xperimental investigation of the micro pulsed inductive...
TRANSCRIPT
Experimental Investigation of the Micro Pulsed Inductive Thruster
David Vaughan Peters
A thesis submitted in partial fulfillment of the
requirements for the degree of
Master of Science in Aeronautics and Astronautics
University of Washington
2009
Program authorized to offer degree: Aeronautics and Astronautics
University of Washington
Graduate School
This is to certify that I have examined this copy of a master’s thesis by
David Vaughan Peters
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and that any and all revisions required by the final
examining committee have been made.
Committee Members:
___________________________________________
Robert Winglee
___________________________________________
Uri Shumlak
Date: _________________________________
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University of Washington
Abstract
Experimental Investigation of the Micro Pulsed Inductive Thruster
David Vaughan Peters
Chair of the Supervisory Committee:
Professor Robert Winglee
Earth and Space Sciences
The Micro Pulsed Inductive Thruster (μPIT) combines elements of pulsed plasma
thrusters and inductive discharges to form a new alternative to the current generation of
micropropulsion technologies. Beginning with an ablative arc as in standard PPTs,
thruster operation continues as an RF antenna couples additional energy into the arc
plasma and associated neutral gas. μPIT employs coaxial electrodes and solid Teflon
fuel. The total system mass, including the power supply, is less than 500 grams. The
RF antenna is proposed as a mechanism for increasing PPT efficiency by improving
overall mass utilization. In this thesis, the roles of the antenna in the breakdown and
energy coupling processes are examined using a variety of diagnostic tools. Langmuir
probe time-of-flight measurements show an exhaust velocity of 30 km/s, strongly
suggesting an electromagnetic acceleration process over an electrothermal one. Radial
and axial mapping of the plume density with Langmuir probes provided information
for an impulse bit estimate of about 120 nN-s. A fast-framing CCD camera captured 20
μs exposures of the breakdown process, revealing macroparticle ejection from the
Telfon face. A comparison between operation with and without antenna coupling
showed that the arc plasma density is consistently 10-40% higher with the antenna
present. The additional plasma density is drawn from a large neutral population
developed by the ablation process. Furthermore, the arc plasma exhaust velocity is
50% higher with antenna coupling. This increased velocity indicates that significant
acceleration occurs within the antenna stage of the thruster; a possible mechanism for
this additional acceleration is an induced azimuthal current crossed with a radially
diverging magnetic field, in the style of a conventional pulsed inductive thruster.
Additional effects specifically attributable to the antenna are a reduced breakdown
voltage requirement and markedly improved shot-to-shot repeatability.
i
Table of Contents List of Figures ................................................................................................................. iii
List of Tables ................................................................................................................... v
1 Introduction .............................................................................................................. 1
1.1 Formation Flying Advantages ........................................................................... 1
1.2 Example Missions Requiring Micropropulsion ................................................ 3
1.2.1 LISA .......................................................................................................... 3
1.2.2 TechSat 21 ................................................................................................. 5
1.2.3 XSS-11 ....................................................................................................... 5
1.3 Leading Micropropulsion Concepts .................................................................. 7
1.3.1 FEEP .......................................................................................................... 8
1.3.2 Colloid Thruster ....................................................................................... 10
1.3.3 PPT .......................................................................................................... 13
1.4 Micro Pulsed Inductive Thruster ..................................................................... 17
1.4.1 μPIT Overview ........................................................................................ 17
1.4.2 Gas-fed μPIT ........................................................................................... 18
1.5 Review of PPT Development .......................................................................... 20
1.6 Outline and Research Goals ............................................................................ 25
2 Experimental Setup and Thruster Design .............................................................. 26
2.1 Vacuum System .............................................................................................. 26
2.2 Control and Data Acquisition ......................................................................... 27
2.3 μPIT Power Processing Unit ........................................................................... 31
2.3.1 Circuit Description ................................................................................... 32
2.3.2 Bench Testing and Tuning Procedure ...................................................... 34
2.4 μPIT Mechanical Design ................................................................................ 35
3 Diagnostics ............................................................................................................ 38
3.1 Rogowski Coil and Stangenes Probe .............................................................. 38
3.2 Voltage Dividers ............................................................................................. 40
ii
3.3 Langmuir Probes ............................................................................................. 42
3.3.1 Standard Double Probe Theory ............................................................... 42
3.3.2 Crossed-Probe Theory ............................................................................. 48
3.3.3 Modified Theory for Double Probes in Crossflow .................................. 50
3.3.4 Probe Construction and Description ........................................................ 52
3.4 CCD Camera ................................................................................................... 55
4 Theoretical Models and Data Reduction ............................................................... 56
4.1 Inductive Discharges ....................................................................................... 56
4.2 Estimated Ionization Rates .............................................................................. 61
4.3 μPIT Performance Estimation ......................................................................... 63
5 μPIT Development and Discussion ....................................................................... 66
5.1 8:120 Transformer ........................................................................................... 66
5.2 8:255 Transformer ........................................................................................... 72
5.3 8:185 Transformer ........................................................................................... 79
5.3.1 Late-time Plasma Production Results ...................................................... 81
5.3.2 Arc Plasma Density Results ..................................................................... 86
5.3.3 Specific Antenna Effects ......................................................................... 92
5.3.4 μPIT Performance Estimates ................................................................... 99
6 Summary and Recommendations ........................................................................ 103
Bibliography ................................................................................................................ 106
iii
List of Figures
Figure 1-1: LISA mission topology ................................................................................. 4
Figure 1-2: FEEP schematic ............................................................................................ 9
Figure 1-3: Colloid thruster schematic .......................................................................... 11
Figure 1-4: PPT schematic ............................................................................................. 13
Figure 1-5: μPIT schematic. .......................................................................................... 18
Figure 1-6: CAD model of the gas-fed μPIT ................................................................. 18
Figure 2-1: Bell jar vacuum chamber ............................................................................ 26
Figure 2-2: Data acquisition arrangement ..................................................................... 28
Figure 2-3: Langmuir probe range of motion ................................................................ 29
Figure 2-4: View along the z-axis of the technique for measuring radial position ....... 30
Figure 2-5: Spice model of the μPIT power supply. ..................................................... 32
Figure 2-6: Pictures of μPIT with and without igniter capacitors and insulation .......... 35
Figure 2-7: Electrode configuration showing the sharpened cathode ........................... 37
Figure 3-1: Rogowski coil schematic ............................................................................ 38
Figure 3-2: Equivalent geometry for an idealized Rogowski coil ................................. 39
Figure 3-3: Voltage divider circuit diagram .................................................................. 41
Figure 3-4: Double probe potential diagram ................................................................. 43
Figure 3-5: Theoretical I-V characteristic for a symmetric double Langmuir probe .... 47
Figure 3-6: Ion current collection theory of Kanal for thin sheaths .............................. 49
Figure 3-7: Normalized I-V characteristic for a double probe ...................................... 52
Figure 3-8: Symmetric double Langmuir probe ............................................................ 53
Figure 3-9: Side and bottom views of the planar double Langmuir probe .................... 53
Figure 4-1: Typical arrangement for an inductively coupled plasma ............................ 57
Figure 4-2: Snapshot of the magnetic field geometry near the electrodes .................... 60
iv
Figure 5-1: Typical density traces for the 8:120 μPIT .................................................. 68
Figure 5-2: Evidence of late-time plasma production ................................................... 69
Figure 5-3: Antenna loading in μPIT ............................................................................. 71
Figure 5-4: Histogram comparing 8:120 and 8:255 breakdown times .......................... 73
Figure 5-5: Time-resolved igniter voltage for the 8:120 and 8:255 transformers ......... 75
Figure 5-6: 8:255 plasma density measured at 5 cm downstream, on-axis. .................. 76
Figure 5-7: 8:255 μPIT plasma density variation with storage cap voltage .................. 77
Figure 5-8: Photographs of the for PPT and PIT configurations ................................... 79
Figure 5-9: A 200 μs exposure that captured an arc in the electronics .......................... 81
Figure 5-10: LTPP comparison between the 8:185 and 8:120 μPIT versions ............... 82
Figure 5-11: Time-resolved images of the 8:185 transformer in PIT mode .................. 83
Figure 5-12: Time-resolved images of the 8:185 transformer in PPT mode ................. 85
Figure 5-13: Quartz tube with residue after PPT mode testing ..................................... 86
Figure 5-14: Light emission from PIT amd PPT configurations ................................... 87
Figure 5-15: PIT mode radial density profile at z = 5 cm downstream ......................... 88
Figure 5-16: PIT mode radial density profile at z = 15 cm downstream ....................... 88
Figure 5-17: PPT mode radial density profile at z = 5 cm downstream ........................ 89
Figure 5-18: PPT mode radial density profile at z = 15 cm downstream ...................... 89
Figure 5-19: Axial density profile comparing the 8:185 PIT and PPT modes .............. 92
Figure 5-20: Breakdown time histogram for 8:185 PIT and PPT modes ...................... 93
Figure 5-21: Minimum breakdown voltage for PIT and PPT modes ............................ 96
Figure 5-22: Exhaust velocities at minimum voltages for PIT and PPT modes ............ 97
Figure 5-23: Peak plasma density dependence on charge voltage ................................ 98
Figure 5-24: Representative storage capacitor voltage waveform .............................. 101
v
List of Tables
Table 1-1: Micropropulsion mission requirements and mission class comparison ......... 6
Table 1-2: Comparison of FEEP, colloid, and PPT micropropulsion technologies ...... 16
Table 2-1: Selected dimensions of the microthruster assembly .................................... 36
Table 3-1: Langmuir probe operating parameters ......................................................... 54
Table 5-1: Electrical characteristics of the 8:120 μPIT ................................................. 66
Table 5-2: Electrical characteristics for the 8:120 and 8:255 μPIT models .................. 72
Table 5-3: Electrical characteristics of three variants of the 8:185 μPIT ...................... 80
Table 5-4: Exhaust velocities at 350 V charge for three 8:185 operating modes .......... 94
Table 5-5: Exhaust velocity measurements for the low current PIT configuration ....... 95
Table 5-6: Estimated impulse bits for PIT and PPT modes........................................... 99
Table 5-7: Energy and shot efficiency comparison for PIT and PPT modes .............. 101
vi
Acknowledgements
I would like to thank the following people for their support and guidance:
Prof. Robert Winglee for helping me chart a course, academic and otherwise
Dr. Tim Ziemba for tirelessly aiding my progress, even while juggling several
other projects at EHT
Prof. Uri Shumlak for serving on my committee and for providing invaluable
guidance
All the graduate and undergraduate students who offered suggestions,
participated in the project, and helped me get μPIT running: Dr. Jim Prager,
Race Roberson, Keith Cowan, Tina Saloutos, and Shawn Campbell
My parents, Don and Marlene Peters, for providing me with everything I
needed to succeed
And finally, Heather Smith, whose unwavering support saw me through this
work
1
1 Introduction
The field of micropropulsion has enjoyed rapid growth in recent years due to increasing
interest in spacecraft missions involving small spacecraft and spacecraft formations. As
mission designers shift from the large, monolithic approach toward these new
architectures, challenging design requirements have emerged. Dramatic reductions in
subsystem mass, volume, and power are essential, as is an increased degree of integration
and cooperation among subsystems. The continuing development of micropropulsion
subsystems, in particular, will serve to enable new classes of missions. Some examples
include space-based interferometers for discovering extrasolar planets and nanospacecraft
swarms for distributed, in-situ measurements of the magnetosphere. In addition to science
missions, micropropulsion technology can also serve as an enabler for military mission
concepts like space-based radar.
Leading micropropulsion concepts include Field Emission Electric Propulsion (FEEP),
colloid thrusters, and Pulsed Plasma Thrusters (PPT). FEEP and colloid thrusters are
variations on electrostatic accelerators best suited to micro-Newton level thrust
applications. PPTs are highly scalable electromagnetic accelerators with a long flight
history. Among these options, PPTs are uniquely flexible due to their pulsed operation,
use of solid fuel, and inherent simplicity. However, PPT performance suffers from poor
propellant utilization efficiency. This thesis describes development efforts underway at
the University of Washington with the goals of improving PPT propellant utilization by
means of RF energy coupling while reducing the overall micropropulsion system mass by
means of a compact, solid-state power processing unit (PPU).
1.1 Formation Flying Advantages
Beginning in the early 1980s, engineers and scientists have speculated about the possible
advantages of employing clusters of smaller spacecraft in place of single, massive
spacecraft [Burke 1981]. Working together, these smaller spacecraft offer greater mission
flexibility, lower cost, lower risk, and potentially greater performance. Small spacecraft
are individually cheaper and faster to build, and when common designs and components
2
are used the overall expense for several small spacecraft can be less than the
corresponding single, large satellite.
For science missions, individual instruments often interfere with one another or present
competing requirements. A formation of small spacecraft can relax these requirements by
isolating individual instruments on each spacecraft, enabling each instrument to collect
data over a greater portion of the total mission time. The modularity of this approach also
allows for a more flexible launch schedule. Individual spacecraft can be launched as their
instruments are completed and the formation can come online gradually. Furthermore,
launching the mission components in pieces this way serves to spread out risk should a
particular launch vehicle fail.
The inherent flexibility of small spacecraft flying in formation is an important asset on
orbit as well. A formation could reconfigure itself dynamically for different missions. For
example, an interferometry mission consisting of several spacecraft could adjust the
spacing between vehicles depending on the current target. Tight, short baseline
formations are better for smaller scale objects, while longer baselines work better for very
large or distant objects. The capability to reconfigure dynamically enables a formation to
continue functioning even as individual spacecraft begin to fail, thus maximizing the
science return. Similarly, the formation architecture creates the opportunity for upgrades
such as new spacecraft, which can be integrated relatively easily.
Applications of formation flying spacecraft include various remote sensing concepts like
interferometers and synthetic aperture radars. Interferometers are valuable tools for
studying astrophysical phenomena and for searching for extrasolar planets, as in the
Terrestrial Planet Finder (TPF) mission. Synthetic aperture radars could be employed for
a wide variety of earth observation tasks, such as studying changes in the environment,
monitoring compliance with international treaties, and space-based surveillance
[Kristiansen 2008]. Spacecraft flying in formation functioning as an interferometer or a
synthetic aperture can collect not only more data but also collect data with greater
sensitivity than could a single, large spacecraft. An additional application for formation
flying technology is autonomous rendezvous for inspection, repair, etc.
3
1.2 Example Missions Requiring Micropropulsion
Mission concepts involving formation flying require advances in the fields of
autonomous control and in micropropulsion. While the control problems associated with
formation flying is a rich, rapidly growing field of research, this chapter will focus on
micropropulsion requirements and current options.
Very often, spacecraft will require propulsion on different scales to handle a variety of
maneuvering tasks. For small spacecraft, ~10 kg, tasks like stationkeeping, drag makeup,
and orbit raising lead to minimum thrust requirements on the order of 1 mN. On the other
hand, attitude control needs for small spacecraft lead to impulse bit requirements on the
order of 1 μN-s. These numbers are borrowed from Mueller [2000] and should only be
taken as an order-of-magnitude estimate because they are computed based on a non-
specific, straw-man mission. However, the important point here is the large difference
between thruster requirements for stationkeeping and for attitude control. A
micropropulsion system that can provide higher thrust for orbit raising, etc. and also
throttle down to provide accurate attitude control at the 1 uN-s level would be at a great
advantage. The following are some specific missions that illustrate the variety of
applications and requirements for micropropulsion systems.
1.2.1 LISA
The Laser Interferometer Space Antenna, or LISA, mission is a joint NASA/ESA
program that will search for gravitational waves. As predicted by general relativity,
gravitational waves are generated by accelerating masses. The largest signals are
expected from rapidly accelerating objects like binary stars and very massive objects like
black holes. In order to detect these waves, the LISA mission concept calls for three
identical spacecraft flying in an equilateral triangle formation as illustrated in Figure 1-1.
Separation between spacecraft will be about five million kilometers. The formation will
orbit at 1 AU and follow about 20 degrees behind the earth. A cubic proof mass will fly
freely inside each spacecraft, whose structure will shield each proof mass from outside
disturbances such as photon pressure and solar wind. LISA’s so-called Disturbance
Reduction System will utilize a capacitance measurement to determine its motion relative
to the proof mass and fire micronewton thrusters in order to avoid contact with the proof
4
Figure 1-1: LISA mission topology, illustrating the triangular formation orbiting the sun and a
gravitational wave source. Image reproduced with courtesy to JPL (lisa.jpl.nasa.gov)
mass. Laser interferometry is used to precisely measure the position of proof masses
relative to each other. This way, the LISA formation can detect gravitational waves as
they produce small oscillations in the relative positions of the proof masses.
In order to advance necessary technologies and reduce risk, ESA will launch the LISA
Pathfinder mission in late 2009. LISA Pathfinder will test two distinct Disturbance
Reduction Systems: a European design, based on FEEP microthrusters and an American,
based on colloid microthrusters [Stebbins 2008]. After downselecting the DRS and other
technologies, LISA is scheduled for launch in 2018. Detecting the tiny changes in
separation generated by gravitational waves requires a level of thruster precision never
achieved before. Requirements for individual thrusters include a nominal thrust of 20 μN,
thrust resolution of 0.1 μN, and a total impulse of 3000 N-s [Cardiff et al. 2004]. Thrust
noise cannot exceed 0.1 0. 𝜇𝑁 𝐻𝑧. An additional requirement unique to LISA is a
minimal change in the spacecraft’s self-gravity over the course of the mission, which in
turn requires high specific impulse to minimize the change in mass as propellant is
expelled. While the thrust requirements are identical for the pathfinder and for LISA
itself, a key challenge is the greatly increased operating lifetime. Pathfinder calls for 2200
hours of operation and 300 N-s of impulse per thruster, while LISA’s microthrusters must
operate for 55,000 hours and generate about ten times the total impulse [Ziemer et al.
5
2006]. Lifetime tests to date of the FEEP and colloid microthruster under consideration
for LISA have only lasted a few thousand hours [Stebbins 2008].
