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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.
Page 2 of 92
Abstract
Microsatellites have become increasingly popular with the advancement in micro-
fabrication 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.
Page 3 of 92
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 Missions Requiring Micro- propulsion .................................................................................. 20
1.3.1 LISA .................................................................................................................................... 20
Mission Overview .......................................................................................................................... 20
Propulsive Requirements .............................................................................................................. 22
1.3.2 International X – ray Observatory (IXO) ........................................................................... 23
Overview ....................................................................................................................................... 23
Propulsive Requirements .............................................................................................................. 24
Page 5 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 Microsatellite Gas Propulsion System .......................................................................... 33
2.3 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
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3.2 Related work in the area of Electro-spray ............................................................................ 58
3.3 Developments in colloid propulsion ........................................................................................... 59
3.3.1 Colloid Propulsion Research at Queen Mary ................................................................ 66
3.4 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
Page 7 of 92
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 8 of 92
Table5.6 Propulsive Requirements ………………………………………………………..79
Table 5.7 Fuel Tank Size Estimation ………………………………………………………..81
Page 9 of 92
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
Page 10 of 92
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
Page 11 of 92
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
Page 12 of 92
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
Nomenclature
𝑎 Specific power of the power plant
𝐴 Surface Area
𝑐 Effective exhaust velocity
𝐶𝑟 Coefficient of solar radiation
pressure
𝐸𝑛 Electric field
𝐸𝑡𝑖𝑝 Electric potential at the tip of the
conical surface
𝑓 Frequency
𝑓𝑠𝑡 Surface tension of the liquid
𝑔 Gravitational acceleration
m Mass
�̇� Mass flow rate
𝑚𝑝𝑎𝑦 Payload mass
𝑚0 Initial mass of the spacecraft
𝑚𝑝 Propellant mass
𝑚𝑝𝑜𝑤𝑒𝑟 Power plant mass
𝑚𝑝 Propellant mass
𝑚𝑆 Structural Mass
𝑀0 Initial mass of the system
𝑀𝑒 Final mass of the system
𝑃 Kinetic power of jet
𝑃𝑒 Pressure in the exhaust area
𝑃𝐸 Electrical power
𝑃𝑎 Atmospheric pressure
𝑃𝑟 Momentum of the mass
Varying system
Q volumetric flow rate
r radius
𝑅𝑐 Principal surface radius of the
curvature
𝑡𝑝 Time of operation or propulsive
time
T Thrust
Page 14 of 92
∆𝑉 Change of velocity
𝑣1 Initial velocity
𝑣2 Final orbit velocity
𝑣𝑐 Characteristic speed
𝑣𝑒 Exhaust velocity relative to the
vehicle
V Voltage
𝑉𝑔 Gravitational speed loss
𝜏𝑏 Burn Time
𝜂 Thruster Efficiency
𝜌𝑠 Charge per unit area
𝜆 Wavelength
𝜀0 Permittivity in vacuum
𝛾 Surface tension of the liquid
∆ 𝐻𝐺𝐸𝑂 Height above GEO for safe
Decommissioning
Page 15 of 92
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
Category Mass
Large Satellites >1000 kg
Medium Satellites 500-1000 kg
Mini Satellites 100-500 kg
Micro Satellites 10-100 kg
Nano Satellites 1-10 kg
Femto Satellites 0.1-1 kg
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 17 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
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
Transfer orbit apogee
Inclined final Orbit
∆V1
∆V2
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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 × 𝐴𝑀
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)
𝐴𝑀
= 𝐴𝑠𝑝𝑒𝑐𝑡 𝑎𝑟𝑒𝑎 𝑡𝑜 𝑑𝑟𝑦 𝑚𝑎𝑠𝑠 𝑟𝑎𝑡𝑖𝑜𝑛 (𝑚2
𝑘𝑔)
1.2.4 Orbit Phasing
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.
Page 20 of 92
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 22 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
Thrust Range 5 − 30𝜇𝑁
Thrust Resolution ≤ 0.1 𝜇𝑁
Thrust Noise ≤ 0.1 𝜇𝑁 √𝐻𝑧⁄
Thrust Response Time ≤ 100 𝑠𝑒𝑐
Specific Impulse ≥ 150 𝑠𝑒𝑐
Cluster power consumption(@30𝜇𝑁) 25 𝑊
Cluster Mass 14.6 𝑘𝑔
Lifetime (Thruster ON) 90 𝑑𝑎𝑦𝑠
Total Impulse 300 𝑁𝑠
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
Page 24 of 92
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)
Manoeuvre ∆𝑉 requirement (𝑚𝑠−1)
Perigee velocity correction 28.0
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
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
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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, 𝐶𝑂2band and 𝐻2𝑂 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
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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 ±1𝑚𝑚 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.
