Potential Propulsion System for Microsatellites

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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: 27

th

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.

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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.

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Acknowledgement

I am really excited that I am writing this long awaited part of my dissertation. It has been a long journey, I suppose. I would like to take the opportunity to thank all those people who helped me to complete this journey.

First of all, I would like to thank my project supervisor Dr Kate Smith, who was kind enough to agree to advise me on this project. Her immense knowledge in the colloid thruster technology, enthusiasm, promptness and the willingness to help have been a great help for me to complete this thesis. My words cannot express my gratitude and appreciation for all the support, guidance and time you provided.

I would like to thank Professor Stark, Professor Vepa, Dr Duddeck and Professor Munjiza for making the lessons so attractive. I am also grateful to Dr M Hasan Shaheed for his invaluable advice during my post graduation in Queen Mary. A huge thanks to you for giving me the offer to work with you for my PhD.

My heartfelt thanks to Alam, my roommate and also my best friend for years. Thank you for tolerating my temper for long seven years! Also I am grateful to my best mate Arif; without his help I could not possibly complete my studies.

Baccha apu, you have inspired me to take this course of study. I wanted to be an Aerospace Engineer like you and here I am! Lets plan about that trip to moon!

Reshma apu, you are the loveliest and simplest sister in the world. I cannot express how much I will miss you.

Thank you Sona for all those messages wishing me luck. Osmosis worked! “ďakujem!”

I would also like to thank my colleague Purushoth for his continuous support during the project. I made a habit of working with you! I will miss all those offline messages in gmail. You better keep them coming!

Peter, I thank you for all those sneaky tea breaks, and particularly for all those random discussions about life, women, space and whatever we talked about!

Last, but very far from least, a huge thanks to my great family. Ammu and Bapi, you are the greatest parents one could ever have. Thanks for always believing in me. I am greatly indebted to my younger brothers-Antu and Soumik for their profound love for me. You guys are the sweetest!

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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

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Table5.6 Propulsive Requirements ………..79

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List of Figures

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

Figure 1.2 Protected Region ………... ……… … 18

Figure1.3 LISA orbits ……… ……… ……… ………19

Figure 1.4 Earth Analog spectrum ………. ……… ……… ………. 26

Figure 2.1 Schematic of a rocket device ……… ……… ……… ………….30

Figure 2.2 SSTL Xenon Gas propulsion system ……… ……… ………34

Figure 2.3 Low power Resistojet………..………...37

Figure2.4 MRIT size comparison ……… ……… …………..39

Figure2.5 MRIT system diagram ………... ……… ……… …………..39

Figure 2.6 Two-Dimensional beam current density profile ………41

Figure 2.7 MRIT thrust vs. time ………... ……… ……… …………...41

Figure2.8 MRIT mass efficiency vs. thrust at multiple propellant flow rates …..…………..42

Figure 2.9 RIT- μX elegant breadboard ……….……….44

Figure2.10 RIT-μX Performance, Specific Impulse as function of total power and thrust level ……… ……… ……… ……… ………..…… …….45

Figure 2.11 RIT- μX 50μN Thrust Stepping ……… ……… …..…………..46

Figure 2.12 Thrust Stepping -wide range ………. ……… …….…………...46

Figure 3.1: Schematic of a colloid thruster ……… ……… ………..51

Figure 3.2 : Charge Concentration change in Electric conductor ……… ……….52

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Figure3.4 Taylor cone geometry with an inner angle𝛼. ……… ……… ………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