1.2.2 TechSat 21
Another mission employing a formation of spacecraft is the TechSat 21 mission. One of
many incarnations of the central mission concept, TechSat 21 was a DOD-funded space-
based radar. Approximately twelve 135 kg spacecraft would form a synthetic aperture
100 meters in diameter in order to provide an advanced surveillance capability. While
this particular program was cancelled in 2003 due to cost overruns, the idea has been
proposed several times over the past few decades and will undoubtedly show up again
[Schilling et al. 2000]. In fact, the micropropulsion technologies currently under
development may help enable a more affordable version of the TechSat 21 mission.
Like the LISA mission, the TechSat 21 development program called for a precursor
mission involving just three spacecraft operating for just one year, followed by the full
mission continuing for ten years. Once again, thrust requirements were similar for the
precursor and full mission, with the mission duration being the only significant
difference. Propulsion requirements for TechSat 21 include a total of 390 m/s ΔV for
orbit raising, stationkeeping, formation maneuvering, and de-orbit. The stationkeeping
requirement is about 20 m/s per year. Assuming three pairs of dedicated ACS thrusters,
just 47 N-s of impulse is required per thruster, or equivalently, a total of about 2 m/s ΔV.
This translates to a thrust requirement of 38 μN per thruster.
A study by Schilling, et al. [2000] recommended a combination of hydrazine thrusters
and PPTs for the precursor mission. However, chemical propulsion was found to be
inadequate for the more demanding full TechSat 21 mission due to very high propellant
mass. The study suggested either a PPT or a low-power Hall thruster for main propulsion
and a micro-PPT for attitude control.
1.2.3 XSS-11
The Defense Department’s Experimental Satellite System (XSS) program is
representative of a third class of missions enabled by micropropulsion. XSS-11 is a single
100 kg satellite that launched in 2005. Over its mission lifetime, the spacecraft performed
6
autonomous rendezvous and proximity operations (30-100 m separation) with several
inactive US satellites and relayed images to the ground in real time [AFRL 2005]. This
technology demonstration opened a path toward several military and commercial
applications. Small satellites like XSS-11 could be used for inspection and repair of
friendly spacecraft, lengthening the lifetime of larger spacecraft and increasing the return
on the original investment. On the other hand, this technology could also be used for
surveillance or even a weapon against other satellites. The spacecraft’s size would make
it harder to detect in such an application.
This mission was similar in several ways to the better-known, DARPA-funded Orbital
Express mission, which involved not only autonomous rendezvous but also docking and
on orbit servicing—in this case, propellant exchange and flight computer replacement.
An important difference, though, is the difference in size: the Orbital Express mission
was composed of a 2100 kg servicing spacecraft (Astro) and a 500 kg mock serviceable
spacecraft (NextSat). XSS-11 is tiny by comparison, and therefore a much cheaper option
for many of the applications described earlier. While Astro employed 0.8 N hydrazine
thrusters for fine maneuvering and attitude control [Dipprey et al. 2003], a smaller
spacecraft like XSS-11 could perform these tasks with a micropropulsion system. Some
performance requirements for this mission class and the other two described previously
are listed in the table below.
Table 1-1 presents some details from each of these missions for comparison. There are
numerous other examples of formation-style missions that fit in these three general
mission classes. JPL’s Terrestrial Planet Finder mission could employ an interferometry
Table 1-1: Micropropulsion mission requirements and mission class comparison
LISA [Cardiff 2004]
Techsat 21 [Schilling, et al. 2000]
XSS-11 [AFRL 2005]
Mission class Interferometer Earth Observation
Autonomous rendezvous
Domain Science Military Military Spacecraft mass [kg] 575 135 100
Spacecraft per formation 3 12 2 Mission lifetime [years] 5 + 10 1-2 Fine control thrust [μN] 1-100 38 -
μThruster total impulse [N-s] 3000 47 -
7
formation similar to LISA [Katz 2005]. JPL’s current A-train constellation consists of a
loose string of earth-observing science spacecraft in low earth orbit, with some spacecraft
spaced as little as 15 seconds apart. The DOD is funding several longer-term small
satellite formation programs such as Tiny Independent Coordinated Spacecraft and
Fractionated Spacecraft System [Dept. of Defense 2007]. Both call for autonomous,
dynamically reconfiguring formations consisting of many 1-10 kg spacecraft. While not
an exhaustive list, these missions serve to illustrate the wide variety of applications
enabled by the combination of micropropulsion technologies and formation flying
architectures.
The missions in Table 1-1 call for up to 12 spacecraft per formation, and this figure is
likely to increase as microelectromechanical systems (MEMS) advance. Also noteworthy
is the mission lifetime of up to ten years, which poses a serious challenge to
micropropulsion development programs considering that micronewton thrusters may
need to operate for a large fraction of the total mission time. This high duty cycle is
responsible for the LISA total impulse per thruster requirement being so much larger than
the Techsat 21 total impulse requirement. A significant conclusion drawn from the table
is that the thrust required for fine maneuvering and attitude control of small satellites is
on the order of 10 μN. This matches roughly with the general mission requirements for
microspacecraft developed by Mueller [2000]. Finally, specific details about the XSS-11
subsystems are presumably classified, so the propulsion requirements are unavailable.
1.3 Leading Micropropulsion Concepts
A broad range of thruster choices are possible for micropropulsion development, ranging
from cold gas to electric. Although some development efforts are underway for
chemical/cold gas microthrusters [Mueller 2000 and Scharlemann et al. 2007], the
inherent specific impulse limitations associated with chemical and cold gas propulsion
most often lead to intolerably high propellant mass, particularly for very small spacecraft
(10 kg or less). Chemical and cold gas are competitive only in a small subset of
microspacecraft programs characterized by high thrust and short mission duration.
Furthermore, miniaturization of both of these systems leads to several fundamental
8
feasibility questions, including high propellant leak rates in microvalves, thermal control,
and increased viscous losses as the scale length decreases [Ketsdever 2000].
Based on these effects, electric propulsion is preferable in most cases for microsatellite
propulsion, especially for the precise maneuvering and attitude control required by many
formation flying architectures. However, not all flight-proven electric propulsion systems
favorably scale down to lower power and thrust. In particular, the gas discharge process
in ion engines and Hall thrusters results in increasingly severe efficiency losses as the
characteristic scale length decreases. As the ratio of discharge chamber volume to surface
area decreases, a larger fraction of free electrons collide with the walls before
experiencing ionizing collisions and more ions strike the walls and undergo charge
exchange without contributing much thrust. Because of this limitation, even scaled-down
Hall and ion thrusters are unlikely to be small enough to meet power consumption and
attitude control requirements for microspacecraft. For example, a micro-Hall thruster
developed at MIT consumes 50 W and generates about 1 mN of thrust at an efficiency
around 5%, which is much lower than the efficiency of a larger Hall thruster and related
to the wall losses mentioned earlier [Khayms et al. 2000].
Three styles of propulsion systems stand out as particularly attractive options. Pulsed
plasma thrusters are highly scalable and come with a long flight history. FEEP thrusters
have some flight history and offer precise thrust at very high specific impulse due to their
electrostatic acceleration mechanism. Colloid thrusters can also provide very precise
thrust at lower Isp and power.
1.3.1 FEEP
Field emission electric propulsion is based on electrostatic acceleration and employs
liquid metal fuel. FEEP thrusters are well suited to micropropulsion applications and
have demonstrated a thrust of 25 μN using around 5 W. Figure 1-2 shows the general
method of operation for a FEEP. A liquid metal propellant, usually indium or cesium, is
fed through a needle either by positive pressure or capillary action. A potential difference
on the order of 10 kV between the needle and the extractor generate an extremely high
electric field (about 109 V/m) at the needle tip [Genovese, et al. 2001]. A balance
between surface tension and electrostatic forces develop a protrusion in the liquid metal
9
Figure 1-2: FEEP schematic. In this design, a single electrode serves a dual purpose: it extracts ions
from the liquid metal pool and accelerates them to tens of km/s. Image reproduced from [Tajmar et al.
2004]
known as a Taylor cone. At the tip of the Taylor cone, individual ions are extracted and
accelerated through the extractor electrode to exhaust velocities as high as 120 km/s.
Finally, the exhaust must be neutralized in order to avoid spacecraft charging.
Two FEEP designs are under development in Europe for LISA. The Austrian Research
Center (ARC) is pursuing an indium FEEP, while Italy’s Centrospazio is working on a
cesium FEEP. Both propellants have the advantages of good wetting properties for easy
propellant feeding, low ionization potential, and being easily storable in light tanks.
Additionally, cesium has a high atomic mass, which improves the thrust generated by
electrostatic acceleration. Indium’s atomic mass is also quite high, though not as high as
cesium. Both metals have a low melting point: cesium melts at 83 ̊F (28 ̊C) while indium
melts at 313 ̊F (156 ̊C). The trade off here is the reduced heater power requirement for
cesium versus the possibility of a liquid propellant during launch. Indium is clearly solid
during launch, which simplifies integration. Cesium is highly reactive with the water
vapor in air and requires special and expensive handling as a result, while indium is
comparatively much safer. The reactivity of cesium also presents thruster and spacecraft
contamination concerns. Finally, the vapor pressure of cesium at typical operating
temperatures is about ten orders of magnitude greater than the vapor pressure of indium
10
in the same temperature range. This means that cesium presents a greater danger for an
electrical breakdown event and that it experiences more propellant losses due to
evaporation [Mitterauer 2001].
In addition to the propellant advantages listed above for indium, FEEP thrusters also have
a very high specific impulse and high thruster efficiency (up to 95%), both due to the
acceleration mechanism. FEEPs are simple, with no moving parts and a passive
propellant feed system. The ARC indium FEEP has demonstrated thrust resolution of just
10 nN and a thrust range of 1-100 μN for a single emitter. Of course, several emitters can
be clustered to together to generate more thrust.
Some limitations for FEEPs include the need for a heater and a neutralizer, both of which
reduce the overall system electrical efficiency. As a gas discharge device, the
conventional hollow cathode has poor efficiency at very low current and simple
thermionic neutralizers have high power requirements. A high efficiency field emission
neutralizer could be developed for this purpose. An issue that may limit FEEP lifetimes is
self-contamination, particularly in the case of cesium propellant. Reaction products can
clog the emitter, leading to thruster failure. Additionally, recent work at ARC has shown
that above a critical extraction current, the indium FEEP begins emitting charged droplets
in addition to ions. Ions generate nearly all of the thrust, so the droplets represent wasted
mass. The propellant utilization efficiency drops off sharply from 100% at 20 nA to about
25% at the full 100 μN thrust rating [Tajmar et al. 2004]. As a result, throttling the thrust
over the range required for attitude control and maneuvering (10 μN to 1 mN) is only
possible with very poor mass utilization. Finally, the high voltage requirements for ion
extraction can lead to massive power supplies even though the power processed is on the
order of a few watts.
1.3.2 Colloid Thruster
Colloid thrusters are very similar to FEEP thrusters due to their common acceleration
mechanism and propellant feeding technique. As illustrated in Figure 1-3, colloid
thrusters draw conductive liquid propellant though a small emitter structure. Propellant
can be actively fed by a bellows or drawn passively by capillary action [Ziemer et al.
11
2006]. Like FEEPs, a strong electric field is applied in the region near the emitter tip and
the propellant forms a Taylor cone. However, rather than individual ions, colloid
thrusters extract and accelerate charged droplets. A typical propellant is glycerin doped
with sodium iodine [Mueller 2000]. This key difference in emitted particles enable
colloid thrusters to trade the high Isp of FEEPs for higher thrust. Because they expel
droplets instead of individual metal ions, the charge to mass ratio of colloid thruster
propellant is generally lower than for FEEPs or ion engines. The higher mass per particle
results in lower specific impulse in exchange for higher thrust density and a greater thrust
to power ratio than can be achieve with FEEPs. A typical colloid thrust Isp is 1000
seconds, while FEEPs can achieve 10,000 seconds or more [Ziemer et al. 2006].
Although self-neutralizing colloid thrusters are possible through the use of two kinds of
dopants (for positive and negative ions) and AC grid operation, practical colloid thrusters
to date have required a separate neutralizer.
Colloid thrusters share several advantages with FEEP systems. Their use of a storable
liquid propellant greatly simplifies propellant handling, reduces tank mass, and eases
constraints on thermal control. They scale well to low power and thrust because the avoid
gas discharges. As electrostatic accelerators, colloid thrusters can accelerate propellant
Figure 1-3: Colloid thruster schematic. This design shows separate extractor and accelerator grids.
As an electrostatic accelerator, the colloid thruster generally requires a neutralizer. Image reproduced
from [Ziemer et al. 2004]
12
very efficiently. As noted earlier, the use of charged droplets instead of ions for
propellant leads to lower Isp but higher thrust than FEEPs. Colloid propellant is in
general a more benign spacecraft contaminant than metal ions, especially cesium. The
Colloid Micro-Newton Thruster (CMNT) developed by Busek is a state-of-the art
micropropulsion system developed for LISA. Consuming around one watt, it can
generate thrust over a range of 5 to 30 μN with a resolution of about 100 nN and very low
thrust noise [Ziemer et al. 2006].
With an available Isp range of 500-1500 seconds, colloid thrusters require higher
propellant mass in comparison to other electric micropropulsion technologies. Despite the
low power requirements, the high voltage necessary for extracting the charged droplets
can lead to a relatively massive PPU. In addition, colloid thrusters generally require
additional power for a neutralizer and a heater for the propellant reservoir. Like FEEPs,
the thrust per emitter is only variable over a small range. Unless many emitters are
clustered together, a micro colloid thruster cannot throttle up enough to handle both
stationkeeping/maneuvering propulsive requirements (mN thrust level) and fine attitude
control (μN thrust level). In fact, Busek is developing a 1 mN-class colloid thruster based
on the same technology as the CMNT in order to accommodate this second set of
maneuvering requirements.
Some possible areas of improvement for future colloid thruster research are lifetime and
propellant utilization. Similar to FEEPs, colloid thrusters can self-contaminate as droplets
strike the emitter grid (termed overspray) or get stuck on the emitter. These effects tend
to reduce thruster lifetime. In addition, bubbles have been observed in the propellant
stream through the emitter in the CMNT. These bubbles disrupt the flow of propellant
and generate an unpredictable impulse spike. The bubbles are believed to be caused by a
small amount of water mixed in with the propellant, and specially developed propellant
and tank handling methods have reduced their frequency. Busek has redesigned the
electrodes to reduce the effects of overspray. In addition, Busek and JPL have
successfully demonstrated the CMNT system’s stability over a 3000 hour lifetime test
[Ziemer et al. 2006].
13
1.3.3 PPT
Pulsed plasma thrusters offer a third option for micropropulsion missions requiring high
specific impulse. The PPT is a combined electromagnetic and electrothermal accelerator
that employs solid Teflon propellant, giving it significant systems-level advantages over
other propulsion technologies. Teflon is comprised of long polymer chains with two
fluorine atoms per carbon. Figure 1-4 shows a simple, breech-fed PPT design with
rectangular electrodes. A capacitor is charged to high voltage (typically a few kV), then a
spark plug is fired to initiate a discharge across the surface of the Teflon. A standard-
sized PPT will store tens of Joules in the main capacitor and develop a peak current in the
kiloamp range. This current pulse ablates material from the propellant surface, generating
a Teflon plasma. Flowing upward in the figure, the current generates a magnetic field.
The resulting 𝑗 × 𝐵 force drives the current sheet along the electrodes and expels the
ablated propellant at high velocity, generating thrust.
However, not all the ablated propellant is entrained in the current sheet, leading to a
significant spread in the exhaust velocity distribution and a large fraction of the total shot
mass exiting at low speed. While the electromagnetically accelerated plasma can be
ejected at tens of km/s, slower particles exit at only a few hundred m/s. Some designs,
such as Illinois gasdynamic PPT, forgo the electromagnetic acceleration in favor of an
Figure 1-4: Notional PPT schematic illustrating the high voltage capacitor, parallel electrodes,
current path, field geometry, and a propellant feed mechanism. Image reproduced from [Busek Co.
2007b]
14
electrothermal approach [Burton et al. 1999]. The current pulse in this case is used to
resistively heat ablated propellant. This leads to a coaxial electrode, side-fed thruster
geometry in which the nozzle becomes very important for converting enthalpy to thrust.
This approach necessarily results in lower Isp but higher thrust. Several different
electrode geometries are possible, with the rectangular and coaxial being the two most
common.
PPTs have a great deal of flight heritage, albeit mostly on a larger scale than would be
appropriate for micropropulsion applications. The first flight model flew on the Soviet
Zond 2 spacecraft in 1964, and PPTs have been studied and improved in the decades
since. In addition to the reduced risk derived from a long flight heritage, PPTs have an
advantage over other micropropulsion systems because of the use of solid fuel. Solid fuel
allows for many benefits related to systems integration, including easy propellant
handling and the elimination of valves, feedlines, propellant tanks, and other fluid system
hardware. PPTs have a very low dry mass due primarily to the solid fuel. Other
advantages stemming from the use of solid fuel include [Burton et al. 2000]:
No warm-up time required
Vacuum compatible
Long shelf life
Chemically inert/safe fuel
Minimal temperature requirements and insensitive to rapid changes in
temperature
The pulsed nature of the thruster is another key advantage. Pulsed thrusters can deliver
variable thrust at optimum operating conditions by simply adjusting the pulse frequency.
As a result, PPTs can operate over a wide range of input power. The ability to provide
discrete impulse bits is particularly useful for attitude control and fine maneuvering, and
small scale or μPPTs are well-suited to this task. Because of the potential for variable
thrust over a wide range, a single μPPT propulsion system may be able to meet the
requirements for both attitude control and stationkeeping/orbit maneuvering.