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Table 1.4 Propulsion requirements for a typical formation flying mission
(Collingwood et al., 2009).
Thrust range (fine) 1 – 150μN
Thrust range (coarse) 150μN to >1mN
Thrust resolution <0.5μN
Thrust noise (1mHz – 1Hz) 1.65μN/√Hz up to 100μN
Thrust linearity and bias 0.5μN
Thrust repeatability 0.5μN (0.5mN coarse)
Thrust response time 60ms
Specific power <50W/mN (coarse)
Specific impulse >1500s @1mN, >90s @12μN
Total impulse 40kNs
Beam divergence <25°
Lifetime 21900 hours
1.3 Aims of the Research Project
Electric propulsion was successfully implemented in various missions like SMART-1
and for Earth Observation or Telecommunication Satellites (Smith et al., 2009). However,
with new mission criteria arising, it is now necessary to incorporate alternative electric
propulsion devices capable of producing very low thrust resolution and very high Specific
Impulse (ISP). Various mission profiles (e.g. Earth observation satellites) may require the
satellite to stay in a low altitude for better resolution images. This results in more atmospheric
interference e.g. atmospheric drag. High thrust controllability and resolution are needed to
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compensate the atmospheric drag. High ISP and very low thrust noise are to be ensured due to
the nature of missions (Leiter et al., 2009). Missions like DARWIN, XEUS and LISA require
high precision formation flying. All these mission profiles require on board propulsion with
precise thrust modulation in 𝜇𝑁 range. In a word, the propulsion system has to be simple,
efficient and highly compact in order to be adapted to micro satellites. The aim of the project
is to identify a potential propulsion system for the micro-satellites for the future missions.
The factors affecting the scaling of the thrusters will be investigated and suggestions will be
made how to minimize those affects without compromising the thruster efficiency.
1.4 Methodology
A literature review focussing on the various propulsion systems that are currently
available has been carried out with particular importance to colloid propulsion. A survey of
various near future missions and their propulsion requirements was also undertaken. Potential
propulsion systems were identified and analyzed further. Various parameters, that affect the
performance of those thrusters, were identified. An empirical model which represents the
relation among those parameters was developed. This model was used to determine the
scaling effect on the performance of the existing thrusters. Finally, the model was fitted into a
STK simulation to justify the findings.
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Chapter 2
2.1 Available Propulsion Technologies
In order to identify a potential propulsion system for the micro satellites, we need to
understand the basic physical mechanism of various propulsion systems. This chapter gives a
brief description of chemical and electrical propulsion systems highlighting the effect of
scaling on Specific Impulse (𝐼𝑆𝑃), thrust, efficiency, lifetime and power consumption.
Due to the limited capability of launch vehicles it is necessary to optimise the
propulsion system for maximum payload. As the launcher only places the spacecraft into an
initial orbit, the onboard propulsion system provides the necessary ∆𝑉 for necessary
correction of the orbital elements. Hence the onboard propulsion system should provide the
maximum performance compared to the required weight and power.