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List of Abbreviations

ADCS Attitude Determination and control subsystem

AU Astronomical Units

EJSM Europa Jupiter System Mission

ELITE European Lisa Technology Experiment

𝐸𝑀𝐼 − 𝐵𝐹4 1 – ethyl – 3 – methylimidazolium tetrafluoroborate

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

ESA European Space Agency

EX-5 Earth Science Experimental Mission 5

FCU Flow Control Unit

FEEP Field Emission Electric Propulsion

GSO Geostationary Orbit

GSTP Gaia Science Team Program

IADC Inter – Agency Space Debris Coordination committee

𝐼𝑆𝑃 Specific Impulse

IXO International X-Ray Observatory

JAXA Japan Aerospace Exploration Agency

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LEO Low Earth Orbit

LIRE Laser Interplanetary Ranging Experiment

LISA Laser Interferometer Space Antenna

LV Launching Vehicle

MAXIM Micro Arcsecond X-ray Imaging Mission

MEMS Micro-electrical and Mechanical Systems

MRIT Miniature Radio-Frequency Ion Thruster

NAI Sodium Iodide

PM Propulsion Module

PPT Pulsed Plasma Thruster

PPU Power Processing Unit

RFIT Radio-Frequency Ion Thruster

s/c Spacecraft

SMART Small Missions for Advanced Research in Technology

SPECS Sub-millimeter Probe of the Evolution of Cosmic Structure

SSTL Surrey Satellite Technology Limited

ST-3 Space Technology 3

TPF-I Terrestrial Planet Finder Interferometer

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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

𝑚̇ 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

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∆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

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Chapter 1

1.1 Prospects of Microsatellites

The idea of microsatellites is not a very recent one. Due to the limited lifting capability of launch vehicles, spacecrafts were traditionally lighter and smaller. Vanguard 1 (85 kg), Explorer 1 (15 kg) and the first earth orbiting satellite Sputnik-1 (85 kg) provide good examples of this. With the advancement of aerospace technology, launching vehicles with heavy lifting capability have been developed (Ariane 5, Proton). During the 1980s and 1990s there was a trend of sending big satellites in space. It became a symbol of superiority of the then superpowers. However, advances in micro-fabrication and computing technologies have changed the scenario. Micro-electrical and Mechanical Systems (MEMS) have a great potential in implementing the ideas of micro and pico satellites. Recent advancement in the electronics industries has increased the functionality of the smaller spacecraft. Moreover, the budget constraints and recent government policies have resulted in a trend of decreasing satellite mass. The reduced cost of production and placing them into the orbit played a vital role in the recent advancement of microsatellites. Microsatellites make it possible to distribute various functionalities of a single spacecraft to a number of smaller satellites. This minimizes the risk and improves the reliability of the system and, at the same time, reduces the individual satellite production cost. The reduction of complexity of individual spacecraft enables rapid prototyping and also lowers the development life-cycle (Khayms, 2000).

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Table 1.1 presents the classification of satellites according to their wet mass (mass including fuel):

Table 1.1 Classification of satellites

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

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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 𝑣2cos 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

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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 ×_𝑀𝐴 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) 𝐴 𝑀 = 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.

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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.

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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

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

Propulsive Requirements

The IXO will be launched with the Ariane 5 ECA or Atlas V 551 launcher. Total launch mass is expected to be around 6500 kg. The final operating orbit will be around L2, the second Lagrangian point of the Sun - Earth system. The main advantage of an L2 halo orbit is its thermally stable environment and a good visibility of sky. The satellite will be injected to the transfer orbit towards L2 and after getting the tracking data (two days after injection) a corrective manoeuvre will be performed to minimize the launcher dispersion error. Another corrective manoeuvre will take place after 10 days of injection. The telescope will be deployed after the second corrective manoeuvre which will enable the commissioning phase of the satellite. The spacecraft will arrive in its final halo orbit approximately after 100 days after launch. Station keeping will be performed monthly which will require a total ∆𝑉 of 2 𝑚 𝑠⁄ /𝑦𝑒𝑎𝑟. The mission is expected to have a lifetime of 10 years requiring a total ∆𝑉 of approximately 120 𝑚 𝑠⁄ (Rando, 2010). Table 1.3 presents ∆𝑉 requirements for various manoeuvres during the mission lifetime of 10 years.

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

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 ±1 𝑚𝑚 𝑠⁄ and ±1 𝑎𝑟𝑐𝑚𝑖𝑛 (Reichbach, 2001). As the mission is in pre development stage, most of the mission requirements are still being researched. However, it is clear from the type and functionality of the mission that TPF-I will need very precise and accurate level of thrust to ensure an accurate position control to carry out the mission goals. The thrusters on board TPF-I satellites have to be able to provide the required ∆𝑉 for a mission lifetime of five years (or possibly ten years). Moreover, the thrust resolution and thrust level has to be within the range of micro – milli newton.