15
Probably the most severe drawback to the use of PPTs is the poor efficiency, which
generally falls in the range of 5-15%. This is due in large part to poor mass utilization
efficiency; PPTs only use a small fraction of the total ablated mass to produce useful
thrust. Much of the ablated mass shows up as a neutral gas, most of which is never
entrained in the current sheet. Instead, these neutrals exit the thruster at low velocity and
contribute very little to the total thrust. Furthermore, the thin layer of Teflon heated by
the arc continues to evaporate neutral gas for perhaps a millisecond after the discharge
ends [Spanjers et al. 1996]. For comparison, a typical PPT arc lasts for about 10 μs.
Spanjers et al. estimated that this so-called late-time ablation could account for up to 40%
of the mass ejected per shot. A third driver of the poor mass utilization efficiency is the
ejection of macroparticles. In another study, a high-speed camera was used to capture
images of these millimeter-sized macroparticles as they were ejected from the Teflon
surface at an estimated 300 m/s, too slow to contribute appreciable thrust [Spanjers et al
1998].
Another limiting factor in PPT performance is related to component lifetimes.
Specifically, high-voltage capacitors will ultimately fail after repeated cycling and/or
voltage reversals [Burton, et al. 1998]. This effect is more severe when capacitors are
cycled near their maximum rating, so this difficulty can be alleviated by taking a mass
penalty and choosing capacitors with a large voltage margin. Spark plugs are frequently
used to initiate PPT discharges in a controllable way and tend to limit the thruster lifetime
as material is eroded away. It should be noted that these lifetime issues are not
insurmountable, as the LES 8/9 PPT successfully fired 20 million pulses [Burton et al.
1998]. The extra mass and power associated with spark plugs makes them undesirable for
micropropulsion. A coaxial μPPT developed by the AFRL employs an alternative
ignition system with less limitation on lifetime [White, et al. 2002].
A typical PPT discharge may be modeled as an LRC circuit. Since the plasma resistance
is very low, this circuit is almost always underdamped, so the current oscillates for a few
periods before the capacitor is fully drained. This tendency to oscillate not only strains
the capacitor, it can also lead to a phenomenon known as restrike in which a second, less
energetic current sheet forms behind the first and ablates more material [Jahn 1968]. This
16
Table 1-2: Comparison of representative FEEP, colloid, and PPT micropropulsion technologies
ARC Indium FEEP [Tajmar et al. 2004]
Busek CMNT [Ziemer et al. 2006]
AFRL μPPT [Busek Co. 2007b]
Isp [s] 1600 - 8000 500-1500 830 Thrust [μN] 0.1 – 100 5-30 80
Min Impulse Bit [μN-s] 0.005 2.5 -
Max Voltage [kV] 10 10 3 Max Power [W] 13 ~ 1 2
Specific Thrust [μN/W] 7.7 20 41
tends to reduce the efficiency since the capacitor is mostly empty at the time of restrike
and because the new conductive path prevents further current from flowing in the
original, energetic current sheet. Diodes can be used to reduce the possibility of restrikes.
Table 1-2 shows some performance parameters for some of the more developed
microthrusters for comparison between state of the art FEEP, colloid, and μPPT systems.
Note that DC performance characteristics for the μPPT are calculated based on a pulse
frequency of 1 Hz. The listed maximum power refers to power input from the spacecraft.
A salient feature from Table 1-2 is the range of specific thrust, which illustrates the
different approaches of these technologies. By employing electrostatic acceleration, the
Indium FEEP can produce a high velocity, high power, low density exhaust stream, while
the μPPT achieves the opposite with a high density but overall slower exhaust stream.
This higher density approach leads to a specific thrust five times greater than the FEEP.
Due to the extremely high exhaust velocity of FEEPs and the need for heaters and
neutralizers, FEEP power consumption is much higher than the other microthrusters.
Also noteworthy is the extremely small impulse bit afforded by the FEEP system. This is
attributable to the passive feed system and rapid response time of the extractor electrode.
By comparison, the CMNT is pressure fed, using a microvalve with a relatively long
response time of about 0.5 seconds [Ziemer et al. 2006]. Finally, the significantly lower
maximum voltage requirement of the μPPT reduces PPU mass as compared to the FEEP
and colloid thruster.
17
1.4 Micro Pulsed Inductive Thruster
In order to improve on the performance offered by current micropropulsion systems, the
micro pulsed inductive thruster was developed by Eagle Harbor Technologies and tested
in vacuum at the University of Washington’s Advanced Propulsion Lab. The project
goals include micropropulsion system mass reduction and improved propellant utilization
efficiency.
1.4.1 μPIT Overview
μPIT is a small-scale coaxial PPT, similar to the AFRL μPPT, but with important
differences. A majority of the mass of the AFRL μPPT comes from the PPU and storage
capacitor. Two of the innovations of μPIT are an alternative ignition mechanism and a
solid-state RF power supply; both of these serve to greatly reduce PPU and capacitor
mass. Additionally, the power supply drives an antenna located near the Teflon face in
order to increase the propellant ionization fraction and to couple additional energy into
the plasma, increasing the average exhaust velocity. The potential exists for ionizing and
heating not only the neutral gas generated by the original arc, but also the neutrals
generated by late-time ablation, greatly improving propellant utilization efficiency.
As suggested in Figure 1-4, a typical PPT takes power from the spacecraft bus, usually at
28 V. A DC-DC converter is used to charge a capacitor to high voltage and then fire a
spark plug, initiating the discharge. Instead of a DC-DC convertor, the power supply for
μPIT drives a resonant LRC network by rapidly switching IGBTs (insulated gate bipolar
transistors), which draws energy from a storage capacitor. The inductor in this network is
itself the antenna used to couple energy into the plasma. The high voltage required for
breakdown is generated via a custom built step-up transformer. This oscillating signal is
rectified then passed to small igniter capacitors to initiate the discharge. In this way the
igniter is eliminated, saving mass and power, reducing complexity, and removing a
potential thruster lifetime limitation. Chapter 2 contains more details about the μPIT
PPU. After the initial arc, a plasma slug passes through the step-up transformer, at which
point the secondary winding presents a higher impedance than the plasma so that the
primary winding couples directly with the plasma in the style of an inductive discharge.
Figure 1-5 shows the electrode geometry and antenna configuration.
18
Figure 1-5: μPIT schematic. The winding L is dual purpose: it serves as the inductor in the tank
circuit, then couples directly to the plasma. Transformer secondary and ignition components are
omitted for clarity.
1.4.2 Gas-fed μPIT
The PPT work described here evolved from an earlier gas-fed microthruster that relied
entirely on an inductive discharge to generate thrust. The thruster is shown in cross-
section in Figure 1-6. Like the solid fuel μPIT described above, this thruster employed an
IGBT-driven resonant LRC network with the white windings in the figure functioning
both as the antenna and as the primary winding of a step-up transformer. The smaller
green windings in Figure 1-6 represent the transformer’s secondary coil, which is directly
connected at both ends to starter electrodes that protrude into the quartz gas feed tube.
Figure 1-6: CAD model of the gas-fed μPIT which uses an RF antenna (white coils) to generate an
inductively coupled plasma. Image reproduced from [Bartone 2007]
C
L Anode
Cathode
Teflon
RF heating stage
19
To generate thrust, Argon gas is injected into the quartz tube and a fire signal begins
switching the IBGTs. The LRC circuit was experimentally found to resonate at a
frequency of 813 kHz by varying the signal driving the IGBTs. As the resonant circuit
rings up, the transformer applies high voltage (estimated at a few kV) across the starter
electrodes; simultaneously, the propellant flows into the region near the starter electrodes
and breakdown occurs as the pressure crosses the threshold given by the Paschen curve.
Since there is very little capacitance in the igniter circuit, the starter electrodes only
generate a small quantity of seed plasma. However, this seed plasma provides enough
electrons for the antenna to couple effectively with the gas via axial and azimuthal
electric field components, further ionizing and heating the propellant. Chapter 4 contains
more details on inductive discharges.
Another design feature evident in Figure 1-6 is the turn spacing in the antenna. The turn
density decreases toward the thruster exit plane, which increases the radial magnetic field
component and reduces the axial component near the thruster exit as the field lines loop
back on themselves. This enhanced, time-varying 𝐵𝑟 is intended to interact with an
azimuithal current in the plasma in the manner of a conventional pulsed inductive thruster
(PIT). The time-varying 𝐵𝑧 generated by the antenna can drive an azimuthal mirror
current in the plasma, resulting in a 𝑗 × 𝐵 force that will accelerate the whole current
sheet.
Although diagnostic measurements did not determine conclusively whether the gas-fed
μPIT was functioning as an electromagnetic accelerator or as an electrothermal device,
the thruster clearly generated and exhausted plasma. Plasma density was measured on-
axis at a peak value of about 1016 𝑚−3 and sustained for several hundred microseconds
[Bartone 2007]. Comparison of antenna current waveforms with and without propellant
revealed a clear reduction in current amplitude when gas was present in the antenna. This
antenna loading corresponds to successful power transfer into the plasma. Based on these
results, the solid-fuel μPIT was developed.
20
1.5 Review of PPT Development
Pulsed plasma thrusters have a development history spanning more than four decades.
The earliest example is an electrothermal PPT that flew on the Soviet Zond-2 spacecraft
in 1964. The first Western PPT was designed and built at MIT’s Lincoln Lab and flew on
the LES-6 spacecraft in 1968 [Burton et al. 1998].
A 1972 study by Thomasson and Vondra [1972] using the LES 6 PPT was among the
first to investigate the exhaust constituents. Using a time of flight technique with
electrostatic probes, they measured an ion velocity of about 40 km/s. However, the mass-
averaged exhaust velocity was found to be just 3 km/s. This discrepancy prompted an
investigation into the exhaust components and their velocities using spectroscopy and a
Faraday cup. Faraday cup measurements showed that 90% of the total ablated mass was
neutral. The most important result was the suggestion that late-time ablation occurs as the
hot Teflon surface cools down on a much longer time scale than the current pulse, which
been verified directly by Spanjers [2002] and others. In a follow-up development in the
mid 1970s the Lincoln Lab designed and flight qualified the LES 8/9, which employs a
rectangular geometry and typical pulse energy of 20 J.
PPT development continued at a modest pace in the 1980s. Launched between 1982 and
1988, the TIP PPT flew on the US Navy’s NOVA navigation satellites and handled
stationkeeping maneuvers [Ebert et al. 1989]. A PPT also flew on the Japanese ETS-IV
mission for in-space testing and stationkeeping. This flight system was studied by Hirata
and Murakami [1982], who used a mass spectrometer for plume analysis to find not only
atomic carbon and fluorine, but also 𝐶𝐹2 and 𝐶𝐹3. They concluded that a significant
portion of the mass was ablated in molecular form, though recombination may have
played a role in these results.
PPT research accelerated in the following decades as interest grew for applications of the
technology to small spacecraft. The Burton group at Illinois developed several PPT
coaxial designs to investigate possible advantages over rectangular electrodes. A design
emphasizing electrothermal heating and acceleration produced significantly less late-time
ablation than the LES 8/9 PPT [Rysanek et al. 2000]. Temperature measurements on
similar PPT showed that up to 50% of input energy was lost to the walls as heat
21
[Bushman et al. 2001]. This level of heat loss may explain the reduced late-time ablation.
If the thruster body provides an effective heat sink, the Teflon can cool more rapidly,
thereby diminishing the evaporation rate after the initial arc.
Turchi et al. [2000] presented a potential technique for optimizing propellant utilization
by controlling the current pulse. The essential idea is to choose a current pulse that will
transfer energy to the solid propellant such that the rate of ablation matches the rate at
which mass can be accelerated away electromagnetically. Inductive energy storage is
suggested as a mechanism for lengthening the current pulse and a simplified model is
presented for performance analysis. Implicit in the analysis is the requirement for quasi-
steady operation instead of truly a pulsed device. While it may result in greater propellant
utilization efficiency, quasi-steady operation significantly increases the minimum
impulse bit a PPT can provide, reducing the capability for fine maneuvering and pointing
required for close spacecraft formations.
An experimental investigation on a modified LES 8/9 PPT by Spanjers, et al. [1996]
produced the first direct measurements of the neutral gas generated by late-time ablation.
However, because only a single laser was used for the measurement, the phase error
associated with external vibrations eventually could not be decoupled from the signal.
Since the vibration phase error increases with measurement time, late-time neutral
density could only be measured reliably to about 300 μs after the discharge. A further
consequence of the single-color interferometer was that the neutral density early in the
pulse could not be measured because of the much larger phase shift due to the plasma.
Despite this, late-time neutral density was found to be about five times the peak electron
density and this ratio was roughly independent of energy. This result is in keeping with
the theory of ablation controlled arcs [Ruchti et al. 1986], which predicts that the total
mass ablated scales with the energy dissipated in the plasma by the arc: 𝑚 ∝ 𝐸0 =
∫ 𝐼2𝑅𝑑𝑡. However, this linear scaling with energy was verified in the energy range from
2 – 40 J, while μPIT operates with discharge energies on the order of 100 mJ.
In another study using the same thruster, Spanjers, et al. [1998] used a high-speed camera
to image solid particles being ejected from the Teflon surface. The average velocity of
22
these macroparticles was estimated at 300 m/s using time of flight. The mechanism
proposed to explain these macroparticles is as follows: radiative heat transfer deposits
energy below the solid Teflon surface, unevenly vaporizing pockets of material. As the
pressure increases, the hot gas is eventually released, expelling solid Teflon with it.
Macroparticle ejection was estimated to account for 40% of the total shot mass, with
other mass utilization losses stemming from late-time ablation and low ionization fraction
in the initial arc plasma. The picture of macroparticle ejection described by Spanjers, et
al. suggests that a lower energy discharge and consequently cooler arc may reduce
macroparticle ejection.
In a similar study of the plume of the AFRL μPPT by Spanjers, et al. [2002], a two-color
interferometer was constructed to simultaneously measure time-resolved plasma and
neutral densities. A 13-pass Herriott cell was developed for the measurement by the
Burton group in order to obtain a greatly improved signal to noise ratio for the neutral
density measurement [Antonsen 2000]. While electron density was successfully
measured with a peak of 1021 𝑚−3, neutral density measurements failed. The Spanjers
study also used a Pockels cell for remote, high-bandwidth electrode voltage
measurements and a high speed camera. The camera once again captured macroparticle
ejection events, this time for the μPPT operating at 5 J.
A modeling effort by Keidar et al. [2004] at Michigan produced some new information
about the ablation process and about Teflon surface charring at low discharge energy, a
phenomenon discovered during lifetime testing of the AFRL μPPT. After firing twenty to
thirty thousand shots, a layer of black char formed Teflon surface, ultimately shorting the
electrodes and leading to thruster failure. It was found that char formation became more
severe at lower discharge energies. Above a critical discharge energy of about 6 J, no
charring occurred. Spectroscopy showed carbon and copper in the charred material,
suggesting back flux from the plasma (the electrodes were made of copper) as the
formation mechanism. A model by Keidar showed low surface temperature and
corresponding low ablation rates matching the regions where the char formed which
seems to reinforce the notion that the char is ablated carbon that has been deposited back
on the surface. Carbon is twice as common in Teflon as fluorine and its lower atomic
23
mass gives it greater mobility, enhancing the carbon flux to the surface. Higher discharge
energy is thought to prevent charring because higher current density increases the Lorentz
force so that the plasma density near the surface falls off more rapidly and carbon back
flux is reduced [Keidar et al. 2006].
In the most detailed experimental study to date of the role of neutrals in PPT discharges,
Koizumi, et al. [2007] presented emission spectroscopy and magnetic field probe data for
a standard macro-scale, rectangular PPT with a typical discharge energy of 10 J. A high
speed camera was used to capture 100 ns exposures showing details of the evolution of
the current sheet and neutral gas. The presence of neutral 𝐶2 in the plume was identified
by the Swan system in the 450-550 nm range, while the presence of 𝐶+was confirmed by
a principal emission line at 428 nm. Through the use of a bandpass filter, only neutral 𝐶2
atoms were imaged, and likewise only 𝐶+ ions were imaged. The superposition of these
two sets of photographs was compared with the broadband emission for the same frames.
The two were very similar, so it was concluded that 𝐶2 and 𝐶+ emissions were in general
representative of the neutral and ion dynamics, respectively. The high-speed images
showed the ions traveling in a current sheet at 10-20 km/s. The images also revealed a
slow-moving neutral layer that forms simultaneously with the current sheet and occupies
the space between the current sheet and the solid Teflon. The neutral layer drifts behind
the current sheet at about 2 km/s, so this neutral gas has no chance of being entrained and
accelerated by the sheet. As the current pulse reverses, a restrike is observed. The restrike
can potentially entrain some of the mass ablated and left behind by the initial pulse, but
the capacitor is mostly drained (24% charged) at the time of restrike, and this event
ablates still more propellant behind the second, less energetic current sheet. Koizumi, et
al. also note that the restrike occurs a small distance downstream of the solid Teflon
rather than at the surface. This is attributed to the axial neutral density gradient which
leads to Paschen minimum in a lower density region slightly downstream of the Teflon.
As a result, the neutral density near the Teflon surface could be estimated to be at least
0.5 Torr.
The most recent PPT hardware tested in orbit is the EO-1 PPT, which flew in 2001. This
thruster can provide an impulse of 90-860 μN-s at an Isp of 650-1400 seconds and an
24
overall efficiency of 8%. Processing 9-56 W at one Hz, the EO-1 PPT provided attitude
control for the spacecraft without negatively impacting the spacecraft bus or onboard
science instruments [Zakrzwski et al. 2002]. The primary improvement over the older-
generation LES 8/9 PPT was reducing the electronics mass to approximately 5 kg,
including two PPTs, common electronics, and mounting hardware [Zakrzwski et al.
2001]. The characteristics of the EO-1 PPT plume were studied in detail by Gatsonis et
al. [2004], who reported 10 eV electrons early in the pulse and decreasing electron
temperature as the density increased. Presumably these electrons lose energy as they
undergo ionizing collisions.