Figure 2.1 Schematic of a rocket device
The basic operating principle of a rocket is to eject matter with kinetic energy at a
controlled rate and in a desired direction. Hence, a thrust is produced by the change of
momentum of the rocket with respect to time and this thrust is used to provide the
𝑣(𝑡)
𝑔(𝑡)
𝑚(𝑡) 𝑃𝑒
Ambient pressure
𝑃𝑎
𝑚 , 𝑐̇
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required ∆𝑉. As the rocket is a mass varying system with a velocity 𝑣(𝑡) , we can calculate
the propellant mass 𝑚𝑝 from Newton’s second law. If the mass flow rate in exhaust is �̇� and
the effective exhaust velocity is c, then the momentum of the mass varying system can be
written as:
𝑃𝑟 = 𝑚(𝑡) 𝑣(𝑡) + ∫ �̇� (𝑣(𝑡) − 𝑐)𝑑𝑡 2.1
So if gravity is the only external force, the equation of motion can be written as:
𝑑𝑃𝑟𝑑𝑡
= −𝑚(𝑡) 𝑔(𝑡)
⇒ 𝑚 𝑑𝑣𝑑𝑡
= 𝑇 − 𝑚𝑔 2.2
Where T is the thrust and is represented by:
𝑇 = 𝑐 �̇� 2.3
As we already know form the rocket momentum equation:
𝑚. 𝑑𝑉𝑑𝑡
= �̇�𝑣𝑒 + 𝐴𝑒(𝑃𝑒 − 𝑃𝑎) + 𝐹𝑒𝑥𝑡 2.4
Where, �̇� = mass flow rate in exhaust
𝑣𝑒 = Exhaust velocity relative to the vehicle
In field free space the external force can be considered 0. So the effective exhaust velocity is:
𝑐 = 𝑣𝑒 + 𝐴𝑒 ( 𝑃𝑒− 𝑃𝑎)�̇�
2.5
If we integrate the momentum equation the required delta v is:
∆𝑉 = 𝑐. ln(𝑀0𝑀𝑒
) 2.6
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where, 𝑀0 𝑎𝑛𝑑 𝑀𝑒 is respectively the initial and final mass of the system. Also the
propellant mass can be written as:
𝑚𝑝 = 𝑚0 − 𝑚𝑓𝑖𝑛𝑎𝑙 2.7
Now, 𝑚𝑝 = 𝑚0 (1 − 𝑒∆𝑉+𝑉𝑔
𝑐 ) 2.8
𝑉𝑔 represents the gravitational speed loss for the total duration of the impulse. 𝑉𝑔 = 0 when
the gravity is considered as zero.
The payload mass, 𝑚𝑝𝑎𝑦can be obtained from the difference between the final mass
and the structural mass 𝑚𝑆 ,:
𝑚𝑝𝑎𝑦 = 𝑚𝑓𝑖𝑛𝑎𝑙 − 𝑚𝑆 2.9
If we ignore gravitational forces, equation 8 clearly shows that for small values of effective
exhaust velocity, c , the payload mass becomes smaller with the large values of ∆𝑉. So we
get the following relation:
𝑚𝑝
𝑚0= 1 − 𝑒
∆𝑉𝑐 2.10
The ratio ∆𝑉𝑐
determines the performance of the propulsion subsystem significantly along with
mass and power. Although the optimisation criteria vary with the mission requirement,
normally a higher exhaust velocity results in a better performance of the thruster.
Specific Impulse , 𝐼𝑆𝑃 is a widely used term to describe the efficiency of the
propulsion devices. 𝐼𝑆𝑃 indicates the change in momentum per unit mass for the propellant.
So for a given mass flow rate, the amount of thrust can be written as:
𝐼𝑆𝑃 = ∫ 𝑇(𝑡)𝑑𝑡𝜏𝑏0
𝑔0 ∫ �̇�(𝑡)𝑑𝑡𝜏𝑏0
= 𝐼𝑔0 𝑀𝑝
2.11
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If the thrust is constant then we can write:
𝐼𝑆𝑃 = 𝑇𝜏𝑏𝑔0 𝑚 ̇ 𝜏𝑏
= 𝑐 𝑚 ̇ 𝜏𝑏𝑔0 𝑚 ̇ 𝜏𝑏
⇒ 𝐼𝑆𝑃 = 𝑐𝑔0
2.12
As 𝑔0 is the gravitational constant, 𝐼𝑆𝑃 is related to the effective exhaust velocity, c. The
mass flow rate, 𝑚 ̇ is considerably low when 𝐼𝑆𝑃 is high and the thrust is kept constant.
Kinetic power in jet is given by 𝑃 = 12
�̇�𝑐2, so equation 3 can be re written as:
𝑇 = 2𝑃𝑐
2.13
2.2 Chemical Propulsion Technology
In case of chemical rocket propulsion, a fuel and an oxidizer react in a high pressure
combustion reaction. The energy from the reaction heats up the product gases to a very high
temperature in the range of 2500℃ to 4100℃.Chemical rockets have relatively low 𝐼𝑆𝑃, very
high thrust, high acceleration and high specific power. Chemical propulsion devices require
heavy electrical power sources to produce the power need for high ejection velocities. (Sutton,
2001). The kinetic energy acting on the gas molecules on the exhaust greatly depends on the
bonding energy of the atoms of the chemicals fuels. However, the bonding energy per unit for
a specific molecule is finite, which restricts the 𝐼𝑆𝑃 to a value of ≤ 500 𝑠(Palaszewski, 1993
and Lozano, 2003).
