2.2 Specifications of Low power Resistojet thruster (SSTL subsystem Datasheet, 2010) Propellant

Nitrogen, Xenon, Butane and most gases

Thrust

≤ 100 mN

Feed Pressure Specific Impulse

100 bar

𝑋𝑒𝑛𝑜𝑛 − 55 𝑆𝑒𝑐 𝑁2 − 99 𝑠𝑒𝑐

Operation Temperature

𝐵𝑢𝑡𝑎𝑛𝑒 − 100 𝑠𝑒𝑐

Mass

500℃

65 gms without valves

Heater Power

50 watts @ 28 Vdc

Operating temperature

−20℃ 𝑡𝑜 + 60℃ Page 37 of 92

2.3.2

Ion thrusters

Ion thrusters work on the principle of accelerating the heavy ions, created in an ionization chamber, to very high exit velocities. Miniature Ion Propulsion devices can be used to provide finite attitude control and also is suitable for missions with high specific impulse requirements. They can also be used for routine satellite station keeping and attitude control for formation flying. They can also be used as primary propulsion devices of micro satellites. They are of high operational efficiency and fuel consumption is very low. 2.3.2.1

Development of Miniature Radio Frequency Ion Thruster (MRIT)

Trudel et al. (2009) worked on the development of a Miniature Radio-Frequency Ion Thruster (MRIT). The title “Design and performance testing of a 1-cm Miniature Radio Frequency Ion Thruster” gives the reader a very clear idea what they included in their report. The abstract was very well constructed. It gave a clear idea of the research work. It expressed the primary goal of the MRIT program which was to design a smaller, micro Newton range RF ion propulsion thrusters to precise attitude control of satellites and to use as a primary propulsion device in micro-satellites. The type of experiment and its outcome was briefly mentioned in the abstract which gave the reader a bird’s eye view of the research. MRIT is a promising device to ensure finite attitude control of spacecrafts requiring precision control. Moreover, it provides high Specific Impulse (𝐼𝑆𝑃 ) and high operational

efficiency with a very low fuel consumption rate. A typical MRIT could produce very low levels of thrust in the range of 1 μN - 50 μN at a precise thrust resolution (4μN-10 μN). Hence, MRIT is very effective in attitude control of formation flying spacecraft.

Page 38 of 92

Figure2.4 MRIT size comparison (Trudel et al. 2009)

2.3.2.2 Experimental Set up

The experimentation was a continuation of the previous work where they used a cylindrical MRIT thruster with a Plasma Chamber of 1.25 cm both in diameter and length. The maximum thrust they gained was 75 μN with an 𝐼𝑆𝑃 of 2400 s. In the latest experiment,

they used a conical Plasma Chamber which was 1.0 cm in both diameter and length. The thruster length was just over 2.0 cm. A schematic of the diagram of the MRIT system is shown below.

Figure2.5 MRIT system diagram (Trudel et al. 2009)

Page 39 of 92

Due to cost effectiveness they used stainless steel to construct the extraction grids. For the same reason they used Argon instead of Xenon. The following table gives a brief description of the material used in the experiment and the materials intended for on flight use.

Table2.3 Materials used for Laboratory vs. material to be used for on flight model

Laboratory Model •

On Flight Model

Extraction grids constructed

•

Molybdenum will be used

•

Alumina ceramic will replace

from stainless steel •

Assembly components were made of Teflon

•

Teflon

•

Propellant was Argon

Propellant will be Xenon

The vacuum chamber used for the experiment was approximately 0.6 meter in diameter and 1.0 m in depth. They used a BOC Edwards IPUP Scroll Pump in addition to a CTI-Cryogenic Cry-Torr10 Series Cryopump to reach a pressure as low as 10−6 𝑇𝑜𝑟𝑟 . Clearly, they did not use SI unit in the paper which is a drawback.

To ensure the pressure inside the vacuum chamber was accurate, they used a MKS series 999 Multi Sensor Pressure Transducer and an Inficon CC3 Cold Cathode Vacuum Gauge. They used a Horiba Stec Mass Flow Controller (MFC) along with an MKS147B control box to control the flow rate of propellant. They used two Bertan 205B series high voltage sources to provide the required voltage. For the experiment they produced RF field of 1.5 MHz with a HP 33120A Arbitrary Waveform Generator. A RF Power Labs Model ML50 RF amplifier was used to amplify the signal. Page 40 of 92

2.3.3.3

Results

The functional propellant flow rates, RF power level, and exit grid potential values are necessary to find out the primary characteristics. The screen grid and acceleration grid needed a potential of +1000V and +200V respectively for a steady state operation. The propellant flow rate was 0.035 sccm (Standard Cubic Centimetres per Minute) and RF power level was 15W. The average current density was in the order of 2.0 𝑚𝐴/𝑐𝑚2 and thrust was

22.5 μN with an 𝐼𝑆𝑃 of 2096 s. They produced a two dimensional beam current density profile as follows (Figure 2.6).

Figure 2.6 Two-Dimensional beam current density profile (Trudel et al. 2009) Figure 2.7 represents the results from the optics throttling tests.

Figure 2.7 MRIT thrust vs. time (Trudel et al. 2009) Page 41 of 92

Figure 2.8 represented the MRIT efficiency in terms of propellant mass efficiency and thrust.

Figure2.8 MRIT mass efficiency vs. thrust at multiple propellant flow rates (Trudel et al., 2009)

It is clear from the graph that the thruster achieved a maximum thrust of 59.0 μN with an 𝐼𝑆𝑃 of 5480s and a mass efficiency of 60%-80% depending on the propellant flow rate.

While concluding, Trudel et al. (2009) gave an overall idea of what they have done in the experiment. They concluded that MRIT thruster could operate at a low RF input power of 13W and mass flow rates of 0.02-0.1 sccm. They operated the steady-state operation of the thruster with a RF input of 15 W and flow rate of 0.035 sccm. The potential difference between the screen and the accelerator was kept at 1200 V. The thruster produced a thrust of 1.45 μN-59.0 μN with a thrust resolution of 4 μN-10 μN and the 𝐼𝑆𝑃 for the maximum thrust

was 5480 s. The paper clearly identified their future work which involves improving the mass efficiency of the MRIT. The authors also considered the conversion of MRIT to fight materials and production of MRIT specific on flight electronics as an important next step. Page 42 of 92

2.3.3.4

Development and Test of the RIT- μX Mini Ion Engine System

The objective of the publication of Leiter et al. (2009) was to focus the results of performance tests of Radio-Frequency Ion Thruster (RIT) thruster under the Gaia Science Team Program (GSTP) of ESA . The authors also mentioned all the parameters (𝐼𝑆𝑃 , Thrust,

Power Consumption etc.) that were investigated during the experiment. The authors state that miniaturized Ion Engines are good for low thrust application for their high propellant efficiency and very low noise level. They can be used in micro satellites as primary propulsion devices. Although they mentioned that the paper was focusing on the functional test results of the project, they did not mention any of the experiments that were performed. Leiter et al. gives a useful literature review before it goes to the experiment section. They state that the altitude of a satellite needs to be reduced in order to get a better resolution. As a consequence, the satellite experiences more atmospheric interference. High thrust controllability and resolution is an effective solution to reduce this atmospheric drag experienced by the satellite. Moreover, sufficient total impulse is essential for this process. The paper gives a brief description about the basic of RIT- μX thruster. RIT involves unique electrodeless ionization of propellant with the use of electromagnetic waves. The implementation (ionization) is very simple which needs only two components: an Ionization Chamber (made of isolating material) and a RF coil surrounding the chamber.