The Dawgstar PPT is a smaller-scale PPT based on the EO-1 design. Operating at 5 J per
pulse and a specific impulse of about 480 s, the Dawgstar PPT was intended to provide
primary propulsion for the Dawgstar university nanosatellite [Rayburn, et al. 2005].
Other characteristics include a 56 μN-s impulse bit and a total system mass of about 4 kg,
which included the PPU and eight thruster modules for three-axis control. Although the
spacecraft was never launched, a primary objective was the demonstration of formation
maneuvering as a part of the ION-F cluster of university nanosatellites.
Another microPPT development effort on an even smaller scale was reported by Simon
and Land [2004]. A prototype 2-D microfabricated PPT with Teflon tape for fuel was
tested over a range of input energy. Propellant charring occurred at low energy levels,
similar to the AFRL μPPT results reported by Keidar. At higher energies, charring did
not occur over the course of lifetime tests but a greater level of late-time ablation is likely
due to increased Teflon surface temperature. Specific performance details regarding
impulse bit and Isp were not reported.
In addition to this solid-fuel microPPT, the investigators are developing a liquid-fueled
variant in order to avoid propellant surface charring. This prototype thruster uses liquid
water propellant which vaporizes and diffuses through a permeable membrane. The use
of liquid water propellant poses some difficulty with propellant loss in a vacuum
environment; zero-power water flow rates of 10-20 μg/s were reported. Since the entire
device holds slightly less than 1 mg of propellant, this leakage represents a serious
problem. Recent thrust stand measurements were performed using a novel vibrating
25
thrust stand technique which showed a minimum impulse bit of about 0.2 μN-s [Emhoff
2007].
1.6 Outline and Research Goals
Primary goals of this project are to demonstrate reliable operation of μPIT and to explore
the operating regimes of the thruster. PPTs almost always use a spark plug igniter to
control the discharge. The elimination of this component, while increasing lifetime and
saving power, also presents reliability challenges since high voltage is required to achieve
breakdown in a repeatable way. In addition, the ignition mechanism selected for μPIT is
susceptible to multiple breakdowns as the antenna rings down over 100-200 μs. The
elimination of the additional breakdown events is important for reliable, repeatable
operation. An additional goal is knowledge of the thruster performance parameters such
as impulse bit and thrust efficiency. These quantities are estimated based on surveys of
the plume. Finally, the impact of the antenna and inductive coupling on overall
performance will be evaluated.
Chapter 2 describes the vacuum system, data acquisition, and other infrastructure
necessary for the experiment. Chapter 2 also contains details about the resonant, IGBT-
driven power supply and ignition system developed specifically for μPIT. Chapter 3
describes the diagnostics implemented for μPIT, including Rogowski coils, high voltage
dividers, a fast-framing CCD camera, and two styles of Langmuir probes. Theory for
inductive discharges and data reduction techniques are presented in Chapter 4. Chapter 5
gives an account of three design iterations and the performance results for each. Also
described is the comparison in thruster performance between PPT mode (without antenna
coupling) and PIT mode (with coupling). Finally, Chapter 6 contains a summary the
major contributions of this project and recommendations for the directions of future
work.
26
2 Experimental Setup and Thruster Design
2.1 Vacuum System
The Advanced Propulsion Lab’s bell jar facility visible in Figure 2-1 was used for this
series of experiments. A mechanical roughing pump drives the pressure down to a few
hundred millitorr in about 15 minutes, then a turbopump takes over, drawing the chamber
down to a base pressure of about 6 × 10−7 torr. However, most of the data presented was
collected in the range 1 × 10−6 to 3 × 10−6 torr. The Varian TV 550 Navigator
turbopump has a pumping rate of 550 L/s for nitrogen at a maximum speed of 42 krpm.
The roughing pump is attached in series with the turbopump so that the turbopump
exhaust flows directly to the roughing pump inlet.
Figure 2-1: Bell jar vacuum chamber enclosed by a plexiglass shield
Glass bell jar
Mounting flange
Power feed-through
Vent valve
Fire control feed-through
Langmuir probes
μPIT
Rogowski and voltage divider signals
27
A Pirani gauge measures pressure in the line between the turbopump output and the
roughing pump inlet. The Pirani senses pressure by passing a constant current through a
filament; the filament temperature increases as the gas pressure drops and a thermocouple
senses the temperature directly. The Pirani gauge can accurately measure pressures
between 1 millitorr and 20 torr. Another Pirani measures low vacuum in the bell jar itself,
while a cold cathode gauge measures high vacuum in the bell jar. The Televac 7E cold
cathode gauge can measure pressures between 1 × 10−8 and 1 × 10−2 torr. The cold
cathode gauge is a type of ionization gauge that measures pressure by bombarding a gas
with electrons and collecting the resulting ion current. With a calibration, this ion current
can be related to the gas pressure. Pressure measurements are displayed by the pressure
readout in Figure 2-2 .
The bell jar measures 46 cm in diameter by about 75 cm high, which provides ample
room to study the evolution of the exhaust plume. The stainless steel mounting flange
adds about 23 cm in height and provides several access ports. Eight one inch radial ports
visible in Figure 2-1 accommodate power leads, input/output data connections, and a
venting valve. Not shown are four more ports located at the base of the mounting flange,
below the table. One of these serves as the inlet for the turbopump while another provides
access for the vacuum gauges. The other two provide Langmuir probe access.
2.2 Control and Data Acquisition
Figure 2-2 shows the bell jar and the peripheral equipment necessary to operate and
gather data from μPIT. Custom Labview software is used to control the optical pulse
generator, which sends optical pulses to equipment such as the function generator and the
camera. However, these devices require electrical digital logic signals, so the optical
pulses are translated to 5 volt TTL (transistor-transistor logic) signals with nearby
custom-made receivers. Pulse width and timing are adjustable from the Labview
software. Optical communication is used to limit the effects of electrical noise on the
control system.
28
Figure 2-2: Data acquisition arrangement
To fire μPIT, the PPU’s onboard storage capacitor is charged, typically to 350 V. Then
the Labview control software is used to trigger a Berkeley Nucleonics function generator,
which in turn produces a TTL pulse train of predetermined pulse width and frequency.
This signal is translated by a custom transmitter box into an optical pulse train and sent to
a similar optical receiver, which passes the signal to the IGBTs to drive the resonant LRC
circuit. The pulse width, frequency, and the length of the pulse train are all set manually
on the function generator prior to a shot. More details about function generator tuning
and PPU operation are presented later in this chapter.
Roughing pump
μPIT
Turbopump
Cold cathode gauge
DC power supply
Optical pulse generator
Digital oscilloscope
Function generator
Control computer
Pressure readout
29
Figure 2-3: Langmuir probe range of motion. The symmetric double probe can only sample the
plume where it exits the RF heating stage, while the planar double probe can sample the plume at
any (r,z) location.
The Langmuir probes visible in Figure 2-1 have different degrees of freedom, as
illustrated in Figure 2-3. A symmetric double probe is inserted radially and dog-legged
so that it can sample the plume approximately at the exit plane, but nowhere else. This
probe can measure the peak on-axis density and is also important for time-of flight
measurements. The asymmetric double Langmuir, or planar probe, enters the bell jar
from below so that it can both translate vertically and rotate azimuthally. In addition to its
role in time of flight measurements, the planar probe is used to map out the axial and
radial density profiles. The probe positions are adjusted manually and displacements are
measured with a precise ruler. Before pumping down to vacuum, the initial positions of
the probes are measured and recorded. Experience has shown that this technique is
accurate to 5 mm or less even after many adjustments of the probe positions.
Radial positions are more difficult to measure, so a technique was developed to find the
change in the angle as the planar probe is rotated. Two strips of 1/8‖ aluminum were
fixed horizontally to the feed-through hardware below the bell jar where the planar probe
enters the chamber, shown schematically in Figure 2-4. One strip was fixed vertically to
μPIT
Planar probe
Symmetric double probe
𝑟 𝑧
30
the knurled steel knob that tightens the O-ring, while the other strip was fixed to the ¼‖
steel tube shielding of the probe itself. This way, when the probe is rotated, the angle θ
between the aluminum strips changes. Because of the awkward position of this
arrangement and because protractors are poorly suited to measuring angles in three
dimensional structures like this, a more accurate technique was developed for finding θ.
To find this change in angle, an isosceles triangle is formed by marking a distance on
each strip from the point where the strip is tangent to its associated cylinder (steel knob or
probe tube). Then calipers are used to measure the distance d accurately to within 0.1
mm. Taking the finite radii of the steel knob and the probe shielding into account, an
implicit equation for θ in terms of d can be derived:
1/22 2
1 2 1cos sin sin cos2 2 2
t t td a a r r a r
(2.1)
In Equation (2.1), 𝑡 is the thickness of the aluminum strips and 𝑟1 and 𝑟2are the radii of
the steel knob and probe shield, respectively. This relationship holds for 𝜃 < 90° and
turns out to be fairly linear up to around 50°. By measuring the change in angle each time
the probe is rotated, the change in θ may be computed, which is then related in a simple
way to the radial position of the probe tip.
Figure 2-4: View along the z-axis of the technique for measuring radial position. The distance “a”
was selected to be relatively long in order to provide good angular resolution. This setup can
accurately resolve angular position to 1-2 degrees.
Steel knob
Planar probe shielding
Moving aluminum strip
Fixed aluminum strip
θ
d
a
a
31
Data acquisition is synchronized by the Labview control software, which sends pulses to
each oscilloscope at the same time that it sends the fire command to μPIT. The Tektronix
TDS 3034 four-channel digital oscilloscopes have a 300 MHz bandwidth and can collect
2.5 GS/s, with a maximum of 10,000 datapoints per waveform. Two Langmuir probe
traces, the antenna Rogowski coil signal, the function generator pulse train, and up to two
voltage divider signals are collected for each shot. Another custom Labview program
acquires the scope data via Ethernet connections and saves them as text files.
2.3 μPIT Power Processing Unit
Important features of the μPIT PPU are the low mass, a result of the solid-state switches
and the resonant LRC network, and the ignition system. Instead of a spark plug, μPIT
uses a step-up transformer to develop a high-voltage breakdown across the Teflon. A
typical PPT power supply uses a DC-DC power supply to charge a large capacitor to a
few kV, resulting in a pulse of tens of Joules. Capacitor mass generally dominates power
supply mass and can account for 30% or more of the mass of the entire thruster package
[Burton et al. 1998]. Even small PPT designs like the AFRL μPPT and the Dawgstar PPT
consume 2-5 J per shot, while the μPIT igniter capacitors store just 100 mJ. Because the
initial discharge is small and additional energy is coupled into the flow by the antenna, a
similar amount of energy can be deposited in the plasma with much smaller capacitors.
This results in significant mass savings over traditional PPTs. Because of the RF power
coupling, μPIT can potentially combine lower system mass with greater propellant
utilization efficiency.
This style of solid-state RF power supply has been previously implemented for the High
Power Helicon experiments at the Advanced Propulsion Lab [Ziemba et al. 2006]. With
much less energy storage and a mass below 200 grams, the μPIT PPU is a significantly
scaled-down version of the HPH power supply.
32
Figure 2-5: Spice model of the μPIT power supply, divided by transformers into three sections:
energy storage and switches, resonant LRC circuit, and igniter components. The entire PPU operates
in vacuum.
2.3.1 Circuit Description
The μPIT PPU can be divided into three sections separated by transformers, as shown in
Figure 2-5. Section 1 includes the energy storage capacitor. This electrolytic capacitor
can charge to 350 V, storing a maximum of about 6 J. Despite initial concerns about
outgassing, this capacitor has performed well in vacuum. Also incorporated in section 1
is the DC power supply used to charge the storage capacitor prior to shot. A 500 V Fluke
power supply was used for this purpose. The DC supply was located external to the bell
jar and energy was transferred to the storage capacitor via a high voltage feed-through
shown in Figure 2-2. Although during testing the power supply remains connected to the
storage capacitor, the charging time scale is much greater than the characteristic shot time
due to limited DC power supply current. As a result, voltage divider measurements of the
storage capacitor charge are accurate over the course of a 150 μs shot.
Also included in section 1 is a parallel array of IGBTs (insulated gate bipolar transistor).
The particular IGBTs selected for the μPIT PPU are rated for up to 1200 V and 240 A
pulsed, but experience has shown that these switches can handle as much as 400 A
pulsed. No decrease in performance is apparent after around 1000 shots. The IGBTs can
respond as fast as 1-2 MHz, but for μPIT they are typically switched at about 300 kHz.
Because the IGBTs require a gate current on the order of one amp, an IXYS driver chip is
used to amplify the pulse train transmitted from the function generator. This amplified
Pulser signal input
1) Energy Storage and IGBTs 2) Tank Circuit 3) Igniter
To electrodes
33
pulse train closes and opens the IGBTs at a frequency selected to resonantly drive the
tank circuit in section 2, which is coupled to section 1 by a 1:1 transformer.
Another important element in section 1 is a high voltage snubber circuit. Due to
inductance at the 1:1 transformer and elsewhere in the system, large voltage spikes
appear across the IGBT terminals when the switches open or close, generating a large 𝑑𝐼
𝑑𝑡.
This is particularly problematic when the IGBTs have just opened because in that case, a
back EMF develops across the inductive elements that add constructively with the storage
capacitor voltage, resulting in a large voltage spike until the current settles down. The
opposite case, just after the switches close, is less severe because the back EMF has the
opposite polarity, so it adds destructively with the capacitor voltage and produces a
smaller voltage spike. The snubber protects the IGBTs by absorbing this brief voltage
spike.
The tank circuit in the middle section of Figure 2-5 is coupled to the IGBT and storage
capacitor by a transformer. The secondary side of this transformer serves as the driving
voltage for a high-Q resonant LRC circuit. Current and voltage measurements show a
very slow decay, indicating little dissipation within the tank circuit. The tank circuit
tuning capacitor is a Cornell Doublier model rated for 2000 VDC. The inductor functions
both as the antenna and as the primary side of the transformer. This way, this inductor
plays two key roles in the operation of μPIT: it helps generate high voltage in order to
produce the initial arc and it is responsible for adding energy directly to the propellant
after the arc occurs.
The oscillating high voltage generated by the transformer secondary is not useful for
charging the igniter capacitors, so a full-bridge rectifier in section 3 is used to change it to
a roughly DC signal. The diodes used in the rectifying bridge are rated for 15 kV and
have a 100 ns recovery time. The fast recovery time is particularly important for this
application because the transformer current reverses approximately every 1.6 μs. During
the development process, three more rectifying bridges were added in parallel with the
first in an attempt to increase the current to the igniter capacitors. Also incorporated into
section 3 are the igniter capacitors. Four 1 nF 15 kV ceramic capacitors are connected in
34
parallel across the electrodes to store energy for the initial arc. As the tank circuit rings
up, the rectifier charges up the igniter capacitors until enough voltage is built up for
breakdown to occur. At a pressure of about 1 × 10−6 torr and an electrode gap of about
0.8 mm, this occurs at around 4 kV. This ignition system eliminates the need for a spark
plug, avoiding a possible lifetime limitation.
2.3.2 Bench Testing and Tuning Procedure
Prior to firing in vacuum, μPIT was tested at low power in air to measure circuit
performance and to calibrate the onboard Rogowski coil and voltage dividers. In addition,
the driving frequency was varied to find the operating parameters that produce the best
resonance in the tank circuit. A Berkeley Nucleonics pulse generator was used to produce
a series of square waves to switch the IGBTs. The two parameters available to control the
PPU tuning are pulse period and pulse width (i.e. duty cycle). In general, a 50% duty
cycle did not produce the best performance. Because these two parameters are
independently adjustable, many combinations were possible, so a methodical approach
was use to optimize the driver pulses. A starting frequency was estimated based on the
resonant frequency for an LRC circuit.
1
LC (2.2)
The antenna inductance was estimated using the long solenoid approximation.
2
0N AL
l
(2.3)
𝑁 is the number of turns, 𝐴 is the cross-sectional area, and 𝑙 is the solenoid length.
Keeping this frequency fixed, the tank circuit response was monitored as the pulse width
was varied. This procedure was repeated for several frequencies above and below the
starting frequency in order to map out performance over the nearby parameter space.
Three parameters were used to identify the best operating point: antenna voltage
amplitude, antenna current amplitude, and the degree to which the waveforms had a
smooth, sinusoidal form. A tradeoff among all three of these was necessary in order to
develop high voltage for the igniter and also couple effectively with the plasma.
35
After optimum the pulse frequency and duty cycle were selected, the full system was
tested in air at low storage capacitor charge and the circuit parameters were measured. A
Stangenes probe was used to measure antenna current and to calibrate a Rogowski coil.
The antenna voltage was measured with a Tektronix differential high voltage probe,
while the step-up transformer output voltage was measured using a second Tektronix
high voltage probe. The peak current and voltages were recorded and a linear
extrapolation was performed to estimate system performance at the maximum storage
capacitor charge of 350 V. Direct measurements at high charge voltage were not possible
due to air arcing. Bench test results for three step-up transformer configurations are
presented in Chapter 5.
2.4 μPIT Mechanical Design
μPIT is a coaxial PPT, as illustrated in Figure 1-5. This geometry was chosen over
rectangular electrodes primarily for its ease of integration with the RF antenna. Spores
and Birkan [1999] pointed out that the coaxial design may help confine EMI and reduce
plume interaction with the spacecraft. Of course, the μPIT antenna introduces a new
source of noise.
Figure 2-6: Pictures of μPIT with and without igniter capacitors and associated Kapton insulation
Position of Teflon face
Teflon propellant
Rectifying bridge
Torr seal insulation
Transformer/antenna
Anode
Cathode
Igniter capacitors
36
Figure 2-6 shows some images of μPIT. The Teflon propellant face is located
approximately at the lower edge of the antenna. This positioning represents a
compromise between operation before and after the initial arc. A gap between the Teflon
face and the beginning of the antenna would reduce the magnetic field near the propellant
surface which facilitates the breakdown process. On the other hand, placing the
electrodes further up so that they overlap with the antenna would reduce direct coupling
with the plasma and increase energy losses as the anode screens a portion of the induced
fields.