2.2.1 Microsatellite Gas Propulsion System
Cold gas thrusters are widely used in space missions. Their operating principle is
quite simple. Gas is stored in a high pressure and is allowed to expand through nozzle to
produce thrust. Traditional adiabatic expansion relations can be used to estimate the
performance of the gas propulsion systems. The primary advantage of gas propulsion systems
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is the simple operating principle, small thrust and specific impulse. These type of thrusters
can be used to rectify launcher injection errors, to main the appropriate orbit, for station
keeping, orbit phasing or for other maintenance of orbits. They are high performance and can
be very cost effective as they do not produce any propellant movement (SSTL subsystem
Datasheet, 2010).
The following is a Xenon Gas Propulsion system for microsatellites which is
developed by Surrey Satellite Technology Limited, UK. Table 3 represents the major
specifications of the thruster. The most significant drawback of the cold gas systems is their
low ISP. The propellant weight will be dominating factor once the ∆𝑉 requirements for the
missions go significantly high (e.g. long term missions requiring high ∆𝑉). Also, the safety
issue of the valves may pose a significant risk as gaseous molecules are more mobile.
Figure 2.2 SSTL Xenon Gas propulsion system
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Table 2.1 Specifications of Xenon Gas Propulsion system (SSTL subsystem
datasheet, 2010)
Propellant 12 Kg Xenon
Thrust 10-50 mN
Maximum total impulse 5.65 kN.s
Storage Pressure 120 bar
Tank burst factor >× 4
Specific Impulse Up to 48 sec
System Volume 7.42 litres
Life duration >7 years
2.3 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.
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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 = 𝑃𝑒𝑚𝑝𝑎𝑦
2.15
If the efficiency of the thruster is 𝜂 then the electrical power input is:
𝑃𝑒 = 𝛼 𝑚𝑝𝑜𝑤𝑒𝑟𝑝𝑙𝑎𝑛𝑡 = 12�̇�𝑣2
𝜂= 𝑚𝑝 𝑣2
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:
𝑚0𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑
= 𝑒∆𝑢 𝑣⁄
1−�𝑒∆𝑢 𝑣⁄ − 1�𝑣2
2 𝛼 𝜂 𝑡𝑝
2.17
Characteristic speed 𝑣𝑐 is given by:
𝑣𝑐 = �2 𝛼 𝜂 𝑡𝑝 2.18
For a given payload fraction ( 𝑚0𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑
) 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
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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 100 bar
Specific Impulse 𝑋𝑒𝑛𝑜𝑛 − 55 𝑆𝑒𝑐
𝑁2 − 99 𝑠𝑒𝑐
𝐵𝑢𝑡𝑎𝑛𝑒 − 100 𝑠𝑒𝑐
Operation Temperature 500℃
Mass 65 gms without valves
Heater Power 50 watts @ 28 Vdc
Operating temperature −20℃ 𝑡𝑜 + 60℃
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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.
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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)
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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
from stainless steel
• Molybdenum will be used
• Assembly components were
made of Teflon
• Alumina ceramic will replace
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.
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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/𝑐𝑚2and 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)
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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.
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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.
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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.
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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 46 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 47 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 48 of 92
Table2.4 Performance and Operating Characteristics of Electric propulsion systems
(Wertz et al., 1999)
Propulsion
Type
ISP s
Thrust
Range
(mN)
Propella
nt
efficien
cy (%)
Energy
Conversion
Efficiency
(%)
Power/
Thrust
(kW/N)
Specific
Power
(kW/kg)
Thrust/
weight
Total
Impulse
(N-S)
Electro-
thermal Arcjet
450-
1500
100-2000 27-37 91-95 6-15 0.25-.5 0.003-
0.005
12,000
Electro-
thermal
Resistojet
150-
700
180-500 35 60 1.3-2 0.4-0.8 0.02-
0.05
300,000
Electro-static
Colloid
1100-
1500
0.001-0.5 75 9 0.0002 >1000
Augmented
Hydrazine
294-
304
180-300 1.5-3 0.5 0.018-
0.036
Radio
Frequency
Ion
3000-
3150
15 71-80 64 39 0.07 0.00017
Field
Emission Ion
4000-
11000
0.001-
1000
33-60
Hall Thruster 950-
1950
11-512 42-67 91-93 16-19 0.1-0.45 0.0006-
0.003
2300
Pulsed Plasma 830-
1200
0.3-0.75 7-9 80 83-100 0.003-
0.005
0.00000
4
15,000-
20,000
Page 49 of 92
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 50 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 51 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.