Page 43 of 92

Figure 2.9 RIT- μX elegant breadboard (Leiter et al., 2009) RIT-μX engines require propellant and electricity. Flow Control Unit (FCU) is used to control the propellant flow, whereas the Power Processing Unit (PPU) controls the electric power supply (Figure 2.9). The neutralizer compensates ion current from the thruster by emitting electrons. RF Generator produces AC current and FCU regulates the Xenon flow to thruster. 2.3.3.5

Tests and Results

Leiter et al. performed various tests to measure the performance of RIT engine. The main problem with their report is that no description of experiments is given. As a result the reader might find it difficult to comprehend the results as the test method is unclear. The paper supported their discussion of results with some clear graphical representation. This could be even better if they have presented the relevant graph in the relevant section rather than putting all the graphs together.

Page 44 of 92

Figure2.10 RIT-μX Performance, Specific Impulse as function of total power and thrust level (Leiter et al., 2009) Graphs clearly shows the relation between total power and 𝐼𝑆𝑃 of the thruster in

different thrust levels. Using different colours for different thrust profile makes it easier for the readers to understand. The main results are presented below which indicates the second paper successfully produced the results for the readers.

Page 45 of 92

Figure 2.11 RIT- μX 50μN Thrust Stepping (Leiter et al., 2009)

Figure 2.12 Thrust Stepping -wide range (Leiter et al., 2009)

Page 46 of 92

Figure 2.12 Thrust Stepping-small range (Leiter et al., 2009) The second paper summarised their findings on the premise that a successful elegant breadboard model for RIT-μX was completed.

2.4

Performance and Operating Characteristics of Electric Propulsion Table 2 presents the performance and operating characteristics of Electric propulsion

systems. Except electro-thermal Resistojet and Hydrazine thrusters, all other thrusters can provide a high ISP of 1500s. Hydrazine thruster has other issues to be incorporated to MEMS technology. Currently MEMS technology uses Silicon at a large scale as working material. Pure Silicon is dissolved by hydrazine. As a consequence the MEMS technology cannot be used in case of hydrazine systems. These electrical systems can be scaled down to be used in microsatellites.

Page 47 of 92

Table2.4 Performance and Operating Characteristics of Electric propulsion systems (Wertz et al., 1999) Propulsion

ISP s

Type

Electro-

450-

thermal Arcjet

1500

Electro-

150-

thermal

700

Thrust

Propella

Energy

Power/

Specific

Thrust/

Total

Range

nt

Conversion

Thrust

Power

weight

Impulse

(mN)

efficien

Efficiency

(kW/N)

(kW/kg)

cy (%)

(%)

27-37

91-95

6-15

0.25-.5

100-2000

(N-S)

0.003-

12,000

0.005 180-500

35

60

1.3-2

0.4-0.8

0.02-

300,000

0.05

Resistojet Electro-static

1100-

Colloid

1500

Augmented

294-

Hydrazine

304

Radio

3000-

Frequency

3150

0.001-0.5

75

180-300

9

0.0002

1.5-3

0.5

>1000

0.0180.036

15

71-80

64

39

0.07

0.00017

0.1-0.45

0.0006-

Ion Field

4000-

0.001-

Emission Ion

11000

1000

Hall Thruster

950-

11-512

33-60

42-67

91-93

16-19

1950 Pulsed Plasma

830-

2300

0.003 0.3-0.75

7-9

80

1200

Page 48 of 92

83-100

0.003-

0.00000

15,000-

0.005

4

20,000

2.5

Feasible Propulsion for microsatellites The nature of the mission will restrict the choice of the onboard propulsion system.

As we are concerned about missions such as LISA , that require precise attitude control for formation flying, we have to consider the required thrust level and the thrust duration. In the previous section it was described that Electro-thermal Resistojet and Hydrazine thrusters does not seem to be very promising because of their low specific impulse (other issues regarding fabrication were also briefly discussed). As the described missions in this report require thrust in the range of micro newton to milli newton, Colloid thrusters, RF Ion, Pulsed Plasma Thruster (PPT) and Field Emission Electric Propulsion (FEEP) can be selected due to their low level of thrust. However, PPT can be excluded because of its high power requirement. As the microsatellites are limited in area as well as power, a power hungry system is to be avoided. Colloid thruster technology is very promising to provide simple and high perforation solution in space. Colloid propulsion systems are already miniaturized due to their operating characteristics. The recent development in microfabrication has enabled effective fabrication and prototyping of colloid thruster. Moreover, they can produce thrust by accelerating both ions and charged droplets. By modifying the ion or charged droplet fraction, colloidal thrusters can be operated with different specific impulse and efficiency. Also the power required to operate is comparatively low (~0.05 𝑊/𝜇𝑁, Smith et al., 2009). This paper will

further investigate the potential application of colloid thrusters as primary propulsion unit of microsatellites. However, the author does not rule out the usability of FEEP or RF Ion thrusters in microsatellites. It is his growing interest and the above mentioned reasons to carry out further research on colloid thruster technology.

Page 49 of 92

Chapter 3 3.1

Physics of Colloid Propulsion Colloid thrusters are a form of electric propulsion in which charged liquid droplets or

ions with high charge per unit mass (200 − 400 C/kg for glycerol with a conductivity of 0.02 𝑆𝑖/𝑚) are accelerated through an electrostatic potential. This phenomenon is also

known as elcectrospraying. Colloid thrusters do not rely on ionization in the gas phase (plasma) which is a high energy process. Propellant is stored in a reservoir. Sometimes the propellant is doped with salt to increase its ability to conduct an electric current (Pranajaya, 1999). Back in early 1960s and 1970s, most colloid systems used glycerol as the propellant. Due to very low conductivity of glycerol (A 19.3% w/v NaI in glycerol has 0.021 Si/m electrical conductivity), very high electrostatic potential (>10 kV) was needed to produce colloid beams with reasonable 𝐼𝑆𝑃

(Gamero-Castano, 2001). The liquid propellant in the reservoir must contain free charges (negative and positive). Generally, solution of salts or molten salts is used as propellant. Liquid water is problematic to use in vacuum although it is a good solvent. Some salts, also known as ionic liquids, remain in liquid state at the room temperature. One of the mostly used salts having this property is 𝐸𝑀𝐼 − 𝐵𝐹4 (1 – ethyl – 3 – methylimidazolium tetrafluoroborate).Molten salts which is also known as ionic liquids can be used to extract ions electrostatically. 𝑬𝑴𝑰 − 𝑩𝑭𝟒

𝐸𝑀𝐼 − 𝐵𝐹4 (1 – ethyl – 3 – methylimidazolium tetrafluoroborate) is an attractive

option for colloid thruster because of their high conductivity. It is possible to operate the thruster in pure Ion regime using this propellant. If positive ions are continuously extracted from EMI-BF4, then the negative ions react over the inner capillary walls blocking the liquid Page 50 of 92

flow. Density of 𝐸𝑀𝐼 − 𝐵𝐹4 is 1130 𝑘𝑔/𝑚3 . It has a conductivity of 1.3 𝑠𝑖/𝑚. Surface tension is 0.052 𝑁/𝑚.

𝑬𝑴𝑰 − 𝑰𝒎

𝐸𝑀𝐼 − 𝐼𝑚 (1-ethyl-3-methyllimidazolium bis (triflouromethylsulfonyl) amide )has a

lower surface tension than 𝐸𝑀𝐼 − 𝐵𝐹4 which makes it possible to keep the starting voltage

relatively lower. Moreover, as it does not contain any fluorine, there is no possibility of emitter damage. However, it is relatively difficult to reach pure ionic regime using EMI-Im compared to EMI-BF4 (Lozano, 2006). EMI-IM has a density of 1.53 𝑔𝑚/𝑐𝑚3 and its molecular weight is 391.31 𝑎𝑚𝑢.