Copper tubing was chosen as the electrode material because it allows for soldering of the
igniter components directly to the thruster body. Figure 2-6 shows the igniter capacitors
and rectifying bridge soldered directly to the electrodes. The close spacing helps reduce
parasitic inductance, concentrating most of the available voltage across the electrodes
instead of elsewhere in the system. Since copper is susceptible to erosion over the long
term, a refractory metal such as molybdenum or tungsten would likely be substituted for
the copper in flight-ready thruster. Dimensions of the electrodes and transformer are
listed in Table 2-1.
The inner electrode was selected to serve as the cathode because of the field-enhancing
effect of the smaller electrode. The enhanced field is more useful near the cathode than
the anode because it aids in the field emission of electrons, promoting breakdown a lower
applied voltage. As a further field enhancement technique, the cathode was sharpened to
a point angled toward the Teflon. This is technique illustrated in Figure 2-7.
Table 2-1: Selected dimensions of the microthruster assembly
Anode ID [mm] 6.4 Anode OD [mm] 8.0 Cathode OD [mm] 4.8 Teflon tube wall thickness [mm] 0.8
Teflon surface area [𝐦𝐦𝟐] 56
Quartz tube OD [mm] 12 Quartz tube wall thickness [mm] 1 Transformer length [mm] 25-30
37
Figure 2-7: Electrode configuration showing the sharpened cathode
In order to prevent alignment errors, the cathode is press-fit into a hole in an adjustable
nylon block. This block was milled to ensure that the mounting hole is perpendicular to
the flat upper surface. A level was used to position the block so that the electrodes and
thruster plume are oriented very close to the vertical.
Three versions of the step-up transformer were constructed, with turn ratios ranging from
8:120 to 8:255. An 8:185 transformer is shown in Figure 2-6. The transformers were
wound on Quartz tube about 1 cm in diameter. Quartz was selected because it tends not
to react with or contaminate the plasma. The primary coil was wound with Kapton-
insulated 30 gauge wire with a DC voltage rating of 4 kV. The primary coil was wound in
an eight-turn helix with about 3 mm spacing between turns. The secondary coil was
wound on top of the primary with Kapton insulation in between. Depending on the
number of turns, the secondary coil consisted of 2 or 3 layers of tightly spaced 32 gauge
magnet wire insulated with formvar. The DC voltage rating for the secondary wire was
about 1.8 kV and the maximum voltage drop expected across each turn was about 75V.
As a result, adjacent turns did not require insulation. However, the transformers
generated an estimated 5 – 8 kV at peak antenna voltage, so each layer of windings was
insulated from the other with a combination of Kapton tape and Torr seal. As is evident
in Figure 2-6, epoxy was applied to the transformer leads and igniter components to
insulate against arcs elsewhere in the system.
Cathode
Teflon Anode
CL
38
3 Diagnostics
Several instruments were used to help characterize μPIT’s electrical systems and exhaust
plasma in order to estimate some key performance parameters. Taken together, these
instruments allowed for a greater understanding of the physics of the device and guided
changes and improvements. This chapter discusses the theory underlying Rogowski coils,
voltage dividers, and Langmuir probes and also outlines the uses of a fast-framing CCD
camera.
3.1 Rogowski Coil and Stangenes Probe
A commonly used current-measuring device is the Rogowski coil, shown schematically
in Figure 3-1. This magnetic diagnostic senses any time-varying current that threads
through it. Threaded current produces an azimuthal magnetic field which varies in time
according to the current. This time-varying flux generates a voltage signal in the
Rogowski’s toroidal coils. Since the return path for the coil is straight through the center
of the torus, this leg of the device picks up no induced voltage and the output voltage
signal is proportional to the time derivative of the current.
Figure 3-1: Rogowski coil schematic. Image reproduced from [Hutchinson 1987]
39
The Rogowski can by treated analytically as long as there is very little variation in the
magnetic field on the scale of the turn spacing, or
B
mB
(3.1)
where 𝑚 is the number of turns per length. Also, let the turn spacing be uniform with a
cross section 𝐴, as illustrated in Figure 3-1. Given these restrictions, the flux linkage over
the loop can be written
l A
m B dldA
(3.2)
The geometry of Equation (3.2) is shown in Figure 3-2.
Now, using Ampere’s law
0 encl
l
B dl I
(3.3)
Equation (3.2) can be rewritten as
0 encl
A
m I dA (3.4)
Evaluating the integral and applying Faraday’s law yields
0V m AI (3.5)
Figure 3-2: Equivalent geometry for an idealized Rogowski coil Image reproduced from [Hutchinson
1987]
40
This output signal 𝑉 can then be integrated to provide a time-resolved current
measurement. Although this technique is accurate, it can be difficult in practice. Noise in
the system can cause problems with the integration. A voltage offset in the digitizer, for
example, will cause the oscillating current signal to lift away from zero because of the
integration. A simple alternative is to calibrate the Rogowski output using a Stangenes
probe. However, this method is specific to sinusoidal current signals like the ones in the
μPIT antenna.
The Stangenes is used to measure the true current signal during bench testing over a
range of storage capacitor charge voltages. The resulting peak current values are recorded
and correlated with the Rogowski output. Generally a highly linear fit is obtained. Even
though the Rogowski signal is not integrated, the oscillatory nature of the signal means
that the integral is about the same shape, just scaled by the frequency. Since unknown
phase errors exist in both the Stangenes and Rogowski signals, this unintegrated but
calibrated waveform contains approximately the same information as an integrated
Rogowski signal would. The Stangenes is itself a Rogowski coil, but it is the self-
integrating variety which allows for direct measurement of the current. In order to self-
integrate, the Stangenes must have an L/R time much less than the characteristic signal
time (one oscillation period). This way, the induced voltage in the Rogowski coil rapidly
drops off and the coil begins flowing current to oppose the applied magnetic field.
Conversely, an ordinary Rogowski coil as described by Equation (3.5) should have
𝐿/𝑅 ≫ 1/𝜔 so that induced voltage can be sustained for a longer time than it takes to
reverse the measured current.
3.2 Voltage Dividers
Voltage dividers were constructed to monitor the storage capacitor and the voltage
applied across the electrodes over the course of each shot. Voltage dividers are attractive
because they are simple and inexpensive to implement. They also provide a natural
discharge path for capacitors that could otherwise hold a potentially dangerous charge for
an extended period of time. However, they also come with some disadvantages. Voltage
dividers require a direct connection to the circuit, applying a high impedance load. In the
case of high voltage measurements like the μPIT electrodes an arc or a failure can lead to
41
Figure 3-3: Voltage divider circuit diagram. Image reproduced from web.mit.edu
high voltage appearing across the sensor signal leads, endangering the digitizer. Voltage
dividers are also highly susceptible to noise. An example voltage divider circuit is
presented in Figure 3-3.
This circuit produces an output voltage given by
2
1 2
out in
RV V
R R
(3.6)
Typically the resistances are chosen such that 𝑅1 ≫ 𝑅2 in order to divide 𝑉𝑖𝑛 down to a
manageable signal. This expression is valid as long as nearly all the current flowing
through 𝑅1 also flows through 𝑅2, which can be enforced by choosing an input
impedance to measure 𝑉𝑜𝑢𝑡 that is much larger than 𝑅2. While this equation is accurate,
large resistances can be difficult to measure with precision. For example, a 25 MΩ
resistor was used to measure the voltage across the electrodes. For this reason, the voltage
dividers constructed for μPIT were calibrated using a high voltage probe.
A voltage divider as was constructed with ¼ Watt carbon film resistors to measure the
charge on the storage capacitor. 𝑅1 was set at 2 MΩ and a proportionality constant of
about 100 was selected. Calculations verified that the peak power dissipation was well
below ¼ W and that the RC time was much greater than the characteristic μPIT shot time.
A voltage divider was also constructed to measure up to 10 kV across the electrodes. A
25 MΩ high-voltage resistor was chosen for 𝑅1 and a proportionality constant of 1000
was selected. Peak power dissipation was well below the rated maximum and the RC
42
time was about 0.1 second. Since a shot lasts for about 150 μs, this high voltage divider
did not affect the operation of the ignition system.
3.3 Langmuir Probes
Langmuir probes were used extensively to measure bulk flow velocity via time-of-flight
and also to measure plasma density profiles. In its simplest form, the Langmuir probe is a
small piece of conductive material inserted into a plasma that draws a current that
depends on the local plasma properties and the applied voltage. The resulting I-V
characteristic contains information about the plasma density, electron temperature, and
floating potential. However, single probes require a conducting wall that is in good
contact with the plasma in order to form a complete circuit. The glass bell jar used to test
μPIT cannot meet this requirement in general. An alternative is the double Langmuir
probe, which is electrically floating with respect to the chamber wall but maintains a
potential difference between two probe tips. First developed by Johnson and Malter
[1950], the double Langmuir probe can provide local density and electron temperature
measurements. Double Langmuir probes were used in this study to map the density
profile in the μPIT plume and to measure exhaust velocity using a time-of-flight
technique. This section outlines the standard Bohm theory for a collisionless, quiescent
plasma and the theory for crossed probes in a flowing plasma. Then a combination of
these is presented for interpreting current measurements from asymmetric double
Langmuir probes in crossflow.
3.3.1 Standard Double Probe Theory
A double Langmuir probe consists of two probe tips immersed in the plasma with a
voltage applied between them. The probe is floating so that it draws no net current.
Figure 3-4 schematically shows the relevant potentials for double probe operation. The
Bohm theory relies fundamentally on the formation of a sheath around the probe tips. A
sheath naturally forms around any object inserted into the plasma because the highly
mobile electrons carry more current to the probe than the ions. If electrically isolated, the
probe will quickly charge to a potential such that the flux of ions and electrons results in
zero net current. The voltage the probe acquires in this process is the floating potential
(𝜙𝑓), and is necessarily more negative than the plasma potential, or space potential (𝜙𝑠).
43
Figure 3-4: Double probe potential diagram. Image reproduced from [Byrne 2002]
As long as the probe potential is less than the plasma potential, ions are accelerated and
electrons are retarded in a sheath region a few Debye lengths across. As the figure
indicates, the application of the bias voltage𝜙𝑏 tends to drive one probe above the
floating potential and the other probe below. This asymmetry combined with the
requirement for zero net current causes 𝑃1 to collect a net electron current while 𝑃2
collects a net ion current.
Several assumptions are necessary to develop Bohm theory, the most commonly used
theory for interpreting double Langmuir probe I-V characteristics. First, a Maxwellian
velocity distribution is assumed. Next, the plasma is assumed to be collisionless on the
length scale of the probe radius
, , , ,p ei ee en ii inr (3.7)
and on the scale of the sheath thickness
, , , ,s ei ee en ii ind (3.8)
Here 𝜆 is the mean free path and the subscripts refer to collisions among ions, electrons,
and neutrals. 𝑑𝑠 is the sheath thickness, given approximately by
44
1/2
022
3s D
e
eVd
kT
(3.9)
In Equation (3.9), 𝜆𝐷 is the Debye length
2
0 /
/ /D
e i
k e
n T n T
(3.10)
and 𝑉0 is the potential difference between the probe tip and the plasma. Probe tips should
be arranged with enough spacing to prevent sheath interactions. For a probe spacing 𝑠
this requirement can be expressed as
1s
s
d (3.11)
It is also assumed in the following analysis that gradients in the collection region are
negligible, that no magnetic fields are present, and that the plasma is stationary. This
way, the probes only collect random current due to thermal motion. The current to each
probe tip (𝐼𝑝) can be expressed in terms of the combined ion and electron current. The
convention used here is to treat collected electron current as positive and collected ion
current as negative.
p e iI I I (3.12)
The electron current will be addressed first. The electron saturation current density can be
expressed in terms of the thermal speed 𝑣𝑡 .
0e e teJ en v (3.13)
The thermal speed for either species is given by
( )
( )
( )
e i
t e i
e i
kTv
m (3.14)
Equation (3.13) gives the electron saturation current density, which is the current density
one could expect if the sheath voltage did not interfere with electron collection. A
45
Maxwellian distribution is assumed in order to account for the fact that only the most
energetic electrons are collected by the probe; the rest are reflected. Consequently, the
electron density in the sheath is reduced by the Boltzmann factor
expes
e
eVn n
kT
(3.15)
The electron current to the probe with sheath area 𝐴𝑠 can then be expressed using this
Boltzmann factor
0 exps p
e s e
e
eI A J
kT
(3.16)
Unlike the electron, the ions are assumed to be cold with an energy approaching zero at
infinity. The ions gain energy as they fall through the sheath. Bohm [1949] showed that
the formation of a sheath with a sheath voltage 𝑉𝑠requires that
2
es
kTV
e (3.17)
in order that the ion flux matches up at the interface between the plasma and the sheath.
Quasineutrality in the region just outside the sheath requires that the ion density match
the electron density given in Equation (3.15). Additionally, the ion velocity entering the
sheath can be approximated by
1/2
2 sis
i
eVv
m
(3.18)
Using Equations (3.15) and (3.18), an expression for the ion current density can be
written
1/2
2exp s s
i s is is s
e i
eV eVI A n v A n
kT m
(3.19)
Equation (3.19) can then be combined with (3.17) to write the Bohm formula
46
1
exp2
i s BI A n v
(3.20)
Where the Bohm velocity is given by
1/2
eB
i
kTv
m
(3.21)
Substituting Equations (3.15) and (3.20) into (3.12) and assuming that the sheath area
equals the probe surface area (𝐴𝑠 = 𝐴𝑝) yields the current collected by a single probe tip.
1/2
0 exp 0.61s p e
p p e p
e i
e kTI A J A n
kT m
(3.22)
The measured current 𝐼 shown in Figure 3-4 is the difference between the currents
collected by the probe tips
1 2I I I (3.23)
This may be combined with the fact that the probe draws no net current
1 2I I (3.24)
With these relationships between 𝐼1 and 𝐼2 and Equation (3.22), a probe I-V characteristic
can be written in terms of the ion saturation current and the probe bias voltage.
tanh2
bi
e
eI I
kT
(3.25)
Note that the two probe collection areas were assumed to be equal in this analysis and
that the sheath areas were assumed to match the probe areas. This latter approximation is
most accurate in the thin-sheath limit, in which
100p
D
r
(3.26)
47
Equation (3.25) is plotted in Figure 3-5. A noteworthy feature is the way the curve
flattens out for large bias voltages. The theory predicts that the current will be
independent of the probe potential, a result of the thin sheath approximation and the
assumption that the ions are cold. Increasing the probe potential while already in
saturation tends to increase the sheath thickness, resulting in a larger collection area and a
slowly increasing saturation current. The assumption that the probe areas are equal leads
to a symmetric I-V characteristic with positive and negative ion saturation values that are
equal in magnitude. Usually double Langmuir probes must be swept through several bias
voltages to map out the curve in Figure 3-5. However, the probe can be operated in one
of these saturation regions and the density can be calculated using an estimated electron
temperature. Because of the short time scales involved and the difficulty of high
frequency voltage sweeping, this saturation approach was used to measure density in the
μPIT plume.
The standard method for extending these results to a flowing plasma is to align the probe
tip axes parallel to the flow so that the probe only collect a random rather than convected
ion current. However, in this study the double probes were aligned perpendicular to the
flow. Mach probe theory was used to adjust the ion current projected by this standard
theory, as described in the following section.
Figure 3-5: Theoretical I-V characteristic for a symmetric double Langmuir probe
4 2 2 4Vb
1.0
0.5
0.5
1.0
48
3.3.2 Crossed-Probe Theory
Several authors have shown that Langmuir probes oriented transverse to a flowing
plasma collect more current [Tilley et al. 1990 and Chung et al. 1975]. The electron
current contribution can be assumed to be approximately the same as the Bohm theory
case as long as the bulk velocity 𝑈 does not significantly shift the electron energy
distribution. This condition can be stated as
1t ev
U (3.27)
If this is true, the collected electron current can still be described using the Boltzmann
factor. The difference, then, must be due to an increased ion current.
The ion saturation current for probes in transverse flow is larger because ions are
transported to the probe surface through a combination of convection and diffusion.
Conversely, in the stationary case, probes collect ion current through diffusion only. To
account for this effect of transverse flow, a current collection model was developed by
Kanal [1964] for use with sounding rocket data collected by probes flying through the
upper atmosphere. An important approximation in Kanal’s theory was the choice of a
cylindrical sheath around the probe tip. In general, an asymmetric sheath around a
cylindrical probe tip would be expected for a probe in crossflow and the particular shape
of the sheath is unknown. Consequently, this approximation introduces an error into the
current collection calculation. Despite this, Kanal’s theory has been successfully used to
interpret the results of crossed probes, or Mach probes. Johnson and Murphree [1969]
applied an additional restriction to Kanal’s ion collection result, producing the widely-
used formulation
2
2
1/20
2 3exp
! 2
m
ii i t i i
m
SJ en v S m
m
(3.28)
Although Kanal’s results are general enough to handle sheaths of varying size, Equation
(3.28) requires the assumption that sheath thickness is negligible. The ion thermal
speed,𝑣𝑡𝑖 , is given by Equation (3.14), while 𝑆𝑖 is the ion Mach number
49
sini
ti
US
v (3.29)
𝑈 is the bulk flow velocity and θ is the angle between the cylindrical probe axis and the
flow vector. In the results presented here, θ is always 90 ̊.
The ion current density given by Equation (3.28) is plotted in Figure 3-6, showing how
the collected ion current varies with ion Mach number. 𝐼𝑟𝑎𝑡𝑖𝑜 refers to the current
collected by a transverse probe normalized by the ion current collected by an identical
probe aligned with the flow.
Figure 3-6 shows that the Kanal theory is appropriate for low to moderate Mach numbers,
but breaks down for ion Mach numbers greater than about 9. In the low speed limit, this
theory correctly predicts that the ion current collected by a transverse probe will be the
same as the ion current collected by an aligned probe. However, in the high speed limit,
this theory predicts that the collected ion current will go to zero. This result is clearly
nonphysical, so an alternative theory is desired for very high speed flows encountered in
some time of flight data obtained from μPIT.