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 𝐸4𝜋 ∈0
(Gamero-Castano, 2000). A thin jet is created at the
Extractor
Capillary
Accelerator
Page 52 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⁄ .
Figure 3.2 : Charge Concentration change in Electric conductor
Therefore, for any electrical conductive liquid charge per unit area can be written as:
𝜌𝑠 = 𝜀0 𝐸𝑛 3.1
3.1.2 Taylor Cone:
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
Electric Field, 𝐸𝑛
Gas
Liquid
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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( 1𝑅𝑐1
− 1𝑅𝑐2
) 3.2
where 𝑅𝑐1𝑎𝑛𝑑 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)
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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, 1𝑅𝑐
= �1𝑅� cos𝛼 = cos𝛼
𝑟 sin𝛼= 1
𝑟cot𝛼 3.3
So the surface traction can be expressed as:
𝜀0 𝐸𝑛2 2⁄ = 𝛾𝑟
cot𝛼
𝐸𝑛 = �2 𝛾 cot𝛼𝜀0 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.
𝛼
𝐸𝑛
𝑅
𝑅𝑐
𝑟
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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 ∅ = 1𝑟2
𝜕𝜕𝑟
� 𝑟2 𝜕∅𝜕𝑟� + 1
𝑟2 sin𝜃 𝜕𝜕𝜃
�sin𝜃 𝜕∅𝜕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 (cos𝜃) 𝑟𝑣 3.6
∅ = 𝐴 𝑄𝑣 (cos𝜃) 𝑟𝑣 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:
r
𝜃
𝜌
𝑧
𝜑
𝑃(𝑥, 𝑦, 𝑧)
𝑦
𝑧
𝑥
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𝐸𝑛 = − 1𝑟
𝜕∅𝜕𝜃
= 𝐴 𝑑 𝑄𝑣𝑑 (cos𝜃)
sin𝜃 1𝑟1−𝑣
3.8
In order to have the normal E-field in equilibrium with the surface tension, the value
of the exponent has to be 𝑣 = 12. So the solution is:
∅ = 𝐴 𝑟1/2 𝑄1/2 (cos𝜃) 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.
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.
𝜃 𝜃
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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:
𝜀0𝐸𝑡𝑖𝑝
2
2= 2 𝛾
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:
𝑉𝑠𝑡𝑎𝑟𝑡 = �𝑑𝑐𝛾2 𝜀0
ln(4𝐷𝑑𝑐
) 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 58 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
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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
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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 25-
25- 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 61 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
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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 Nano-
electrospray colloid thruster which consisted of switchable emitter clusters. It was designed
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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
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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
(V)
1400 1600 1800 2000 2200 2400 2600 2800
Thrust (𝜇𝑁) 1.36 1.4 1.49 1.55 2.0 3.25 3.95 4.85
Standard
Variance (𝜇𝑁)
0.12 0.126 0.13 0.14 0.19 0.34 0.5 0.5
Kirtley (2002) suggested a conceptual design of a colloid micro thruster with an electro-
dynamic 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.
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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,
2007
? Formamide 502
emitters
13𝜇𝑁. 3000s 1500V
Paine,
2005
Micro
fabrication
Triethylene
glycol
(0.02 𝑆𝑖/𝑚)
4 emitters 6.8 𝜇𝑁 ? 1970 V
Velasquez,
2004
MEMS Formamide ? 100 - 450𝜇𝑁 200-350s ?
Velasquez, MEMS 𝐸𝑀𝐼 − 𝐵𝐹4 ? 1-3 𝜇𝑁 3800- ?
Page 66 of 92
2004 8500s
Xiong,
2002
MEMS Formamide
with 30%
NAI
192
copper
emitters
6.8 𝜇𝑁 ? 1-3 kV
Hruby,
2001
? Formamide
(0.5 𝑆𝑖/𝑚)
57
stainless
steel
? 20-190
𝜇𝑁
400s
Pranajaya,
1999
? Glycerol and
Iso-propyl
alcohol
Stainless
Steel
? ? ?