Extractor

Accelerator

Capillary

Figure 3.1: Schematic of a colloid thruster As shown in figure 3.1, the liquid passes through a capillary tube. A high electric potential difference is maintained with respect to the extractor electrode, which results a strong electric field at the capillary tip (Figure 3.1). The fluid surface becomes unstable and deforms into a conical meniscus when the potential difference reaches a certain threshold limit which is given by 1.7 𝑒𝑉 − �

𝑒𝑣 𝐸

4𝜋 ∈0

(Gamero-Castano, 2000). A thin jet is created at the

Page 51 of 92

tip of the meniscus which later ejects small charged droplets. The same electrostatic field is used to accelerate the droplets to produce thrust (Khayms, 2000). 3.1.1

Surface Charge

Let us assume that a strong normal electric field 𝐸𝑛 is applied to a liquid surface. If

there are free ions in the liquid, the opposite polarity will be attracted to the surface. The

charge per unit area , 𝜌𝑠 can be determined by integrating the control volume indicated in the

figure using Gauss’ law ∇ . 𝐸�⃗ = 𝜌𝑐ℎ ⁄𝜀0 .

Electric Field, 𝐸𝑛

Gas

Liquid

Figure 3.2 : Charge Concentration change in Electric conductor Therefore, for any electrical conductive liquid charge per unit area can be written as:

3.1.2

Taylor Cone:

𝜌𝑠 = 𝜀0 𝐸𝑛

3.1

From previous experimental observations it is known that the surface of a conductive liquid deforms when it experiences high electric potential. The electrostatic pull is increased in a cascading effect due to the increase of charge concentration in the surface area. If the applied electrostatic potential reaches a certain limit, the liquid surface forms the shape of a Page 52 of 92

cone. A very thin, fast moving jet is emitted from the apex of the cone. Taylor explained and also experimentally verified this behaviour of the liquid. The surface traction generated by the strong electric field must be balanced by the surface tension on the conical surface. Surface tension of the liquid can be expressed per unit of area as: 𝑓𝑠𝑡 = 𝛾(

1

𝑅𝑐1

−

1

𝑅𝑐2

)

3.2

where 𝑅𝑐1 𝑎𝑛𝑑 𝑅𝑐2 represent the principal surface radii of the curvature. The surface traction

experienced by the liquid is 𝜀0 𝐸𝑛2 ⁄2 (Martinez-Sanchez, 2001).

Figure 3.3 Cone jet structre for Ethylene – Glycol (Fenandez, 1994)

Page 53 of 92

𝑅𝑐 𝑅

𝐸𝑛

𝛼

𝑟

Figure3.4 Taylor cone geometry with an inner angle 𝛼. Meusnier’s theorem (1776) states that “all curve lying on a surface S and having at a given point p∈ 𝑆 the same tangent line have at this point the same normal curvature”. Therefore,

1

𝑅𝑐

1

= � � cos 𝛼 = 𝑅

cos 𝛼

𝑟 sin 𝛼

=

1 𝑟

cot 𝛼

3.3

So the surface traction can be expressed as: 𝛾

𝜀0 𝐸𝑛2 ⁄2 = cot 𝛼 𝑟

𝐸𝑛 = �

2 𝛾 cot 𝛼 𝜀0 𝑟

3.4

Let us consider the spherical coordinate system in figure 9 to determine the external electric field with which the cone is in equipotential.

Page 54 of 92

𝑧

𝑃(𝑥, 𝑦, 𝑧) 𝜌

𝜑

𝑧

r 𝜃

𝑥

𝑦

Figure3.5 Spherical Coordinate system The electric field is inversely proportional to 𝑟 and it exhibits singularity as 𝑟 → 0. If

the region outside the cone is considered to be charge-free, the field is described by Laplace’s equation. ∇2 ∅ = 0. For conical section the Laplace’s equation is: ∇2 ∅ =

1

𝜕

𝑟 2 𝜕𝑟

� 𝑟2

𝜕∅ 𝜕𝑟

�+

1

𝜕

𝑟 2 sin 𝜃 𝜕𝜃

�sin 𝜃

𝜕∅

𝜕𝜃

�

3.5

As we need the solution outside the liquid conical section, 𝜃 is measured from inside the cone. The solution for equation 22 is in terms of Legendre polynomials: ∅ = 𝐴 𝑃𝑣 (cos 𝜃) 𝑟 𝑣

∅ = 𝐴 𝑄𝑣 (cos 𝜃) 𝑟 𝑣

3.6 3.7

𝑃𝑣 has singularity at 𝜃 = 180° and 𝑄𝑣 has singularity at 𝜃 = 0°. The solution in terms of 𝑄𝑣

is accepted as we need the solution outside the conical section and the singularity in this case is inside the cone. So the normal field can be written as follows: Page 55 of 92

𝐸𝑛 = −

1 𝜕∅ 𝑟 𝜕𝜃

=𝐴

𝑑 𝑄𝑣

𝑑 (cos 𝜃)

sin 𝜃

1

3.8

𝑟 1−𝑣

In order to have the normal E-field in equilibrium with the surface tension, the value 1

of the exponent has to be 𝑣 = . So the solution is: 2

∅ = 𝐴 𝑟1/2 𝑄1/2 (cos 𝜃)

3.9

The function 𝑄1/2 has a single zero at 𝜃 = 49.29°. This angle is independent of the property

of the liquid, geometry of the liquid or the applied potential. Taylor verified this value experimentally but it does not hold when strong charge effects are acting on the liquid cone.

𝜃

𝜃

Figure 3.6 Plot of Legendre polynomials (Lozano, 2003) Also the electrode geometry affects the value. The actual electrode set up may not resemble that of the Taylor’s model. Moreover, the charged jet modifies the potential distribution of the liquid cone which leads to a deviation from the value of Taylor’s angle.

Page 56 of 92

However, Taylor model succeeds to explain the behaviour of conical section of the liquid and is valid in a specific region of the jet to the cone’s base. 3.1.3

Starting Voltage:

A certain electric field has to be induced on the liquid surface in order to result the Taylor cone. The electrostatic field can be expressed as: 2 𝜀0 𝐸𝑡𝑖𝑝

2

=

2𝛾 𝑅𝑐

3.10

where, 𝜀0 = permittivity in vacuum ,

𝐸𝑡𝑖𝑝 = Electric potential at the tip of the conical surface

𝑅𝑐 = principal surface radius of the curvature and 𝛾 = surface tension of the liquid.

For a meniscus diameter 𝑑𝑐 and extractor to meniscus distance D, Eyring (1927) derived the following expression for the electric field around solid metal tips: 𝑑 𝛾

4𝐷

𝑉𝑠𝑡𝑎𝑟𝑡 = � 𝑐 ln( ) 2𝜀 𝑑 0

𝑐

3.11

Equation 28 is just an approximation as it does not consider the fluid dynamic nature. Moreover, the tip is assumed to be an equi-potential hyperboloid for the approximation to hold.