An early theory for collected ion current in a flowing plasma developed by Langmuir and
Mott-Smith can be used for very high speed flows. However, this model is valid only in
Figure 3-6: Ion current collection theory of Kanal for thin sheaths
50
the Orbital Motion Limited (OML) regime, which assumes a collisionless plasma and a
very large sheath so that
1p
D
r
(3.30)
To reduce the Langmuir theory to a case of interest, we choose a 90 ̊ angle of attack and
also assume that the flow is fast compared with the Bohm velocity
1/e i
U
kT m (3.31)
With these additional constraints, the Langmuir theory yields a simple expression for ion
current density collected by the probe
iJ en U (3.32)
The corresponding collection area is just the cross-sectional area of the cylindrical probe
tip: 𝐴⊥ = 2𝑟𝑝 𝑙 . Since this expression for ion current density is accurate only in the very
large sheath limit, it is likely to overestimate the plasma density in situations more
appropriate to the thin-sheath approximation.
3.3.3 Modified Theory for Double Probes in Crossflow
Using the ion current densities developed in section 3.3.2 with the standard Langmuir
probe theory given in section 3.3.1, current-voltage models specific to the μPIT plume
operating regimes are developed. The current to a probe is given by Equation (3.12).
Choosing a generic ion current density 𝐽𝑖0 and allowing for unequal probe areas yields the
following system
1
1 1 1 0
2
2 2 2 0
exp
exp
s
t e i
e
s
t e i
e
eI A env A J U
kT
eI A env A J U
kT
(3.33)
51
𝐽𝑖0 is some function of the flow velocity. Note that quasineutrality is assumed here: 𝑛𝑒 =
𝑛𝑖 = 𝑛. Following the method in section 3.3.1, this system can be combined with the
relationships in (3.23) and (3.24) to find an I-V characteristic
1
2
0 2 1 2 1
1
2
exp
exp
b
e
i
b
e
A
T AI J A A A A
A
T A
(3.34)
Here, 𝑇𝑒 is in eV. Two limiting cases appear for this equation. When 𝑒−𝜙𝑏/𝑇𝑒 ≫ 𝐴1⊥/
𝐴2⊥, the ratio of exponentials goes to 1. This reduces Equation (3.34) to
1 02 iI A J (3.35)
Since the bias voltage is defined as 𝜙𝑏 = 𝜙1 − 𝜙2in accordance with Figure 3-4, the
applied bias voltage is generally positive. This means that the situation described by
(3.35) is rarely encountered, but that the alternative limiting case is frequently useful.
When 𝑒−𝜙𝑏/𝑇𝑒 ≪ 𝐴1⊥/𝐴2⊥, the ratio of exponentials goes to -1, reducing (3.34) to
2 02 iI A J (3.36)
These limiting cases are illustrated in Figure 3-7. Equation (3.34) can be rewritten in
nondimensional form
1
21 1
0 2 2 2 1
2
exp
1 1
exp
b
e
i b
e
A
T AA AI
J A A A A
T A
(3.37)
The resulting I-V curve is plotted in Figure 3-7, where 𝐼∗ is the nondimensional current.
In addition to the difference saturation currents, the unequal probe areas shifts the
characteristic off the origin so that current flows even for zero bias voltage. The greater
the difference in probe areas, the more the asymmetric double probe behaves like a single
Langmuir probe. In the case of a single probe, the vacuum chamber walls take the place
of the double probe’s larger collection tip.
52
Figure 3-7: Normalized I-V characteristic for a double probe with a probe area ratio of 500:1
In the case of equal probe collection areas, Equation (3.34) readily reduces to the more
familiar form
02 tanh2
bi
e
I A JT
(3.38)
At this point, the appropriate ion current density may be substituted into Equations (3.34)
and (3.38). For low to moderate speed, (3.28) is appropriate, while (3.32) is used for ion
Mach numbers greater than 9.
3.3.4 Probe Construction and Description
A symmetric double Langmuir probe was constructed and used to sample the plume near
the thruster exit plane, as illustrated in Figure 2-3. The probe tips are made of 1.1 mm
tungsten wire which are fed through a four-bore alumina tube. These probe tips are
electrically connected via crimps to probe leads which are fed through a ¼‖ stainless steel
tube. The grounded steel tube serves to reduce electromagnetic and electrostatic noise.
The formvar-insulated probe leads are twisted together to minimize magnetic pickup. The
probe tips are 4.6 mm long and present a cross-section to the flow of about 5 𝑚𝑚2. The
symmetric double probe is shown in Figure 3-8.
53
Figure 3-8: Symmetric double Langmuir probe
A planar double Langmuir probe was also constructed to sample the μPIT plume, as
shown in Figure 3-9. Probe construction is very similar to the symmetric double
Langmuir, with the primary difference being the large collection plate. This stainless steel
plate is fixed in place with a boron nitride ring and Torr seal. The stinger probe is 0.78
mm tungsten wire and is 9.2 mm long. The plate diameter is 12 mm in diameter and the
area ratio of the plate to the stinger is about 16. The planar probe represents a tradeoff
between highly localized measurements and strong signals. Despite its large size, the
planar probe is useful in this application because the plasma density is very low 10-20 cm
away from the thruster.
Figure 3-9: Side and bottom views of the planar double Langmuir probe
54
Table 3-1: Langmuir probe operating parameters
Parameter Requirement Symmetric double probe Planar probe
Collisionless probe volume
1i e
pr
135 191
Collisionless sheath 1i e
sd
644 644
Thin sheath 100p
D
r
29 20
Electrons unaffected by flow
1tev
U
7.9, 3.0 7.9, 3.0
High speed flow 1B
U
v
6.2, 17 6.2, 17
Several estimates were necessary to generate Table 3-1. Based on previous PPT plume
measurements [Byrne 2002], electron and ion temperatures were estimated at 2 and 1 eV,
respectively, In order to compute the sheath thickness, the maximum voltage difference
between the probe and plasma potential was estimated at 30V. Density and average ion
mass were estimated at 1017 𝑚−3 and 2.77 × 10 −26 kg. The parameters in the final two
rows were calculated for two different exhaust velocities: 30 km/s and 80 km/s. These
velocities were selected based on time of flight data.
Table 3-1 shows that the planar probe and symmetric double Langmuir operate in
roughly the same regimes. The requirement that the plasma be collisionless on the scale
of the probes is easily met. The thin sheath approximation is not a very good one for
these plasma parameters, however. A theory developed by Laframboise [1966] exists for
sheaths that are larger than the thin sheath limit but are poor approximations to the
infinite sheath limit as well. However, the results of Laframboise are not applicable to
transverse plasma flow, so the theory outlined in this chapter is the most appropriate
choice. For a flow speed of 30 km/s, assuming that the electrons are unaffected by the
flow is a reasonably good approximation, but for 80 km/s flow, this is a relatively poor
approximation. This approximation is necessary to treat the issue of electron collection
using the Boltzmann factor. Also, the assumption that the flow is high speed—fast
55
enough to use the Langmuir theory from section 3.3.2—is appropriate only in the 80 km/s
case.
3.4 CCD Camera
The final diagnostic tool used to study the μPIT plume was a CoolSnap CF fast-framing
CCD camera. This camera is controlled with a TTL pulse sent from the Labview control
software. The camera can capture images with a minimum exposure time of
approximately 10 μs. The camera was used to capture images of each iteration of μPIT
including shots in PPT mode. Since the shot time is much longer than the minimum
exposure, several images were captured at different stages during the current pulse in
order to gain information about the time evolution of the plume. A delay of about 6 μs
exists between the time the camera receives a trigger pulse and the start of an exposure. A
small variation in this delay from shot to shot contributed to uncertainty in the portion of
the current pulse that was actually imaged. The camera was also used with longer
exposure times to capture images of entire shots.
56
4 Theoretical Models and Data Reduction
This chapter outlines an analytical model for inductive discharges. In particular, power
absorption is addressed and the influence of the RF fields on breakdown between the
electrodes is discussed. Estimated ionization rates are presented to illustrate the potential
advantages of the μPIT inductive coupling technique. The chapter concludes with a
description of data reduction techniques used to estimate important μPIT performance
parameters.
4.1 Inductive Discharges
Inductive discharges are commonly used in the semiconductor industry for etching
applications. Inductively coupled plasmas also appear in some electric propulsion
applications, such as Astrium’s Radiofrequency Ion Engine (RIT) [Zeuner et al. 2003]
and Stutgartt University’s Adjustable Throttle Inductively Afterburning Arcjet (ATILLA)
[Laure et al 2001]. The RIT series of thrusters are electrostatic accelerators that use an
inductive discharge to generate ions as an efficient alternative to the Kaufman-style ion
source. ATILLA is a 100-kW class spacecraft propulsion concept combining arcjet and
inductive discharge technology. First an arcjet ionizes and heats the propellant, then the
resulting plasma flows through a solenoidal antenna which further heats the propellant
inductively. Intended to provide a variable specific impulse capability, this thruster
employs the concept of inductive coupling in a way similar to μPIT, though on a much
larger scale.
Figure 4-1 shows a common arrangement for an inductive discharge. In contrast to
capacitive discharges, inductive discharges have no electrodes in contact with the plasma
and can maintain relatively low sheath voltages. Additionally, inductive discharges can
operate with greater efficiency, making them a more appealing choice than capacitive
discharges.
57
Figure 4-1: Typical arrangement for an inductively coupled plasma, employing a helical antenna.
Image reproduced from [Bartone 2007].
In an inductive discharge, electric fields generated by the helical antenna transfer power
to the plasma by accelerating the electrons. The heated electrons redistribute their energy
via collisions with neutrals and charged particles and through wave-particle interactions.
This process occurs in a layer with a thickness on the order of the skin depth near the
plasma edge. The antenna can be modeled as a standard solenoid with oscillating current
with a frequency ω and an amplitude 𝐼0
0 expI I i t (4.1)
The electric field components can be calculated from Maxwell’s equations. To begin, a
uniform oscillating axial magnetic field is assumed. The radial and azimuthal field
components are taken to be zero. This field can be calculated for a solenoid using
Ampere’s law by assuming that the field outside the core is negligible. For an 𝑁-turn coil
of length 𝑙, the magnetic field is
0z
N IB
l
(4.2)
The induced azimuthal electric field can be calculated using Faraday’s law with the time
varying field given by Equation (4.2), yielding
58
0
2
N I rE r
l
(4.3)
Since 𝐸𝜃 is linear in radius 𝑟, the induced electric field is strongest at the outer edge of
the plasma, which is typically where power transfer to the plasma occurs. An oscillating
axial electric field component also exists due to the voltage drop along the coil. For this
idealized geometry, the inductance can be calculated from its definition
2 2
0R NNL
I l
(4.4)
for a solenoid of radius 𝑅 with linked magnetic flux Φ. The voltage across the coil is
related to the current by the inductance
d I
V Ldt
(4.5)
Since electric field is just voltage per length, Equations (4.4) and (4.5) can be combined
to form an equation for the axial electric field component
2 2
0
2z
R N IE
l
(4.6)
Equations (4.3) and (4.6) represent the dominant coupling mechanism for inductive
discharges in E-mode and H-mode, respectively. When the RF current is large enough,
the discharge will transition from E-mode to H-mode. In the early stages of a discharge,
the plasma density is low and 𝐸𝑧 plays a more important role in transferring energy from
the antenna. As the density increases, 𝐸𝑧 is shielded out and 𝐸𝑟 dominates at higher
densities. For a current density 𝐽 and an electric field 𝐸, the power transfer per volume is
given by
P J E
(4.7)
Using the constitutive relation 𝐽 = 𝜎𝐸 where 𝜎 is conductivity, Equation (4.7) may be
integrated over the appropriate volume to find the total power absorbed by the plasma. In
the case of E-mode, the electric fields penetrate the whole volume inside the cylindrical
59
antenna since the plasma density is too low to screen them out. For H-mode, power is
only absorbed in a thin region at the outer edge of the plasma with a thickness equal to
the skin depth. Lieberman (1) developed power absorption scaling for E-mode
2
0absP n I (4.8)
and for H-mode
2
1/2
0absP n I (4.9)
The most important feature from these scaling results is the proportionality with the
square of the RF current, showing that high antenna current is critical for good coupling.
Also noteworthy is the dependence on plasma density, 𝑛0. For sufficiently high RF
currents, an inductive discharge can rapidly transition from E-mode to H-mode since the
power absorbed scales with plasma density, and more absorbed power leads to higher
density. However, this positive feedback effect ends when enough plasma is present for
H-mode operation, as the power absorbed begins to fall with increasing density. At very
high densities, it is possible for the electron-ion collision frequency to be larger than the
electron-neutral collision frequency. In this situation, the plasma density dependence
drops out, leaving
2
absP I (4.10)
While this regime is probably not accessible in μPIT, it may be possible to operate this
way in a larger-scale PPT like the one used on EO-1.
The equations for the electric field components developed above give an accurate picture
of the fields inside the antenna that produce the inductively coupled plasma, but the
assumption of a uniform axial magnetic field is too restrictive to analyze the fields
outside the coil. Specifically, the magnetic and electric fields change close to the ends of
the coil as the magnetic field begins to curve outward, as illustrated in Figure 4-2.
60
Figure 4-2: Snapshot of the magnetic field geometry near the electrodes. The field completes one
oscillation every 3 μs. In μPIT, the separation between the first antenna turn and the Teflon face is
only a few mm.
The oscillating fields in this region can play a role in the breakdown process. Since the
electrons are magnetized through the majority of each cycle, the radial magnetic field
component provides a path for electrons, reducing the minimum voltage required for
breakdown. This radial magnetic field component also modifies the azimuthal electric
field result given by Equation (4.3). Relaxing the earlier assumption about the magnetic
field, we allow a radial component: z rB B B
. Applying this field formulation to
Faraday’s law gives
z rB BE z r
t t
(4.11)
Writing curl in cylindrical coordinates and assuming azimuthal symmetry so that no
gradients in θ are allowed, we have
1 1z r
EB Bz r rE z r
t t r r r z
(4.12)
The 𝑧 component of Equation (4.12) gives the same equation for 𝐸𝜃 𝑟 developed earlier,
but the 𝑟 component accounts for the effect of the radial magnetic field component. The
effect of the curving magnetic fields is to add a gradient in the z direction to the
azimuthal electric field. This gradient comes from the decreasing magnetic flux though a
surface in the r-θ plane as the distance to the antenna increases. Since 𝐸𝜃 drops off with
𝑧, positioning the antenna adjacent to the Teflon face maximizes energy coupling into the
arc plasma and associated neutrals.
Anode
Teflon
Cathode
B field lines
Antenna
61
4.2 Estimated Ionization Rates
A kinetic model was used in order to determine the potential usefulness of the antenna in
ionizing neutrals during and after the arc. The number of ionizing collisions per volume
per time between species A and B with relative velocity 𝑣 can be calculated from
0 A B in n n v (4.13)
where 𝜎𝑖 is the ionization cross section. The quantity 𝜎𝑖𝑣 is the product of the
ionization cross section and velocity averaged over all possible velocities. This can be
computed by convolution with the speed distribution ( )v
0
, ,i iv v T v T vdv
(4.14)
where the speed distribution is given by
3/2 2
24 exp2 2
m mvv v
kT kT
(4.15)
In general, the ionization cross section may be a function of velocity and temperature
which can either be measured or calculated using quantum mechanics. The average cross
section-velocity products have been tabulated for many elements and species
temperatures. The following values were taken from Japan’s National Institute for Fusion
Science online database: for fluorine at 4.9 eV, 𝜎𝑖𝑣 = 2.1 × 10−9; for carbon at 4.9 eV,
𝜎𝑖𝑣 = 3.4 × 10−9 [𝑐𝑚3𝑠−1]. While the relative concentrations of charged particles in
PPT plumes have been measured by several groups [Koizumi et al. 2007, Hirata et al.
1982, and Shumlak et al. 2004], the relative concentrations of neutral particles remains
unknown. In the following analysis, the neutral particles are assumed to consist of atomic
fluorine and carbon in the ratio 2:1.
In order to apply Equation (4.13), 𝑛𝐴 will be replaced with the electron density and 𝑛𝐵
will be the neutral density. While the electron density can be measured using Langmuir
probes, the neutral density is unknown. Based on the results of Spanjers et al. [1996], the
62
neutral density will be estimated at five times the peak electron density. The vapor
pressure 𝑝 can be modeled with an exponential relationship to the surface temperature 𝑇𝑠
exp cc
s
Tp p
T
(4.16)
where 𝑝𝑐 and 𝑇𝑐 are the tabulated characteristic pressure and temperature for the material.
The surface temperature, in turn, decays with time as the Teflon cools due to radiation
and evaporation. However, the time constant for this cooling is assumed to be much
larger than a typical shot time of 150 μs. As a result, the neutral density is taken to be
approximately constant throughout a shot. This model implies that the neutral particles
produced in a PPT discharge greatly outnumber the charged particles, which is consistent
with the measurements of Koizumi and Spanjers.
In an additional simplification, the quantity 𝜎𝑖𝑣 for carbon and fluorine is combined as
a weighted average 𝜎𝑖𝑣 𝑎𝑣𝑔 to account for the greater concentration of fluorine. With
these assumptions, Equation (4.13) can be rewritten
2
0 5 e i avgn n v (4.17)
Equation (4.17) allows for the calculation of an ionization rate. In order to compare the
quantity of plasma produced by the antenna to plasma produced by the arc, the total
number of ions produced 𝑁𝑖 is estimated using the following method
0i r eN A c n t dt (4.18)
An exhaust velocity 𝑐𝑒 is assumed and multiplied by the time integral of the density as
the plasma flows out of the antenna; this is equivalent to integrating the density along the
𝑧 direction, leaving units of 𝑙𝑒𝑛𝑔𝑡−2. The expression is multiplied by a reference area
𝐴𝑟 to convert the result to a number of particles. To calculate 𝑁𝑖 for the arc plasma, a
typical density trace collected near the antenna exit plane is used along with an exhaust
velocity measured using a time-of-flight technique. To calculate 𝑁𝑖 for the antenna
plasma, an ionization rate from Equation (4.17) is computed using the peak electron
63
density from the arc plasma data. This ionization rate is multiplied by a typical shot time
of 150 μs to estimate an antenna plasma density. Equation (4.18) is evaluated over the
shot time by assuming this density to be constant and by choosing a reasonable exhaust
velocity.