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 × 1𝑐𝑚2silicon 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 67 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
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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)
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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 𝐼𝑆𝑃with 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)
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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:
𝑐 = �2 𝑞 𝑉𝑚
3.12
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It is clear from equation 3.12 that the exhaust speed depends on the value of 𝑞 𝑚
. Colloid
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 = 𝑐𝑔0
, 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
high 𝑞𝑚
(Fernandez, 1994). Castano (1999) showed that droplets with 5000 C/𝑘𝑔 could be
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:
𝑉𝑚𝑖𝑛 = �𝛾 𝑅𝜀0
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 𝑞𝑚
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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 𝑡ℎ𝑟𝑢𝑠𝑡𝑇ℎ𝑟𝑢𝑠𝑡 𝑜𝑏𝑡𝑎𝑖𝑛𝑒𝑑 𝑓𝑟𝑜𝑚 𝑎𝑛 𝑖𝑛𝑑𝑖𝑣𝑖𝑑𝑢𝑎𝑙 𝑒𝑚𝑖𝑡𝑡𝑒𝑟
3.14
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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 30 × 10 × 10 𝑐𝑚
Mass 2907.262 gm
Orbit Around the moon
Lifetime 1 year (initial)
Payload USB camera
Communications UHF Amateur Band
Key Components 3-axis Reaction wheel
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.
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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
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
5.3 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 432-
438 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.
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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
and sun sensor data
1-3 Arcsec
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
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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
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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 = 𝜋2 ∆𝑖
𝑣0𝑣𝑓−cos�𝜋2 ∆𝑖�
5.1
where, ∆𝑖 = 𝑐ℎ𝑎𝑛𝑔𝑒 𝑜𝑓 𝑖𝑛𝑐𝑙𝑖𝑛𝑎𝑡𝑖𝑜𝑛 𝑜𝑓 𝑡ℎ𝑒 𝑠𝑐
𝑣0 = 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑐𝑖𝑟𝑐𝑢𝑙𝑎𝑟 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝑜𝑓 𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡
𝑣𝑓 = 𝑐𝑖𝑟𝑐𝑢𝑙𝑎𝑟 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝑜𝑓 𝑠𝑝𝑎𝑐𝑒𝑐𝑟𝑎𝑓𝑡 𝑖𝑛 𝑡ℎ𝑒 𝑓𝑖𝑛𝑎𝑙 𝑜𝑟𝑏𝑖𝑡
The initial velocity of the spacecraft is given by :
𝑣0 = �𝜇𝑒𝑟0
5.2
where 𝜇𝑒 is the gravitational parameter of earth (4 × 105 𝑘𝑚3/ 𝑠2 ) 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:
𝑣𝑓 = �𝜇𝑚𝑜𝑜𝑛𝑟0
5.3
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where 𝜇𝑚𝑜𝑜𝑛 is the gravitational parameter of earth (~5 × 103 𝑘𝑚3/ 𝑠2) 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 cos𝛽0 −𝑣0 sin𝛽0
tan�𝜋2 ∆𝑖+ 𝛽0� 5.4
⇒ ∆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 7460
Me Kg 3.1
ISP S 1500
g0 m/s2 9.8
C m/s 14700
M0 Kg 5.1
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Mf Kg 1
Burn time Sec 8640000
Mass Flow rate kg/sec 1.1574E-07
Thrust N 0.0017
Power W 85
Total Impulse N-S 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 𝑐𝑚2active solar area is required. The CUBESAT has a surface area of
300𝑐𝑚2which 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 W-
hr/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 80 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 fuel 0.000885 𝑚3
Volume of tank (5% margin) 92.92 𝑐𝑚3
Inner Radius of tank 1.5 cm
Outer Radius 2 cm
Cylinder Height 13.15 cm
Thickness 0.5 cm
Volume of Metal 72.27 𝑐𝑚3
Density of Metal (Titanium) 4.506 gm/𝑐𝑚3
Fuel Tank Structure Mass 325.65 gm
Page 81 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 82 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 + 𝐿𝑇 A2
𝑤ℎ𝑒𝑟𝑒, 𝐿𝑇 𝑎𝑛𝑑 𝐺𝑇 𝑟𝑒𝑝𝑟𝑒𝑠𝑒𝑛𝑡𝑠 𝑡ℎ𝑒 𝑙𝑜𝑠𝑠𝑒𝑠 𝑎𝑛𝑑 𝑔𝑎𝑖𝑛𝑠 𝑜𝑓 𝑡ℎ𝑒 𝑡𝑟𝑎𝑛𝑠𝑚𝑖𝑡𝑡𝑒𝑟 𝑟𝑒𝑠𝑝𝑒𝑐𝑡𝑖𝑣𝑒𝑙𝑦
The space loss is calculated from the following expression:
𝐿𝑆 = � 𝜆4𝜋𝑆
�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:
𝐸𝑏𝑁0
= 𝐸𝐼𝑅𝑃 + 𝐿𝑆 + 𝐺𝑅 − 10𝑙𝑜𝑔𝑘 − 10𝑙𝑜𝑔𝑇𝑆 − 10𝑙𝑜𝑔𝑅 A4
And the signal to noise ratio is given by:
𝑆𝑁
= 𝐸𝐼𝑅𝑃 + 𝐿𝑆 + 𝐺𝑅10 log 𝑘 − 10 log𝑇𝑆 − 10 log𝑅 A5
The link budget was calculated for CUBESAT when it reaches the perigee of the lunar orbit giving the worst case scenario.