Page 57 of 92

3.2

Related work in the area of Electro-spray Zeleny (1917) pioneered the field of elecrospray through his experimental

investigations. He showed that a stable conical meniscus could be obtained from a liquid surface flowing through a capillary tube if it experienced an electrostatic potential. Taylor (1964) was the first one to describe the formation of the cone. Taylor showed that the cone was the result of the electrostatic stress acting on the liquid surface which reacted with the surface tension of the liquid. He derived the required semi-vertical angle to be 49.3° based

on general assumptions that the conical surface was equipotential and the cone was in steady state equilibrium. This was later confirmed experimentally. Although Taylor was successful and widely accepted in explaining the geometric properties of the liquid cone, his analytical model was unable to explain the formation of the thin jet issuing at the apex of the cone. Moreover, it was shown in various experiments that

the fluid meniscus deviated from Taylor angle when operating conditions varied.

For

example, droplets are mobility limited in air. Fernandez (1992) showed that the deviation from Taylor angle occurred due to space charge created by the charged droplets in the jet which disturbed the electrostatic field of the conical surface. Mestel (1994) developed a model to explain the mechanism for the formation and physics of the jet for liquid flows at high Reynold’s number. Later work was carried out by Fernandez (1994). He developed another model to describe the instability near the jet tip. He proposed that this behaviour was due to the convection associated with the liquid flow which transported the net surface charge towards the cone tip. He developed a sink flow model to determine a rough scaling of tip current and flow rate. Later he carried out a joint experimental work with Juan to determine the charge distribution, size of droplets and the charge to mass ratio. A Vienna type Differential

Page 58 of 92

Mobility Analyzer (DMA) was used to sample the spray drops and to measure the associated current or to pass them to an aerodynamic size spectrometer. The charge to volume ratio of the droplets showed that charge is frozen in the liquid surface during breakup. Earlier studies involving glycerol solutions and the recent works involving highly conductive ionic solutions ( 𝑓𝑜𝑟 𝑒𝑥𝑎𝑚𝑝𝑙𝑒 , 𝐸𝑀𝐼𝐵𝐹4 ) indicate that ions with much higher

charge to mass ratio can be produced from the liquid cone tip along with charged droplets. Benignos (2005) worked to create a numerical tool to analyze the current, droplet size, velocity and electrostatic potential for a fixed geometry and a specific fluid at a fixed flow rate. Lozano (2007) also carried on study on the pure ion emission regime. From the experiment he found that for a 500 V applied voltage, the electric field for ion emission site is 0.15 𝑉/𝑛𝑚. He also showed that it low interception of emitted beam on extractor electrode was possible to achieve.

3.3 Developments in colloid propulsion Krohn (1962) pioneered the use of electrospray for space propulsion. He experimented with liquid metals and highly viscous organic liquids (e.g. glycerol) for the first time. He observed the mixed regime of ion and charged droplets emitted from the liquid propellant. The main concentration was to develop a suitable configuration for electrodes and liquids for the use of space propulsion. As on board propulsion demanded high thrust density, the operating voltage for the colloid thruster proved to be very high (in the range of 10-15 KV). Kaufman Ion Engine, which was relatively simple in operating principle, could achieve similar performance. So the initial idea of colloid propulsion was interrupted. The idea of colloid thruster was regenerated in the 1990s. The increasing number of missions involving microsatellites (10-100 kg) revolutionalized the field of colloid propulsion. Fenn (1989) used the “soft” ionization technique to apply mass spectrometric analysis on Page 59 of 92

large and fragile polar molecules.

As stated earlier, colloid thrusters deemed to be a

promising primary propulsion system due to its operational range. Shtyrlin (1995) published a paper describing the work on colloid thrusters in the Moscow Aviation Institute (MAI) for 35 years. MAI developed a colloid thruster which could operate at a thrust range of 0.5 − 1𝑚𝑁 and required a 30 W DC power supply. The

electrostatic potential had to be kept at a very high level of 15-25 kV. The data from Table 5

shows that the thruster was suitable for the missions involving satellites with masses of 2525- kg. On board power requirement for the mission profile was 1 W/kg. The expected ∆𝑉 budgeting of the targeted mission was 40-400 m/s.

Table 3.1 Specification of MAI colloid thruster developed at MAI (Shtyrlin, 1996) Attributes

Values

Thrust range

0.5-1mN

Power Requirement

30 W DC

Voltage Supply

15-25 KV

Suitable for S/C ranged

25-250 kg

On board power requirement

1 W/kg

∆𝑉 budgeting

40-400 m/s

Perel (1998) developed a thruster for the purpose of micro electric propulsion which was capable of operating at a high specific impulse in the range of 1000s. The thruster used glycerol as propellant and used was valveless in operation. The neutralization of charged ions was achieved by operating two emitters which produced negative and positive ions. The “micro-volcano emitter” was an integrated MEMS design for principal on-board propulsion system for microsatellites. Page 60 of 92

Figure 3.7 Prototype of the 100-nozzle thruster (Pranajaya, 1999) Pranajaya (1999) developed a one-nozzle and 100-nozzle emitter prototype (Figure 11) under a university nanosatellite project called Emerald. The prototype consisted of two 3 × 3 𝑐𝑚 brass plates- one for the source and one for the extractor. Undoped glycerol and

isopropyl alcohol were used as propellants during the experiment. A stainless steel capillary with an inner diameter of 0.002 inch and an outer diameter of 0.006 inch was used as the emitter. Experimental result showed emitter current in pico-ampere level. However, the report does not mention anything about the conductivity of the propellant. Khayms (2000) worked under Martinez-Sanchez during his MSc and PhD. Khayams worked with miniature hall thruster during his MSc but in his PhD he developed scaling laws for the most promising thruster systems retaining the basic non dimensional qualities that determine the thruster performance. He presented the formulations of the physics of colloid propulsion, which included non-zero flow rate effect and non-zero electrical conductivity. The effects of different electrode geometry, pressure, fluid inertia and electrostatic inertia were also discussed. The developed model could explain the base of the conical tip as well as the emission jet in a geometric scale ratio of 10,000:1. The model can be used to describe

Page 61 of 92

various profiles of colloidal thruster with different geometric definition and for different operating conditions. Reichbach (2001) analyzed promising propulsion systems for several missions namely Space Technology 3 (ST-3), LISA, Terrestrial Planet Finder (TPF), Micro Arcsecond X-ray Imaging Mission (MAXIM) and Sub-millimeter Probe of the Evolution of Cosmic Structure (SPECS). All these missions required precise position and attitude control. Performance, cost and technical feasibility were the three criteria he used to select a propulsion system for the formerly mentioned formation flying missions. He identified colloid thruster and FEEP thruster suitable for the precision formation flying spacecrafts.

Figure 3.8 Testing arrangement of prototype (Left); Schematic of Prototype (Right) [Paine, 2005] Paine (2001) started working with the Micro-fabricated colloid thruster arrays. However the array designed by him could not be tested because of electrical breakdown. However, he carried on his research and in 2005 he presented the concept of a Nanoelectrospray colloid thruster which consisted of switchable emitter clusters. It was designed

Page 62 of 92

to provide “precisely throttled thrust over wide ranges”. Nozzles which were capable of electrospraying were grouped in clusters. Each cluster of emitters (Figure 12) was independent and could operate without the interference of other clusters. The idea was novel and very important in the sense that this thruster could provide thrust over a wide range and could easily be tailored as per the mission requirements. Hruby et al. (2001) built a colloid thruster system for NASA Jet Propulsion Laboratory (JPL) which could continuously produce thrust in the range of micro-newtons. The thruster was integrated with a cathode neutralizer and a power processor to control the liquid feed system. The envelope containing the thruster system was 9 × 5 × 5 𝑖𝑛𝑐ℎ𝑒𝑠 in

dimension. The mass of the system was 2.5 kg including enough propellant to produce thrust for 3000 hours. The prototype had 57 needles or emitters with double grid facilities (extractor grid and accelerator grid) to ensure 0.1 𝜇𝑁 thrust stability. The required power for steady state operation was < 6 𝑊. The Zeolite heater, which was used for the feed system, needed a

maximum of 4W power. The thruster was designed to be compatible with NASA missions such as LISA, Laser Interplanetary Ranging Experiment (LIRE) and Earth Science Experimental Mission 5 (EX-5). Performance of the thruster was measured with a formamide propellant which is measured to have a conductivity of 0.5 𝑆𝑖/𝑚. The thrust achieved from

the thruster was in the range of 20 − 190 𝜇𝑁 . However, due to unexpected lower propellant conductivity, the maximum 𝐼𝑆𝑃 was 400s.