1.2.4 Summary

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

𝑃_𝑟 = 𝑚(𝑡) 𝑣(𝑡) + ∫ 𝑚̇ (𝑣(𝑡) − 𝑐)𝑑𝑡 2.1

So if gravity is the only external force, the equation of motion can be written as: 𝑑𝑃𝑟

𝑑𝑡 = −𝑚(𝑡) 𝑔(𝑡) ⇒ 𝑚 𝑑𝑣

𝑑𝑡 = 𝑇 − 𝑚𝑔 2.2

Where T is the thrust and is represented by:

𝑇 = 𝑐 𝑚̇ 2.3

As we already know form the rocket momentum equation:

𝑚.𝑑𝑉_𝑑𝑡 = 𝑚̇𝑣𝑒 + 𝐴𝑒(𝑃𝑒− 𝑃𝑎) + 𝐹𝑒𝑥𝑡 2.4

Where, 𝑚̇ = 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 ( 𝑃_𝑚̇𝑒− 𝑃𝑎) 2.5

If we integrate the momentum equation the required delta v is:

∆𝑉 = 𝑐. ln(𝑀0

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where, 𝑀₀ 𝑎𝑛𝑑 𝑀_𝑒 is respectively the initial and final mass of the system. Also the propellant mass can be written as:

𝑚𝑝 = 𝑚0− 𝑚𝑓𝑖𝑛𝑎𝑙 2.7

Now, 𝑚_𝑝 = 𝑚₀ (1 − 𝑒∆𝑉+𝑉𝑔𝑐 ) 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, 𝑚_𝑝𝑎𝑦can be obtained from the difference between the final mass and the structural mass 𝑚_𝑆 ,:

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 ∆𝑉

𝑐 2.10

The ratio ∆𝑉

𝑐 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 , 𝐼_𝑆𝑃 is a widely used term to describe the efficiency of the propulsion devices. 𝐼_𝑆𝑃 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)𝑑𝑡 𝜏𝑏 0 𝑔0 ∫ 𝑚̇(𝑡)𝑑𝑡0𝜏𝑏 = 𝐼 𝑔0 𝑀𝑝 2.11

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If the thrust is constant then we can write:

𝐼𝑆𝑃 =_𝑔₀𝑇𝜏_{𝑚 ̇ 𝜏}𝑏 _𝑏= _𝑔𝑐 𝑚 ̇𝜏₀_{𝑚 ̇ 𝜏}𝑏_𝑏

⇒ 𝐼𝑆𝑃 = _𝑔𝑐₀ 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 𝐼_𝑆𝑃 is high and the thrust is kept constant. Kinetic power in jet is given by 𝑃 =1

2 𝑚̇𝑐2, so equation 3 can be re written as:

𝑇 = 2𝑃_𝑐 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 𝐼_𝑆𝑃, 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 𝐼_𝑆𝑃 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.

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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,

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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 𝑃_𝑒 to the mass of the power plant 𝑚_{𝑝𝑜𝑤𝑒𝑟}, is defined as:

𝛼 = 𝑃𝑒

𝑚𝑝𝑎𝑦 2.15

If the efficiency of the thruster is 𝜂 then the electrical power input is:

𝑃_𝑒 = 𝛼 𝑚_{𝑝𝑜𝑤𝑒𝑟𝑝𝑙𝑎𝑛𝑡} = 1 2𝑚̇𝑣2 𝜂 = 𝑚𝑝 𝑣2 2 𝜂 𝑡𝑝 2.16

where, 𝑡_𝑝 is the time of operation or propulsive time.