Page 83 of 92
Table A1: Link Budget
Transmitter Parameters Maximum Output Power 3 W 4.771213 dBW Antenna Efficiency 0.3 Linear -5.22879 dB Antenna Gain 5 Linear 6.9897 dB
Transmitted Signal
Signal Bandwidth MHz 435 Carrier Signal Frequency GHz 0.435 Wavelength λ m 0.689655172
Receiver Parameters (Ground Station)
Receiver Antenna specifications
Antenna Efficiency (η) 0.55 Linear -2.60 dB
Diameter D (Assuming Circular) 2.00 m
Beam width: Theta 3dB [ = 75 * λ / D ] 25.86 Degrees
Gain [ = η ( πD / λ)² ] 45.65 Linear 16.59 dB
Receiver System Gain and Noise Temperature
Receiver System Noise Temp TS 300.00 K 24.77 dBK
Receiver System Gain 251.19 Linear 24 dB
Transmission Path Satellite-Earth Station Distance S 4.057E+08 m Clear Air Atmospheric Loss LA 0.50 Linear -3.00 dB Rain Loss LR 0.02 Linear -18.00 dB Other Losses LO 0.63 Linear -2.00 dB
Free Space Path Loss LS [ = (λ/4πR )² ] 1.82997E-20 Linear -197.38 dB
Carrier Power Results Transmitter Antenna Gain 1.9 Linear 2.80 dB Receiver Antenna Gain 16.0 Linear 12.04 dB Free Space Path Loss 1.759E-18 Linear -177.55 dB All Other Losses 5.012E-03 Linear -23.00 dB
Earth Station Received Carrier Power 1.203E-19 Linear -189.20 dB
Page 84 of 92
Table A2: Link Budget (Continued)
Noise Power Results Boltzmann's Constant
1.380E-23 J/K 228.6 dBW/K/Hz
Receiver System Noise Temperature
300.00 K 24.8 dBK
Noise Bandwidth
430.00 MHz 26.335 dBHz Receiver Noise Power
6.021E-13 W -122.20 dBW
C/N Ratio or SNR
signal-to-noise-density-ratio or carrier to noise ratio, C/No 1.517E+01 Linear 11.8 dB
Bit Rate R 1200 bits/sec 30.79181 dB
Eb/N0 -465.10 dB
Page 85 of 92
References
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droplet cone-jet mode”, PhD thesis, MIT, Feb 2005.
Castano M. G. and Fernandez, J., “Direct Measurement of Ion Evaporation of Small Ions
from Droplets”. Journal of Chem. Phys., 1999.
Chobotov, V. A., “Orbital Mechanics Second Edition”, AIAA Education Series, Reston, VA,
1996.
Clyde space Battery Data Sheet, “http://www.clyde-space.com/cubesat_shop/batteries
/279_ cubesat-standalone-battery
”, accessed on 20th July, 2010.
Collingwood, C. M. et al., “The MiDGIT Thruster: Development of a Multi-Mode
Thruster”. IEPC-2009-271, 31st International Electric Propulsion Conference, Michigan,
USA. September 20-24, 2009.
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