Xiong et al. (2002) developed an integrated colloidal micro thruster using PCB (print Circuit Board) technology. They used Formamide with 30% sodium Iodide (NAI). The thruster array consisted of 81 copper emitters each of which had an internal diameter of 0.3mm and an external diameter of 0.5 mm. Emitters were placed 1.5 mm apart from each other. The design was later perfected using Microelectromechanical Systems (MEMS) based

Page 63 of 92

technique. This time the array was built with a total of 192 emitters. Xiong et al. conducted a test for an array of four emitters on vacuum. Their report indicates that the thruster could be operated with an external voltage of 2000V to produce a thrust of 6.8 𝜇𝑁. Xiong et al.

published another paper in 2004. In this paper they described their work in developing another colloid thruster by silicon processing. They reduced the thruster dimension and also the external voltage level was considerably low in a range of 1-3 kV. The starting voltage was found to be 1400V. The thrust and its variance were compared against the applied voltage which is given in Table 6. It is clear from the table that the mean thrust of the colloid thruster was 1.36 𝜇𝑁 with a standard variation of 0.12 𝜇𝑁. Table 3.2 Thrust and Thrust variance vs. applied voltage (Xiong et al, 2004) Applied Voltage 1400

1600

1800

2000

2200

2400

2600

2800

(V) Thrust (𝜇𝑁)

1.36

1.4

1.49

1.55

2.0

3.25

3.95

4.85

Standard

0.12

0.126

0.13

0.14

0.19

0.34

0.5

0.5

Variance (𝜇𝑁)

Kirtley (2002) suggested a conceptual design of a colloid micro thruster with an electrodynamic linear accelerator. The same hydraulics was used but the electrodynamic linear accelerator was used instead of the electrode system. This approach of accelerating the exit jets gives a higher 𝐼𝑆𝑃 and simple physics for acceleration as well as high acceleration efficiency.

Velasquez (2004) worked on fabricating a colloid thruster with dense emitters. He demonstrated that using MEMS it was possible to construct a highly dense colloid thruster arrays. His work involved two different engine concepts: one linear thruster array which was fed internally with doped Formamide and the other engine was fed externally by 𝐸𝑀𝐼 − 𝐵𝐹4 . Page 64 of 92

The first engine was intended to work in the charged droplet regime. This thruster was suggested for missions requiring a thrust in the range of 100 - 450𝜇𝑁. The specific impulse ranged between 200 - 350 s. The second planer array of thrusters was suitable for missions requiring a larger 𝐼𝑆𝑃 of 3800 – 8500 s and a thrust in the range of 1-3 𝜇𝑁 . Gassend (2007) fabricated a fully integrated colloid thruster which had a weight of 5 gms. The thruster had a specific impulse of 3000 s when it was operated in the ion regime with 1 kV extraction voltage. The experiment also showed that with an electrostatic potential of 1500 V it was possible to produce a thrust of approximately 13𝜇𝑁. The power consumed

by the colloid thruster was 275 mW. The thruster consisted of 502 emitters spreading over an area of 113 𝑚𝑚2 which gave approximately 26 nN thrust per emitter. The efficiency of the

thruster was assumed to be 85%.

Table 3.3 accumulates all the characteristic values of different colloidal thrusters produced: Table 3.3: Colloid thrusters developed to date Developer

Technology

Propellant

Emitters

Thrust

ISP

Voltage

Gassend,

?

Formamide

502

13𝜇𝑁.

3000s

1500V

4 emitters

6.8 𝜇𝑁

?

1970 V

?

100 - 450𝜇𝑁

200-350s ?

?

1-3 𝜇𝑁

3800-

2007

emitters

Paine,

Micro

2005

fabrication

Velasquez, MEMS

Triethylene glycol

(0.02 𝑆𝑖/𝑚) Formamide

2004 Velasquez, MEMS

𝐸𝑀𝐼 − 𝐵𝐹4

Page 65 of 92

?

2004

8500s

Xiong,

MEMS

2002

with

Hruby,

?

192

30% copper

NAI

emitters

Formamide

57

(0.5 𝑆𝑖/𝑚)

2001

Pranajaya, 1999

Formamide

?

6.8 𝜇𝑁

?

1-3 kV

?

20-190

400s

𝜇𝑁

stainless steel

Glycerol and

Stainless

Iso-propyl

Steel

?

?

?

alcohol 3.3.1

Colloid Propulsion Research at Queen Mary

The interest in colloid thruster has recently been renewed with the increasing interest to develop smaller satellites and with the advancement of micro-fabrication technique. Queen Mary, University of London and Rutherford Appleton Laboratory took a joint programme to develop an integrated micro fabricated colloid thruster (Stark et.al. 2003). Krpoun et al. (2007) designed, fabricated and tested an electrospray micro-thruster. The design was based on the limitations of the state of the art microfabrication process and the mission specification for missions like LISA or DARWIN. The final prototype of the thruster was based on a surface area of 1𝑐𝑚2 . They developed two designs. One of which

consisted of capillaries emitting all at a time and the other design allowed each and every capillary to be independent of the functionality of the other emitters. As stated earlier, 1 × 1𝑐𝑚2 silicon wafer was used for each thruster layout. Figure 3.9 illustrates that each capillary was 70 𝜇𝑚 high and the distance between two adjacent capillaries was 250 𝜇𝑚. By

Page 66 of 92

varying the emitter and extractor diameter and the spacing of the emitter, different designs were constructed.

Figure 3.9 Capillary Geometry (Krpoun et al., 2007)

Figure3.10 Schematic cross section of the colloidal thruster (Krpoun et al., 2007) The chosen materials were doped with silicon to increase conductivity and a borosilicate glass was used as insulator (Figure 3.10). After manufacturing the capillary emitters and the extractor electrodes, the thruster is assembled chip- wise using glue. 𝐸𝑀𝐼 − 𝐵𝐹4 was used (ambient pressure) to fill the thruster assemblies. The emitters were operated at positive or negative voltages using a high voltage source (±5𝑘𝑉). A Faraday cup Page 67 of 92

was used to measure the tip current with a picometer. The electrospraying of a single emitter with an inner and outer diameter of respectively 20 𝜇𝑚 𝑎𝑛𝑑 14 𝜇𝑚 was measured during the experiment. Votage vs. Current curve was obtained for the configuration where the extractor electrodes and the emitter was 25 𝜇𝑚 (figure 3.11). Same graph was obtained when the spacing was 40 𝜇𝑚 (figure 3.12).