For a peak electron density of 1017 𝑚−3and an ion thermal speed of 3 km/s
(corresponding to 10 eV ions), the ratio of antenna plasma particles to arc plasma
particles is 0.29. While this scenario only ionizes small portion of the neutrals, it can still
have a significant impact on impulse bit and mass utilization efficiency. The 𝑛2 scaling in
Equation (4.17) means that this technique for ionizing neutrals would be particularly well
suited to large-scale PPTs with high discharge energies. For example, choosing a peak
electron density of 1018 𝑚−3 and the same exhaust velocity yields an antenna/arc particle
ratio of about 3. To reach a particle ratio much larger than 1, a peak electron density on
the order of 1020 𝑚−3 is required, easily within the range of the LES 8/9 and EO-1 PPTs.
4.3 μPIT Performance Estimation
A technique was developed to estimate impulse bit and thrust efficiency using radial
density profiles, time-of-flight velocity results, and capacitor voltage measurements. To
compute the impulse bit, knowledge of the velocity 𝑐𝑒 and mass 𝑚𝑝 of the plume is
required.
bit e pI c m
(4.19)
The plume was assumed to be axisymmetric so that integration of the density was only
required in 𝑟 and 𝑧 to find the propellant mass. Propellant mass is given in cylindrical
coordinates by
,p im dV m n r z rdrdzd (4.20)
where 𝜌 is mass density and 𝑚𝑖 is weighted-average ion mass, assuming two fluorines
per carbon. Integrating over all 𝜃 and converting to time with 𝑑𝑧 =𝑑𝑧
𝑑𝑡𝑑𝑡 = 𝑐𝑒dt yields
64
0
0 0
2 ,
ftR
p i em m c n r t dtrdr (4.21)
for a cylindrical volume with radius 𝑅0 and a shot time ending at time 𝑡𝑓 . To account for
the vector nature of Equation (4.19), the exhaust velocity is rewritten as
cose ec c
(4.22)
Here, 𝜙 is the angle between the z-axis and the radial coordinate. Physically, this term
ensures that only the component of momentum flux that produces useful thrust is
included in the impulse bit calculation. Based on the geometry, the cosine can be
rewritten in terms of 𝑟
0
1/22 2
0
cosz
z r
(4.23)
𝑧0 is the axial coordinate where radial density profile was measured. This cosine
relationship was also used to correct the planar probe area as viewed by the plume for
off-axis locations. Radial density data was collected at 𝑧0 = 5 cm and 15 cm. Combining
Equations (4.21)—(4.23) with the expression for impulse bit in (4.19) yields
0
2 0
1/22 2
0 0 0
2 ,
ftR
bit i e
rzI m c n r t dt dr
z r
(4.24)
This equation was evaluated numerically in Matlab using several density traces. Three
density traces were collected and averaged for each radial position, with a spacing in 𝑟 of
about 1 cm. At 5 cm and 15 cm downstream, 𝑅0 was 2.8 cm and 7.5, respectively. The
general data reduction approach involved a Matlab script for converting Langmuir
current measurements to density according to the theory outlined in Chapter 3. The script
also eliminated any vertical offset present in the waveform, smoothed the data with a 10-
60 point boxcar filter depending on the noise level, and saved the waveform to a data file.
It is important to note that the formulation given by Equation (4.24) only accounts for the
impulse generated by charged particles in the plume. While neutral gas is expected to
65
make up the majority of ablated mass per shot, the low velocity associated with neutrals
means that they contribute very little impulse. For example, neutrals evaporating from a
Teflon surface at 600 K would have a thermal velocity of about 550 m/s compared with
20-55 km/s for the electromagnetically accelerated propellant in typical PPTs [Burton et
al. 1998]. 600 K is the approximate minimum temperature for Teflon decomposition
[Turchi et al. 2000].
The exhaust velocity is a critical element of both the impulse bit estimate and the density
calculations for a flowing plasma. To measure exhaust velocity, time-of-flight (TOF) was
used. The symmetric double Langmuir was positioned a few millimeters downstream of
the antenna exhaust plane and the planar Langmuir probe was located 5 to 15 cm further
downstream. The separation between the probes was always known to within 5 mm (and
usually to within 1 mm). Both probes were positioned on-axis. Density data was collected
on the same oscilloscope to ensure synchronized waveforms. The time between the
density peaks was measured to an accuracy of about 500 ns. The known probe separation
and time lag between density peaks enables the calculation of the time-of-flight for the
plume. Based on the accuracy of the time and distance measurements, velocity
uncertainty is estimated at 14% for a 5 cm probe separation and 8 % for a 15 cm
separation.
The exhaust velocity was also necessary for calculating thrust efficiency. Defining thrust
efficiency as the ratio of impulse generating energy out to input energy results in the
equation
2 2
0
2 e bitt
f
c I
C V V
(4.25)
The denominator represents the energy drained from the storage capacitor, so 𝐶 is the
capacitance, 𝑉0 is the initial storage cap voltage, and 𝑉𝑓 is the final voltage. This
formulation of thrust effiency does account for losses within the RF power supply, but
does not include losses associated with the small DC-DC supply used to charge the
storage capacitor.
66
5 μPIT Development and Discussion
Following the proof of concept prototype, three μPIT iterations were constructed and
tested in vacuum. The most important feature distinguishing each iteration was the
transformer, so the different versions are referred to as the 8:120, 8:255, and 8:185
models. Testing of the original prototype showed a high density arc plasma and evidence
of late time plasma production (LTPP). However, some limitations of the prototype were
also apparent. Specifically, the breakdown process demonstrated poor repeatability,
sometimes requiring a gas puff to continue firing. The increased difficulty in achieving a
reliable breakdown is a consequence of the choice of ignition mechanism for μPIT.
Additionally, thruster lifetime was fairly limited; the prototype failed after a few hundred
shots. Due to these findings from the original prototype, the first goal of the subsequent
models was to enable μPIT to fire reliably without a gas puff for a greater number of
shots. After repeatable operation was achieved, additional goals included a more detailed
study of LTPP, investigation of the effects of the antenna on the arc plasma and
breakdown physics, and estimation of impulse bit and thrust efficiency.
5.1 8:120 Transformer
The first iteration was designed with a theoretical transformer step-up of 15, increased
from 10 in the original prototype. The rationale for this increase was that a higher voltage
would reliably force a breakdown across the Teflon. A tuning procedure was performed
to choose the optimum operating frequency, which was found to be 317 kHz. The
antenna and igniter characteristics were measured at low voltage in air and a linear
extrapolation was performed to estimate power supply performance at the 350V storage
capacitor charge. The electrical characteristics are displayed in Table 5-1.
The listed antenna current and voltage are the maximum peak-to-peak measurements
over the course of an entire shot. The igniter voltage was measured using a high voltage
divider and specifically refers to the igniter capacitors’ voltage. Since these capacitors are
Table 5-1: Estimated electrical characteristics of the 8:120 μPIT with a storage cap charge of 350 V
Antenna current [A] Antenna voltage [V] Igniter voltage [V] Transformer efficiency 1130 1240 6080 65 %
67
located on the output side of the rectifying bridge, they only acquire half the peak-to-peak
voltage generated on the transformer secondary. The transformer efficiency is defined
here as the ratio of the measured voltage on the secondary to the theoretical maximum
voltage based on the turn ratio. The relatively low transformer efficiency is mainly
attributable to flux leakage. At lower operating frequencies, higher efficiency is
attainable through the use of ferrous transformer cores which drastically reduce flux
leakage. However, in the 100+ kHz range the resistive losses from eddy currents in the
core material begin to overwhelm the flux leakage savings. It has recently been
established that more closely spaced turns in the primary winding can significantly
improve the transformer coupling efficiency in the μPIT ignition system.
Also noteworthy in Table 5-1 are the high antenna current and voltage; due to the 𝐼2
scaling, this current is critical for significant ionization and heating as the propellant
flows through the antenna core. The resonant tank network drives the initial storage
capacitor voltage more than three times higher, reducing the burden on the transformer to
generate high voltage for breakdown.
Despite the transformer step-up increase, vacuum testing of the 8:120 model
demonstrated that more voltage was necessary to achieve reliable breakdown. The system
did not fire on its own at a charge of 350 V, so it was fired several times using an argon
gas puff to aid in the discharge. This is believed to have a cleaning effect on the
electrodes and Teflon face. After 10-20 gas shots, the μPIT begain firing on its own.
Successful operation without gas continued for about 25 shots, then another round of
argon shots was required. Density data was collected from each shot with both the
symmetric and planar Langmuir probes and a typical set of waveforms is presented in
Figure 5-1.
68
Figure 5-1: Typical density traces for the 8:120 μPIT (without gas puff). Storage capcitor charge was
350 V and both probes were positioned on-axis.
Zero on the time scale is defined as the antenna switch-on time and a breakdown time of
28 μs is fairly typical of the 8:120 μPIT. A noteworthy feature in Figure 5-1 is the high
density measured by the symmetric double Langmuir just downstream of the antenna.
However, the plasma density in the plume falls by two orders of magnitude over 6.5 cm,
suggesting severe plume divergence. The high plasma density at the antenna exit is
unlikely to be an instrumental effect because the ratio of the symmetric double Langmuir
signal to the planar probe signal was consistently on the order to 100 for the 8:120 model,
while this ratio was closer to 10 for the subsequent 8:255 and 8:185 μPIT iterations. No
changes were made to the planar probe, and the only change to the double probe was the
replacement of the current source capacitor with a larger one, which would increase the
current drawn by the probe if anything.
Another feature evident in Figure 5-1 is the broadening of the axial extent of the plume
between 5 mm and 7 cm. This implies a spread in the velocity distribution of the arc
plasma. While the velocity distribution was not measured, an average exhaust velocity
69
was measured for each shot using time-of-flight. The average TOF exhaust velocity was
82 km/s. This is 2-3 times higher than typical exhaust velocities for electromagnetic
PPTs. In fact, the high velocity measurements in this and subsequent iterations suggest
that μPIT functions primarily as an electromagnetic accelerator, as opposed to a thermal
accelerator. However, the dataset was limited to less than ten useful shots and the data
exhibited significant scatter. The standard deviation, for example, was 35 km/s. As this
was the only velocity data available for the 8:120 μPIT, it was used in conjunction with
the fast-flow theory of Langmuir and Mott-Smith outlined in Chapter 3 to reduce the
Langmuir probe data.
Figure 5-2 shows one of the new contributions that μPIT brings to the field of PPTs.
Late-time plasma was generated throughout the shot and detected by both Langmuir
probes. These measurements are consistent with other researchers’ reports of neutral gas
remaining near the Teflon face on a time scale much longer than the initial arc. The
Figure 5-2: Evidence of late-time plasma production after the arc plasma has passed by. No gas was
injected.
70
antenna turns on at t = 0 and shuts off at 158 μs. Breakdown occurs at 16 μs and a large
density pulse is evident on both probe signals. The density traces from both probes
clearly settle into a mode corresponding to low-level, DC plasma production lasting
several times longer than the arc plasma. The plasma density at both axial locations drops
off at a time closely matching the antenna shut-off time. Based on the late time neutral
density estimates of Koizumi [2007] and Spanjers [1996], the plasma produced by the
antenna in Figure 5-2 only represents a small fraction of the neutral population available
in the 300+ μs after the arc. Nevertheless, these results illustrate μPIT’s potential to make
significant gains in PPT efficiency.
A high voltage divider was used to monitor the igniter capacitor voltage over the course
of each shot. The average voltage just before breakdown occurred was 5.2 kV. For a gap
size of 0.8 mm, this corresponds to an electric field of 6.5 MV/m applied across the
Teflon face. Since the igniter capacitors are so small, the corresponding arc energy is just
52 mJ. This arc energy is 1000 times less than the standard energy per pulse used in the
LES 8/9 PPT [Spanjers 1996] and just 2.5% of the energy per pulse used in Busek’s
flight model of the AFRL μPPT [Busek 2007b]. The linear scaling of peak electron
density with energy suggested by Spanjers et al. was tested using a modified LES 8/9
PPT over a range of energies between 2 and 80 J. If peak plasma density is linear with
energy, the ratio of peak density to energy should be the about the same for the LES 8/9
and μPIT measurements. For LES 8/9, 𝑛𝑝𝑒𝑎𝑘
𝑊= 5 × 1020 𝑚−3𝐽−1 while for μPIT, the
ratio equals 2 × 1019 𝑚−3𝐽−1. These differ by an order of magnitude, while Spanjers’
results vary by a factor of 2 at most. This discrepancy may result from a difference in the
energy/area ratio, another figure of merit for comparing PPTs. The 𝐸/𝐴 ratio for the LES
8/9 PPT was about 2 𝐽/𝑐𝑚2, about 20 times higher than the 𝐸/𝐴 ratio for μPIT. A higher
𝐸/𝐴 produces a higher Teflon surface temperature, and Equation (4.16) shows that an
increased surface temperature corresponds to greater neutral ablation. Increasing the
energy to area ratio for μPIT by using larger igniter capacitors would likely cause the
density results to match better with the results of Spanjers et al.
71
Figure 5-3: Antenna loading in μPIT
In addition to the effects of the antenna on the late-time neutrals, the antenna also appears
to couple directly with the arc plasma as it passes through the transformer. Figure 5-3
shows an unloaded current trace in the upper panel. This current trace was collected
during one of the late breakdown shots mentioned earlier in which breakdown occurred
well after the antenna shut off. Fortuitously, this resulted in the capability to measure the
antenna current at full power but without any perturbations due to breakdown or the
presence of plasma in the antenna. The lower panel of Figure 5-3 shows the same
unloaded current trace superimposed on a loaded current trace collected from a shot in
which breakdown occurred as intended at around 27 μs. A density trace from the planar
probe located on axis, 7 cm downstream is also shown. The density and perturbed current
traces were collected from the same shot. Up to around 26 μs, the loaded and unloaded
current traces match almost exactly. After breakdown occurs, the loaded current trace
shows a significant reduction in amplitude as the arc plasma presents a high impedance to
the antenna and absorbs power in the process. The effects of the antenna on the arc
plasma are explored further in Section 5.3 with the 8:185 iteration.
72
The presence of plasma in the antenna also leads to a phase shift in the current pulse as
compared with the vacuum case. The induced mirror current in the plasma decreases the
circuit inductance, which in turn increases the antenna frequency. An antenna voltage
diagnostic would allow for a calculation of the power in the vacuum and plasma cases
using 𝑃 = 𝐼𝑉. Comparing the power in these two cases, the power transfer from the
antenna to the plasma could be estimated. Additionally, a comparison with the increase in
thrust power could be made. A high voltage divider was developed for this purpose, but it
produced a noise signal an order of magnitude greater than the expected signal. A
combination of magnetic and electrostatic pickup from the floating electrodes was
suspected, and this diagnostic was removed.
5.2 8:255 Transformer
Because of the difficulties involved in collecting useful data from a thruster that cannot
fire reliably, the primary goal after testing the 8:120 model was to redesign the ignition
system to ensure repeatable, gas-free operation. With that goal in mind, a new
transformer was constructed with a theoretical voltage step-up of nearly 32. Assuming
approximately the same efficiency, this transformer was expected to generate about twice
the peak voltage as the previous version, or about 12 kV. At more than twice the average
breakdown voltage of 5 kV, this peak transformer voltage was expected to eliminate all
late and failed breakdown events.
Table 5-2 lists electrical characteristics of the antenna and ignition system for the 8:255
transformer. The corresponding figures for the 8:120 model are repeated here for
comparison. Like the 8:120 figures, these results are estimated based on a linear
extrapolation of measurements at low power. Using the calibrated Rogowski for tests in
vacuum at high power, the projected antenna current shown here was found to be
accurate to within 5 %.
Table 5-2: Estimated electrical characteristics for the 8:120 and 8:255 models. Storage capacitor
charge is 350 V.
μPIT transformer
Antenna current [A]
Antenna voltage [V]
Igniter voltage[V]
Transformer efficiency
8:120 1130 1240 6080 65 % 8:255 850 1410 10,500 50 %
73
The most salient difference between the two transformers is the peak igniter voltage,
though the igniter voltage for the 8:255 is not quite twice that of the 8:120 transformer
due to the reduced efficiency. The reduced efficiency is probably caused by the additional
flux leaked by the way the 8:255 transformer was wound. In order to fit 255 turns on the
secondary over the same 2.5 cm long quartz tube, smaller gauge wire was used and a
third layer of turns was wound. Because of the greater distance from the primary
winding, this third layer is likely to link with less flux than the first and second layers,
resulting in a lower overall coupling efficiency. Also notable in Table 5-2 is the shift in
the antenna toward higher voltage at the expense of current. While this shift allows for
higher igniter voltage, it also leads to a reduction in power transfer from the antenna to
the plasma.
Figure 5-4 shows a histogram of breakdown times for the 8:120 and 8:255 transformers.
This data was collected from about 30 shots with each transformer. This figure illustrates
Figure 5-4: Histogram comparing breakdown times for the 8:120 and 8:255 transformers
74
the greatly improved reliability of the 8:255 transformer. Note that five of the 8:120 shots
have a negative breakdown time; these denote shots that either failed to break down at all
or broke late, after the antenna shut off. In addition to these five irregular shots, the 8:120
breakdown times are relatively diffuse and cover a wide range. Conversely, the 8:255
breakdown times are sharply peaked at about 20 μs. Furthermore, every shot breaks down
within 40 μs for the 8:255, while shots with the 8:120 occasionally do not break down at
all. The increase in voltage associated with the 8:255 transformer successfully met the
goal of repeatable breakdown. However, this increased voltage also created new
problems.