Equation13, 14, 15 can be used to obtain the following relation for the payload mass fraction:

𝑚0

𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑 =

𝑒∆𝑢 𝑣⁄ 1−�𝑒∆𝑢 𝑣_{2 𝛼 𝜂 𝑡𝑝}⁄ − 1�𝑣2

2.17

Characteristic speed 𝑣_𝑐 is given by:

𝑣𝑐 = �2 𝛼 𝜂 𝑡𝑝 2.18

For a given payload fraction ( 𝑚0

𝑚𝑝𝑎𝑦𝑙𝑜𝑎𝑑) and characteristic speed (𝑣𝑐), an optimum

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

2.3.1 SSTL Low Power Resistojet

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

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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

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Page 38 of 92 2.3.2 Ion thrusters

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

2.3.2.1 Development of Miniature Radio Frequency Ion Thruster (MRIT)

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

MRIT is a promising device to ensure finite attitude control of spacecrafts requiring precision control. Moreover, it provides high Specific Impulse (𝐼_𝑆𝑃) 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.

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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 𝐼_𝑆𝑃 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.

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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 𝐼_𝑆𝑃 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 (𝐼_𝑆𝑃, 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 𝐼_𝑆𝑃 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.

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

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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.

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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

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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.

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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 𝐼_𝑆𝑃 (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

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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 𝐸

4𝜋 ∈0 (Gamero-Castano, 2000). A thin jet is created at the

Extractor Capillary

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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 𝐸_𝑛 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 , 𝜌_𝑠 can be determined by integrating the control volume indicated in the figure using Gauss’ law ∇ . 𝐸�⃗ = 𝜌_𝑐ℎ⁄ . 𝜀₀

Figure 3.2 : Charge Concentration change in Electric conductor

Therefore, for any electrical conductive liquid charge per unit area can be written as:

𝜌_𝑠 = 𝜀₀ 𝐸_𝑛 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, 𝐸_𝑛 Gas

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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(_𝑅1_𝑐1 − _𝑅1_𝑐2) 3.2

where 𝑅_𝑐1𝑎𝑛𝑑 𝑅_𝑐2 represent the principal surface radii of the curvature. The surface traction experienced by the liquid is 𝜀₀𝐸_𝑛2⁄ (Martinez-Sanchez, 2001). 2

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Figure3.4 Taylor cone geometry with an inner angle𝛼.

Meusnier’s theorem (1776) states that “all curve lying on a surface S and having at a given

point p∈ 𝑆 the same tangent line have at this point the same normal curvature”.

Therefore, 1 𝑅𝑐= � 1 𝑅� cos 𝛼 = cos 𝛼 𝑟 sin 𝛼= 1 𝑟cot 𝛼 3.3

So the surface traction can be expressed as:

𝜀0𝐸𝑛2⁄ =2 𝛾_𝑟cot 𝛼

𝐸𝑛 = �2 𝛾 cot 𝛼_𝜀₀_𝑟 3.4

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

𝛼

𝐸𝑛

𝑅 𝑅𝑐

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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_{∅ =} 1 𝑟2 𝜕 𝜕𝑟 � 𝑟2 𝜕∅𝜕𝑟� + 1 𝑟2_{sin 𝜃} 𝜕 𝜕𝜃 �sin 𝜃 𝜕∅ 𝜕𝜃� 3.5

As we need the solution outside the liquid conical section, 𝜃 is measured from inside the cone. The solution for equation 22 is in terms of Legendre polynomials:

∅ = 𝐴 𝑃𝑣 (cos 𝜃) 𝑟𝑣 3.6

∅ = 𝐴 𝑄𝑣 (cos 𝜃) 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 = − 1_𝑟 𝜕∅_𝜕𝜃= 𝐴 _{𝑑 (cos 𝜃)}𝑑 𝑄𝑣 sin 𝜃 _𝑟1−𝑣1 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 =1

2. So the solution is:

∅ = 𝐴 𝑟1/2_𝑄

1/2 (cos 𝜃) 3.9

The function 𝑄_1/2 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.

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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 𝑑_𝑐 and extractor to meniscus distance D, Eyring (1927) derived the following expression for the electric field around solid metal tips:

𝑉𝑠𝑡𝑎𝑟𝑡 = �_{2 𝜀}𝑑𝑐𝛾₀ln(4𝐷_𝑑_𝑐) 3.11

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

Potential Propulsion System for Microsatellites