Figure 3.11 Current vs. voltage curve- 25 𝜇𝑚 spacing (Smith et al, 2007)

Figure 3.12 Current vs. Voltage –extractor and emitter distance 25 𝜇𝑚 (Smith et al, 2007) Page 68 of 92

Another test was carried out to measure the performance of the thruster designed with emitters capable to operate independent of each other. At 570 V a current of 150n A was measured. However, when the voltage was increased, short circuit occurred due to a leakage. Based on the experimental results a simple current-voltage model was proposed. Another experiment was carried out by Smith et al. (2009) to determine the beam properties. Both micro fabricated emitters and off the shelf ESMS (Electrospray Mass Spectrometry ) were tested. The thruster was studied in two different modes- high 𝐼𝑆𝑃 with

low thrust density and low 𝐼𝑆𝑃 with high thrust density. 𝐸𝑀𝐼 − 𝐵𝐹4 was used as propellant.

The first measurement was carried out with a single ESMS emitter (30 𝜇𝑚 ) with one

extraction grid. The second test was to study the performance of an assembly of emitters, extraction grid and an acceleration grid.

During the characterization test of the beam, 30𝜇𝑚 silica tips (tip dimension accuracy of ±2 𝜇𝑚) were used as emitters. These emitters were mounted on the ionic liquid reservoir. A submerged electrode of stainless steel ensured the contact of the liquid propellant and the high voltage power supply.

Figure 3.13 Hybrid Colloid Thruster (Smith et al., 2009) Page 69 of 92

Figure 3.13 shows the hybrid colloid thruster consisting of 19 emitters and a conventional fluid reservoir and acceleration grid. From the experimental results it was suggested that two modes of colloidal thruster operations could be achieved (Table 3.4) for the purpose of station keeping and low thrust maintenance to meet the future micro satellite missions. Table 3.4 Proposed thruster flight experiment (Smith et al., 2009)

3.4

Applicability of Colloid thrusters This section briefly reviews the advantages of the colloid thrusters and their

applicability to micro propulsion. As we discussed in the previous section, the liquid tip of the propellant (e.g. ionic salt) is exposed to high electrostatic potential 𝑉 with respect to an

external cathode. Consequently, the meniscus tip deforms as the perturbing potential is high enough to overcome the surface tension of the liquid. This results into a jet of small charged droplets. Droplets are electro-statically accelerated to produce thrust. If the mass of an individual droplet produced by the unstable jet is 𝑚 and the charge it carries is 𝑞 then the exhaust speed is 𝑐 can be obtained from the conservation of energy: 2𝑞𝑉

𝑐= �

𝑚

Page 70 of 92

3.12

𝑞

It is clear from equation 3.12 that the exhaust speed depends on the value of . Colloid 𝑚

thrusters operating in the droplet regime does not have a higher ISP as it is difficult to produce droplets with a high charge to mass ratio (Gassend, 2007). It requires very high voltage (~10,000 V) as well. However, recent developments in the field of colloid propulsion showed that it is possible to operate the thruster in droplet region with a low voltage by choosing a propellant with desired fluid property, by controlling the geometry of the capillary tube and by ensuring suitable operating condition (Xiong, 2006). For liquids having an electrical conductivity of ≤ 0.1 𝑆𝑖/𝑚, the charge to mass ration

obtained is in the range of 200 − 400 𝐶/𝑘𝑔. As we know that the specific impulse of thruster is given by 𝐼𝑆𝑃 =

𝑐

𝑔0

, a low charge to mass ratio results into a low exhaust speed as well as a

low 𝐼𝑆𝑃 . It has been shown that high conductive solutions at a low flow rate give droplets with 𝑞

high (Fernandez, 1994). Castano (1999) showed that droplets with 5000 C/𝑘𝑔 could be 𝑚

produced during experiments, but still needed a high voltage of 10 kV. It is possible to produce ions with larger charge to mass ratio than droplets. So the thruster has a longer 𝐼𝑆𝑃 in

the ion regime. It is possible to ensure a desired value for 𝐼𝑆𝑃 with help of a combination of droplets and ions (Khayms, 2000).

Although developments have been made to operate the colloid thruster in a lower voltage, still a threshold voltage is required to overcome the surface tension: 𝛾𝑅

𝑉𝑚𝑖𝑛 = � 𝜀

0

3.13

Equation 30 is an approximation as it assumes the liquid meniscus to have a simplified spherical profile and also it does not take into account the effect of external electrodes. The mixed regime case is not optimal. “More energy is spent accelerating particles with higher

Page 71 of 92

𝑞

𝑚

than the extra thrust derived from them” (Khayms, 2000). So even for higher values of charge to mass ratio (droplets ~ 6,000 C/kg, Ions ~ 240,000 C/kg) the efficiency of the thruster is about 48%. However, advantages of colloid propulsion outweigh its disadvantages while it comes to the applicability to micro-propulsion. Colloidal thrusters are already miniature and they use liquid propellant. So the propellant tanks used are lighter and compact than gaseous systems. The energy required to produce charged particle is very low (7-8 eV) for 𝐸𝑀𝐼 − 𝐵𝐹4 . The charged ions emerge from the extractor with a low energy speed. This can be made even lower by reducing the power,

thus decreasing the electric potential of the acceleration electrode. So colloid Thrusters can be operated at a wide range of 𝐼𝑆𝑃 (Lozano, 2006). Even the use of a neutralizer can be

overcome by placing two thrusters in a combination. As the ionic liquids can produce

negative or positively charged ions, they can neutralize each other (Maloney, 1969). It should be noted that each emitter in a colloid thruster produces thrust in the range of 𝜇𝑁 level while

consuming power in the level of 𝑚𝑊. So to provide a certain level of thrust (for example, 30 𝜇𝑁 for LISA) one has to fabricate 𝑛 number of emitters: 𝑛=

𝑅𝑒𝑞𝑢𝑖𝑟𝑒𝑑 𝑡ℎ𝑟𝑢𝑠𝑡

𝑇ℎ𝑟𝑢𝑠𝑡 𝑜𝑏𝑡𝑎𝑖𝑛𝑒𝑑 𝑓𝑟𝑜𝑚 𝑎𝑛 𝑖𝑛𝑑𝑖𝑣𝑖𝑑𝑢𝑎𝑙 𝑒𝑚𝑖𝑡𝑡𝑒𝑟

Page 72 of 92

3.14

Chapter 5 5.1

Proposed Lunar CUBESAT Mission CUBESAT is a proposed lunar cubesat mission which will follow a low thrust

trajectory to enter into a lunar orbit from an initial low earth orbit. The mission is designed using off the shelf products. Table 5.1 General Specifications of the spacecraft Attributes

Values

Dimensions Mass

30 × 10 × 10 𝑐𝑚

Orbit

Around the moon

Lifetime

1 year (initial)

Payload

USB camera

Communications

UHF Amateur Band

Key Components

3-axis Reaction wheel

2907.262 gm

Table 5.1 represents the general specifications of the spacecraft. The satellite is 30 × 10 × 10 𝑐𝑚 in dimension. Its mass is around 3 kg. As stated earlier the satellite will

orbit the moon at an altitude of 100km. Initially the satellite lifetime is 1 year which can be extended as most of the components are designed for a longer lifetime. However, an extension in lifetime is very unlikely because of the nature of its payload. It has only a USB

Camera onboard as payload. The satellite will use Ultra High Frequency (UHF) amateur band for communication. A 3-axis reaction wheel will be used for attitude control.

Page 73 of 92

5.2

Power Requirement Power requirements of the satellite have been shown in Table 5.2. All the components

are carefully picked keeping in mind that the cubesat will be limited in power generation. All the subsystems will be briefly discussed in the following sections.

Table 5.2 Power Requirements

5.3

Purpose

Unit

Typical Consumption

Max

On board computers

W

0.1

0.1

ADCS

W

1.5

4.5

Transceiver

W

0.000033

2.7126

Payload requirement

W

2

2

Thruster

W

85.06

85.06

Total

W

88.67

94.38

Communication In order to communicate with the ground station Amateur radio band will be used. A

Nanocom UHF half-duplex transceiver will be used to transmit and receive data. The specifications of the transceiver are given in Table 5.3. It operates in the frequency band 432438 MHz. It needs a single supply of 3V and can be operated at -30℃ 𝑡𝑜 +70℃. The link

budget for the mission was done in a separate spreadsheet that can be found in Appendix A.

An ISIS deployable Cubesat Antenna System for single UHF Dipole Antenna will be used. One or two radios can connect to the antenna by miniature RF connectors. Maximum radio frequency power is 2W. Insertion loss (𝑃𝑅 /𝑃𝑇 ) is 1.5 dB. It is very light weight (100gm) and operates in the range of 390-450 MHz.

Page 74 of 92

Table5.3 Transceiver Specification

Attribute

Value

Single Supply Voltage

3.3 V

TX: Current Consumption

800 mA (Max)

TX power

2.64 W (Max)

RX current Consumption

22 mA

RX power

0.0726 W (Max)

Total Tx/Rx Power

2.7126 W (Max)

Standby Power

0.000033 W

5.4

Attitude Control System The proposed cubesat orbits around the moon. An IMI-101 miniature 3 axis reaction

wheel will be used. It has a pointing accuracy of 1° . Table 5.4 Operating Characteristics of the reaction wheel Parameter

Note

Max

Unit

Lifetime

LEO

1

Year

Pointing Accuracy

Using built in magnetometer 1-3

Arcsec

and sun sensor data System Bandwidth

Typical 0.05

Hz

Power

12-28VDC@200mA

Telemetry

5

Hz

Momentum storage

Per wheel

1.1

mNms

Torque

Per wheel

0.635

mNm

Page 75 of 92

Other operating characteristics of the ACS can be found in Table 5.4.

5.5

Total Mass budget Table 5.5 represents the total mass budget of the cubesat. The total wet mass of the

spacecraft has been calculated as 3357.262 gm (~3.4kg). The dry mass is about 3 kg.

Table 5.5 Spacecraft total mass budget

Spacecraft Element

Mass (g)

Propellant

1000

Propellant Tank Mass

325.65

S/C Structure Mass

580

Transceiver

75

Antenna

100

Solar panel

150

Battery

1000

ACS

910

EMCO DC to HV DC converter

100

On board Computer

100

Payload

62

S/C MASS with Propellant

3357.262

Dry Mass

2907.262

Page 76 of 92

5.6

Propulsive Requirement The CUBESAT will be placed in a parking orbit of 800 km with an inclination of

𝑖 = 28.5°. The satellite has to use its onboard propulsion system to enter into a low thrust transfer trajectory towards moon.

Let us assume that the thrust acceleration magnitude is constant during the transfer of the spacecraft. So the thrust vector yaw angle 𝛽0 can be obtained from the following expression (Chobotov, 1996):

𝛽0 =

𝜋 ∆𝑖 2 𝑣0 𝜋 −cos� 2 ∆𝑖� 𝑣𝑓

5.1

∆𝑖 = 𝑐ℎ𝑎𝑛𝑔𝑒 𝑜𝑓 𝑖𝑛𝑐𝑙𝑖𝑛𝑎𝑡𝑖𝑜𝑛 𝑜𝑓 𝑡ℎ𝑒

where,

𝑣0 = 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑐𝑖𝑟𝑐𝑢𝑙𝑎𝑟 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝑜𝑓 𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡

𝑠

𝑐

𝑣𝑓 = 𝑐𝑖𝑟𝑐𝑢𝑙𝑎𝑟 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝑜𝑓 𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡 𝑖𝑛 𝑡ℎ𝑒 𝑓𝑖𝑛𝑎𝑙 𝑜𝑟𝑏𝑖𝑡

The initial velocity of the spacecraft is given by :

𝜇

𝑣0 = � 𝑒 𝑟

5.2

0

where 𝜇𝑒 is the gravitational parameter of earth (4 × 105 𝑘𝑚3 / 𝑠 2 ) and 𝑟0 is the

initial orbital radius around the earth (7200km). So the initial velocity can be calculated as 7.45 km/s.

The velocity of the spacecraft while orbiting the moon is:

𝜇𝑚𝑜𝑜𝑛

𝑣𝑓 = �

𝑟0

Page 77 of 92

5.3

where 𝜇𝑚𝑜𝑜𝑛 is the gravitational parameter of earth (~5 × 103 𝑘𝑚3 / 𝑠 2 ) and 𝑟𝑓 is the

radius of the final orbital around the moon (1838 km). So the final velocity can be calculated as 1.65 km/s.

The following equation can be used to calculate the ∆𝑉 required for this low thrust

trajectory transfer (Chobotov, 1996):

∆𝑉 = 𝑣0 cos 𝛽0 −

𝑣0 sin 𝛽0

5.4

𝜋

tan� 2 ∆𝑖+ 𝛽0 �

⇒ ∆𝑉 = 7.46 𝑘𝑚/𝑠

So in order to place the satellite in the final orbit through a low thrust transfer the onboard propulsion system has to provide a ∆𝑉 𝑜𝑓 7.46 𝑘𝑚/𝑠. As the satellite will orbit around the moon there will be no atmospheric drag. The only external perturbation acting on the satellite will be solar radiation. The total delta V required for cancelling the effect on the orbital parameters by solar radiation and earth oblateness is very low (~2 m/s). Table5.6 presents the spacecraft attributes where the CUBESAT mission demands a ∆𝑉 𝑜𝑓 7.46 𝑘𝑚/𝑠 during its lifetime of 1 year. Table5.6 Propulsive Requirements ΔV

m/s

Me

Kg

ISP

S

g0

m/s2

C

m/s

M0

Kg

Page 78 of 92

7460 3.1 1500 9.8 14700 5.1

Mf

Kg

1

Burn time

Sec

Mass Flow rate

kg/sec

Thrust

N

0.0017

Power

W

85

Total Impulse

N-S

8640000 1.1574E-07

14700

From literature review (Smith et al., 2009) it has been found that the power requirement for the EPFL thruster is ~~0.05 𝑊/𝜇𝑁 which gives us a total thruster power

requirement of 85.1 W during the transfer period. Table 5.2 indicates that the spacecraft demands a maximum power of about 9 W whereas the typical requirement 4 W power. If a safety margin of 20% is added then the maximum power consumption will be 10.8 W. To generate 10.8W, 288 𝑐𝑚2 active solar area is required. The CUBESAT has a surface area of

300𝑐𝑚2 which is sufficient in this case. The power required by the thrusters is 85 W. Li-ion batteries are a good option to provide this power as they have high specific energy (150 Whr/kg), light weight and have very little self-discharge (Clyde space Battery Data Sheet, 2010). Eight lithium polymer batteries has a capacity of 80 Watt hours which itself is sufficient to power the thruster. As we have to supply high voltage for the colloid thruster, an EMCO Q series Ultra-Miniature DC to HV DC converter has to be used. EMCO Q101 takes 0-5V as input and outputs maximum of 10000V which will be adequate for electrospraying process.

Table 5.7 represents fuel tank size of the colloid micro thruster. From the proposed thruster flight experiment in Table3.4 (Smith et al., 2009) it can be deducted that a thrust level of 442𝜇𝑁 (specific impulse~ 1000s) can be produced with an array of 194 emitters. So

Page 79 of 92

to produce the thrust for the proposed lunar CUBESAT mission will need an array of about 776 emitters as the total thrust scales linearly with emitter number.

Table 5.7 Fuel Tank Size Estimation

Attribute

Value

Mass of fuel

1 kg

Density of fuel (𝐸𝑀𝐼 − 𝐵𝐹4 )

1130 𝑘𝑔/𝑚3

Volume of tank (5% margin)

0.000885 𝑚3

Inner Radius of tank

92.92 𝑐𝑚3 1.5 cm

Outer Radius

2 cm

Cylinder Height

13.15 cm

Thickness

0.5 cm

Volume of Metal Density of Metal (Titanium)

72.27 𝑐𝑚3

Fuel Tank Structure Mass

325.65 gm

Volume of fuel

4.506 gm/𝑐𝑚3

Page 80 of 92

Chapter 6 Conclusions and future work In this thesis, propulsion system requirements for the near future missions requiring formation flying were studied. Colloid thrusters were identified to be the most suitable primary propulsion system of the microsatellites that require very precise attitude control. The literature related to the development in the field of colloid thrusters was reviewed. The state of the art microfabricated thrusters were studied. It has been found that the colloid thruster operation is still limited due to the geometric structure of the emitters and the nature of the liquid propellant. Propellant with high conductivity needs to be studied further to obtain operational electrospray propulsion system. Moreover due to the low level of thrust generated, the emitter density needs to be increased if thrust in the range of milli newton is to be achieved. Otherwise several thrusters units would be required which would lead to complication of the spacecraft structure. Also the operational lifetime of the thruster needs to be tested. Detailed research has to be carried out before any concrete recommendations can be made regarding the performance of colloid thrusters. Also a CUBESAT mission was proposed which will follow a low thrust trajectory to reach a lunar orbit. The goal was to demonstrate the feasibility of the mission using a colloid thruster. The required ∆ 𝑉 for the total mission lifetime was calculated. It has been shown that it is possible to provide the necessary ∆ 𝑉 with colloid thruster technology. High voltage supply is required for electrospray process. This problem was overcome with an EMCO DC to HV DC converter which can produce a maximum 10,000 V with a minimal input of 5V. The onboard power supply was met by the solar panel. The power required by the thruster during the low thrust transfer could be supplied by lithium polymer batteries. Further feasibility studies need to be carried out to develop an empirical performance model of the state of the art thrusters system which could not be completed because of the time constraint. Page 81 of 92

Appendix A. Link Budget In telecommunications a link budget is the accounting of all the gains and losses from the transmitter to the receiver through the medium. It provides the designer with values of transmitter power and antenna gains for the various links in the system. Therefore, link budget is a key to the overall system design in space communications revealing many characteristics of the overall system performance. Link budget provides the ratio of received energy-per-bit to noise density

𝐸𝑏

𝑁0

. As the power is

limited in case of CUBESAT, it is necessary to verify that the available power is sufficient for communication to a certain level. If the system efficiency is denoted by η and the available power for communication is𝑃0 , then the power available for the transmitting antenna is given by: 𝑃𝑇 = ηP0

A1

All the calculations are done in dB where 1dB=10𝑙𝑜𝑔10

The Effective Isotropic Radiated Power (EIRP, the power that effectively leaves the antenna) is represented by: 𝐸𝐼𝑅𝑃 = 𝑃𝑇 + 𝐺𝑇 + 𝐿 𝑇

A2

𝑤ℎ𝑒𝑟𝑒, 𝐿 𝑇 𝑎𝑛𝑑 𝐺𝑇 𝑟𝑒𝑝𝑟𝑒𝑠𝑒𝑛𝑡𝑠 𝑡ℎ𝑒 𝑙𝑜𝑠𝑠𝑒𝑠 𝑎𝑛𝑑 𝑔𝑎𝑖𝑛𝑠 𝑜𝑓 𝑡ℎ𝑒 𝑡𝑟𝑎𝑛𝑠𝑚𝑖𝑡𝑡𝑒𝑟 𝑟𝑒𝑠𝑝𝑒𝑐𝑡𝑖𝑣𝑒𝑙𝑦 The space loss is calculated from the following expression: 𝐿𝑆 = �

𝜆

4𝜋𝑆

2

�

A3

𝑤ℎ𝑒𝑟𝑒, 𝜆 𝑖𝑠 𝑡ℎ𝑒 𝑤𝑎𝑣𝑒𝑙𝑒𝑛𝑔𝑡ℎ 𝑎𝑛𝑑 𝑆 𝑡ℎ𝑒 𝑝𝑜𝑤𝑒𝑟 𝑝𝑒𝑟 𝑢𝑛𝑖𝑡 𝑎𝑟𝑒𝑎 𝑎𝑡 𝑠𝑜𝑚𝑒 𝑑𝑖𝑠𝑡𝑎𝑛𝑐𝑒

The ratio of received energy per bit to noise density is given by: 𝐸𝑏

𝑁0

= 𝐸𝐼𝑅𝑃 + 𝐿𝑆 + 𝐺𝑅 − 10𝑙𝑜𝑔𝑘 − 10𝑙𝑜𝑔𝑇𝑆 − 10𝑙𝑜𝑔𝑅

A4

And the signal to noise ratio is given by: 𝑆

𝑁

= 𝐸𝐼𝑅𝑃 + 𝐿𝑆 + 𝐺𝑅 10 log 𝑘 − 10 log 𝑇𝑆 − 10 log 𝑅

A5

The link budget was calculated for CUBESAT when it reaches the perigee of the lunar orbit giving the worst case scenario.

Page 82 of 92

Table A1: Link Budget Transmitter Parameters Maximum Output Power Antenna Efficiency Antenna Gain

3 W 0.3 Linear 5 Linear

4.771213 dBW -5.22879 dB 6.9897 dB

Transmitted Signal Signal Bandwidth Carrier Signal Frequency Wavelength λ

MHz GHz m

435 0.435 0.689655172

Receiver Parameters (Ground Station) Antenna Efficiency (η) Diameter D (Assuming Receiver Antenna Circular) specifications Beam width: Theta 3dB [ = 75 * λ / D ] Gain [ = η ( πD / λ)² ] Receiver System Noise Receiver System Temp TS Gain and Noise Temperature Receiver System Gain

Transmission Path Satellite-Earth Station Distance Clear Air Atmospheric Loss Rain Loss Other Losses Free Space Path Loss

S LA LR LO LS [ = (λ/4πR )² ]

Carrier Power Results Transmitter Antenna Gain Receiver Antenna Gain Free Space Path Loss All Other Losses Earth Station Received Carrier Power

Page 83 of 92

0.55

Linear

-2.60

dB

2.00

m

25.86 45.65

Degrees Linear

16.59

dB

300.00 251.19

K Linear

24.77 24

dBK dB

4.057E+08 0.50 0.02 0.63

m Linear Linear Linear

-3.00 -18.00 -2.00

dB dB dB

1.82997E-20 Linear

-197.38

dB

1.9 16.0 1.759E-18 5.012E-03

Linear Linear Linear Linear

2.80 12.04 -177.55 -23.00

dB dB dB dB

1.203E-19

Linear

-189.20

dB

Table A2: Link Budget (Continued) Noise Power Results Boltzmann's Constant Receiver System Noise Temperature Noise Bandwidth Receiver Noise Power C/N Ratio or SNR signal-to-noise-density-ratio or carrier to noise ratio, C/No

Bit Rate Eb/N0

R

1.380E-23

J/K

300.00 430.00 6.021E-13

K MHz W

1.517E+01

Linear

1200 bits/sec 30.79181 dB -465.10 dB

Page 84 of 92

228.6 24.8 26.335 -122.20

11.8

dBW/K/Hz dBK dBHz dBW

dB

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