The first difficulty presented by the increased voltage in the 8:255 transformer was a
tendency for μPIT to occasionally arc within the transformer or ignition system
electronics. The system would generally continue functioning after such an arc occurred,
suggesting a robust power supply and ignition electronics. However, arcing of this kind
ruins data with a cascade of noise and is highly undesirable from an operational
perspective because of the increased risk to equipment. A combination of epoxy and
Kapton tape was applied to high voltage sections of μPIT to insulate the surfaces against
arcing and the standard charge voltage was reduced from 350 V to 275 V. These
measures greatly reduced the occurrence of arcs.
A second problem arising from the increased voltage is the reduced current available to
charge the igniter capacitors. Figure 5-5 shows two typical high voltage divider traces
that demonstrate how the igniter capacitors charge prior to breakdown. The antenna
switches on at t = 0 and breakdown occurs at 12.5 and 18.1 μs for the 8:255 and 8:120
transformers, respectively. The 8:120 charges to about 5.3 kV while the 8:255 only
reaches 1.8 kV before breaking down. Since the same Teflon, electrodes, and other
equipment were used for both shots, the same voltage is expected to initiate breakdown in
each case. It is important to note that the 8:255 shot started with a 275 V charge on the
storage capacitor, while the 8:120 started at 350 V. With an initial charge of 275 V the
8:255 secondary is projected to reach a peak voltage of 8.2 kV and peak antenna current
of 680 A, compared with 1130 A in the 8:120 primary. Assuming 100% efficient
transformers, power is conserved between the primary and secondary, resulting in a
75
Figure 5-5: Time-resolved igniter voltage for the 8:120 and 8:255 transformers
secondary current that is reduced by a factor 1
𝑁 for a secondary/primary turn ratio 𝑁. As
pointed out in Table 5-2, the transformers have significant losses and the 8:255 is
substantially less efficient than the 8:120. This means that the secondary current is
reduced even further, particularly for the 8:255 transformer. The combined effects of the
reduced charge voltage, reduced transformer efficiency, and increased turn ratio lead a
significantly lower current available to charge the igniter capacitors. Reduced current to
the igniter capacitors, in turn, should result in a substantially reduced arc energy and
corresponding plasma and neutral densities.
While it accounts for the reduced arc energy evident in Figure 5-5, the lower current in
the 8:255 secondary does not explain why the 8:255 breaks down at less than 2 kV while
the 8:120 required an average of 5.2 kV. Since the electrodes are connected to the
secondary leads directly through a rectifier, the high voltage 300 kHz signal generated by
the transformer appears across the electrodes as well. A possible explanation for the arc
at an apparent voltage of 1.8 kV is that breakdown occurs as soon as the antenna rings up
enough to apply around 5 kV across the electrodes even though the igniter capacitors are
not yet charged to match that potential. The limited current prevents the capacitors from
76
Figure 5-6: 8:255 plasma density measured by the planar Langmuir probe at 5 cm downstream, on-
axis.
charging fully, resulting in a less energetic discharge. The transient voltage spikes that
cause the earlier breakdown should be visible in the igniter voltage current trace.
However the waveforms in Figure 5-5 have been low-pass filtered and smoothed with a
boxcar due to large-amplitude, high frequency noise on the raw signal. The poor signal to
noise ratio in the data obscures the 300 kHz oscillation from the antenna.
The idea that the igniter capacitor isn’t fully charged when breakdown occurs is
supported by Figure 5-6. This figure shows plasma density traces for the 8:255
transformer at several different storage capacitor charge voltages. A typical density trace
from the 8:120 transformer is shown for comparison. Although the peak densities for
both transformers are about the same, the width of the 8:120 trace is significantly greater
than any of the 8:255 traces, implying a greater quantity of plasma released in the arc as
long as the exhaust velocities are approximately the same. The same peak density as the
8:120 but decreased pulse width is consistent with the following conditions. The 8:255
μPIT charges to around 5 kV with approximately 10% of the energy stored in the igniter
capacitors of the 8:120 μPIT. The capacitors discharge at 5 kV in both cases, developing
77
about the same peak density, but the 8:255 μPIT drains its capacitors more rapidly and
runs out of current more quickly than the 8:120. The result is a thinner density pulse as
viewed by the planar probe at the same downstream position.
Another noteworthy feature of Figure 5-6 is the variation in breakdown times: the higher
the charge voltage, the shorter the breakdown time, and the 8:120 breaks down last. This
pattern is due to the fact that the antenna rings up to the breakdown threshold more
rapidly as the charge voltage is increased.
Figure 5-7 shows a selection of zoomed-in density traces to examine LTPP in the 8:255
μPIT along with some 8:120 traces for comparison. The breakdown times have been
shifted to zero to facilitate comparisons. As no TOF measurements were performed for
the 8:255 transformer, a velocity of 80 km/s was assumed for the density data analysis in
Figure 5-7: Zoomed-in plasma density traces from the 8:255 μPIT at various storage cap voltages
78
order to facilitate comparisons. The black trace in the upper panel appears to follow the
pattern of LTPP observed in the 8:120 μPIT, though at a lower level. The waveform
settles down to a low-level DC trace with some oscillation from antenna noise, then drops
out near the end of the shot. The corresponding planar probe signal is shown in blue in
the lower panel, which was collected about 5 cm further downstream. This waveform
exhibits the same pattern and more clearly shows a density signal that exceeds zero.
However, comparison with the 8:120 traces from both probes demonstrates that the 8:255
transformer generates less late-time plasma. This is likely due to the decreased arc energy
and corresponding decreased in neutral density in the antenna after the arc plasma is
expelled.
Although some LTPP traces were recorded from the 8:255 transformer, it did not always
generate late-time plasma. The lower panel in Figure 5-7 also shows two examples of
alternative operating modes. At a charge of 350 V, the high level of noise following the
arc plasma suggests that a quasi-DC low-current arc develops across the Teflon face.
However, the low current leads to a very low ablation rate which, like the antenna
coupling, is known to scale like 𝐼2 [Ruchti and Niemeyer 1986]. This mode of operation
is only seen to occur at a 350 V storage capacitor charge. Presumably, this occurs
because at 350 V charge, the electrodes continuously see > 5 kV through the duration of
the shot, preventing any substantial charging of the igniter capacitors. An operating mode
that occurs interchangeably with the LTPP mode is the red ―sawtooth‖ mode shown in
the lower panel. This behavior appears to be caused by the repeated charging and
discharging of the igniter capacitor, which occurs when there are too few neutrals and
residual electrons for the antenna to couple with directly. Although the sawtooth and
LTTP modes occur for the same operating conditions, the LTTP mode was more
common early in the 8:255 vacuum testing campaign, while the sawtooth mode came to
dominate after a few dozen shots. This transition may indicate a dependence on the
condition of the Teflon surface, which was shown by Keidar et al. [2004] to accumulate
carbon especially at low discharge energy.
79
5.3 8:185 Transformer
After observing the changes in operation due to the higher voltage transformer, the
ignition system was modified. Additional rectifying bridges were added in parallel with
the original one in order to reduce the possibility of a current bottleneck there and a new
lower-voltage transformer was constructed. The new transformer was designed for a step-
up halfway between the 8:120 and 8:255. A hybrid of the two μPIT models was sought
that would combine the higher arc energy and density of the 8:120 with the reliable
breakdown of the 8:255. Ideally, the new transformer would generate a high enough
voltage to ensure that breakdown occurs in a repeatable way without gas, but also a low
enough voltage to avoid arcing before the igniter capacitors are fully charged. A density
single pulse followed by a sustained, DC late-time plasma was desired.
The 8:185 transformer was also used to quantify the effects of the antenna on the PPT
discharge. With this goal in mind, the 8:185 μPIT was tested in the standard
configuration (PIT mode), then tested again in a configuration referred to as PPT mode.
In PPT mode the antenna was still present in the system, but it was physically separated
from the arc plasma so that it only served as the inductor in the resonant LRC network.
Figure 5-8: Photographs of the for PPT mode (left) and PIT mode (right) configurations for the 8:185
μPIT
80
Table 5-3: Electrical characteristics of three variants of the 8:185 μPIT at 350 V charge
Configuration Antenna
Current [A] Antenna
voltage [V] Igniter
voltage [V] Transformer
efficiency PIT, low I 620 1580 7640
42 % PPT 940 1550 7540 PIT, high I 1190 1390 6740
Figure 5-8 shows how the experimental arrangements for PPT mode and PIT mode
compare. A nearly identical quartz tube was placed over the electrodes in PPT mode
maintain any wall effects present in PIT mode. The μPIT was tested under vacuum first
in PIT mode, then in PPT mode. Afterward, additional radial cut and TOF data in PIT
mode was desired, so the system was reconfigured again. Interestingly, the antenna
current was found to increase significantly with this system rebuild. One possible
explanation is that the solder joint in the earlier PIT configuration contributed an
unusually high contact resistance, reducing the tank circuit Q factor. Electrical
characteristics of both the high and low current PIT modes and PPT mode are listed in
Table 5-3.
μPIT was driven at 345 kHz in both PIT modes and at 317 kHz in PPT mode. The
transformer efficiency is the same in each case because a single antenna/transformer
assembly was used in each operating mode; only its physical location varied. The
transformer efficiency is unexpectedly lower than that of the 8:255 transformer. This
reduction is perhaps due to additional Kapton insulation in the 8:185 transformer which
could lead to marginally higher flux leakage losses. Table 5-3 shows that as intended, the
maximum igniter voltage about halfway between the peak voltage of the 8:120 and the
8:255 transformers, which were 5.3 kV and 10.5 kV, respectively. Greater coupling with
the plasma and neutral gas is expected for the high-current PIT mode because of the
increased current.
Despite the insulation improvements, an arc occurred early in the first round of vacuum
testing of the PIT configuration of the 8:185 μPIT. An image of the event was captured,
as shown in Figure 5-9. Since this picture was taken through the glass bell jar, some
reflections are visible in the upper right portion of the image. While the transformer
81
Figure 5-9: A 200 μs exposure that captured an arc in the ignition system electronics
windings withstood the high voltage, the figure indicates an arc occurred at one of the
transformer leads. The other breakdown point needed to complete the circuit is not
visible, perhaps obscured in the saturated portion of the CCD array. Inspection in air after
the event did not reveal the other breakdown point, so a large quantity of epoxy was
applied throughout the ignition system. Along with the previous improvements in
insulation, this modification prevented any unintended arcs from occurring. The system
was subsequently fired several hundred times without incident.
5.3.1 Late-time Plasma Production Results
After the insulation problems were solved, a study of the LTPP performance of the 8:185
μPIT was undertaken. Figure 5-10 shows zoomed-in density traces collected with the
symmetric double Langmuir and the planar Langmuir probe. The breakdown times have
been shifted to zero for comparison. The double Langmuir was located 4 mm
downstream of the quartz tube exit plane and the planar probe was located an additional 5
cm downstream. The upper panel shows typical double Langmuir waveforms for the
8:185 transformer in PIT and PPT modes. In PIT mode, μPIT exhibits a low-level, quasi-
DC density signal, following the general pattern of LTPP in the 8:120 μPIT. However,
82
this signal drops to zero shortly before the 100 μs mark, well before the antenna shuts off.
The PIT mode fires 50 pulses at 345 kHz, resulting in a shot time of 145 μs, while the
PPT mode shot time is 158 μs. In the remaining time before the end of the shot, the
igniter capacitors charge up again and a smaller arc occurs at about 135 μs, following the
sawtooth pattern recorded in some 8:255 shots. In other words, elements of both the
8:120 and 8:255 transformer show up in the 8:185 data. The PPT mode trace in the upper
panel shows a more chaotic signal that hovers closer to zero. This may be interpreted as
repeated, low current arcing across the electrodes that ablate very little mass.
The lower panel in Figure 5-10 shows late-time plasma waveforms collected with the
planar probe with a representative LTPP trace from the 8:120 μPIT for comparison. The
data in the upper and lower panels with matching colors are collected from the same shot.
The timeframe in which the low current PIT mode appeared to be generating late-time
plasma (20-40 μs) is off-scale in the lower panel because the arc plasma is still flowing
past the planar probe. The arc plasma signal swamps the LTTP signal, if one is present.
Figure 5-10: LTPP comparison between the 8:185 and 8:120 μPIT versions
83
However, this indicates that the plasma generated by the antenna may have a bulk
velocity on the order of the arc plasma’s. Another feature to note in the lower panel is the
greater axial extent of the arc plasma for the 8:185 μPIT. This situation could arise from
greater arc energy in the 8:185 than the 8:120; the antenna coupling with late-time
neutrals is worse perhaps because of the increased voltage in the ignition system of the
8:185. While more neutrals are likely generated by the higher 8:185 arc energy, the
elevated voltage applied across the electrodes appears to be too high for the Teflon to
stand off, resulting in repeated arcs instead of the low-level DC discharge seen in the
8:120 density data.
Figure 5-11 shows several images of PIT mode operation captured in 20μs exposures
using the CoolSnap fast-framing CCD camera. The listed exposure times are accurate to
Figure 5-11: Time-resolved images of the 8:185 transformer in PIT mode. Exposure times are in μs.
1-21 8-28 16-36 25-45 32-52
45-65 52-72 66-86 71-91 82-102
84
±6 μs and 𝑡 = 0 is when breakdown occurs. Since the camera cannot capture several full-
resolution images in quick succession, each of these images was taken on a different shot.
The images show blue and orange emissions between 0 and 100 μs that peak in intensity
around 25 μs. The presence of light emission after the arc plasma was ejected would
suggest sustained LTPP and these images were taken to check this possibility. However,
the lack of sustained light emission after about 100 μs indicates that any LTPP is either
not occurring at all or at such a low level as to be inconsequential. These images seem to
confirm the results of the Langmuir data which indicate that no sustained LTPP occurs in
the 8:185 PIT-mode plume.
Another element of the discharge revealed by these photographs is the occasional orange
streak. These streaks are believed to be the trails left by macroparticles (reported by
Spanjers et al. [1998]) as they are ejected from the Teflon or electrode surfaces.
However, based on the exposure time and the streak lengths of around 2 cm, the
minimum velocity for these particles is on the order of 1 km/s. Conversely, Spanjers
reported macroparticle velocities of around 300 m/s and estimated that macroparticles
accounted for up to 40% of the shot mass in the AFRL μPPT. With higher exhaust
velocities, the negative impact of macroparticles upon 𝐼𝑠𝑝 and efficiency is reduced,
though the reason for higher velocity macroparticles ejected from μPIT is not clear.
A set of fast images were collected from the 8:185 μPIT in PPT mode as well. The
images in Figure 5-12 were captured in 20 μs exposures from different shots like the
images in Figure 5-11. These images are much brighter than their PIT mode counterparts
because the antenna blocks much of the light near the electrodes during operation in PIT
mode. Although the net emission in PPT mode is much greater, higher emission intensity
in the region just downstream of the quartz tube exit plane is evident in Figure 5-11,
suggesting higher plasma and neutral gas densities in PIT mode. The PPT configuration
allows for easy viewing of the Teflon and electrode surfaces during breakdown. Many
streaks are visible early in the shot, which are quickly overwhelmed by emission from the
arc plasma in the 0-20 μs frame. The emission becomes more uniform and acquires a
domed shape that peaks on axis. The dome shape implies lower energy at the edges due
85
to losses at the walls. The emission intensity begins decreasing around 70-80 μs,
revealing additional streaking most likely due to macroparticles. These images indicate
that many macroparticles are ejected from the Teflon and electrodes throughout every
shot and that only a small fraction of these exit the thruster with any visible emission.
Streaks are even visible in the last frame, more than 100 μs after the discharge began. The
images also show that the arc position changes from shot to shot, encouraging even
propellant utilization.
Figure 5-12: Time-resolved images of the 8:185 transformer in PPT mode. Times are in μs.
0-20 25-45
50-70
12-32 45-65
86-106
-4 - 16
111-131 126-146
86
Figure 5-13: Quartz tube with residue after PPT mode testing
Figure 5-13 is a photograph of the quartz tube used to test μPIT in PPT mode. Note the
dark residue distributed symmetrically around the tube. The electrodes and Teflon face
were positioned about a third of the way down the tube, where the deposits begin. The
symmetry of the deposits is further evidence that the arc position shifts from shot to shot,
resulting in azimuthally uniform Teflon ablation. The residue is most likely composed of
carbon, fluorine, and copper that arrives on the quartz by a combination of macroparticle
ejection as seen in Figure 5-12 and vapor deposition.
5.3.2 Arc Plasma Density Results
The fast-framing CCD camera was also used to capture images of entire shots from the
8:255 and 8:185 iterations of μPIT. The shot lengths for each of the images in Figure
5-14 range from 145-160 μs. Each image was taken with a 200 μs exposure time in order
to capture each shot in its entirety. Figure 5-14 (a) shows a typical shot of the 8:255
transformer, while (b) shows a representative 8:185 shot in PIT mode. The obvious
difference between the two is the greater quantity of light and more extensive plume in
the 8:185 image. Since plasma density is roughly proportional to arc energy, this result is
consistent with the idea that the 8:255 breaks down with significantly less energy than the
8:185. The third image (c) shows a representative PPT mode shot. The plasma inside the
quartz tube is extremely bright, but emission from the region immediately downstream of
the quartz tube exit plane approximately matches the emission from the middle image.
87
Figure 5-14: Light emission from (a) 8:255 μPIT, (b) 8:185 μPIT, PIT mode, (c) 8:185 μPIT, PPT
mode
However, the lack of obvious differences between these images does not rule out the
possibility that the antenna couples significantly with the arc plasma.
Also noteworthy in Figure 5-14 is a certain measure of asymmetry in all three plumes.
The emission asymmetry is attributable to the orange regions in each plume. Based on the
streaks visible in the 20 μs exposures, these orange regions are likely caused by several
macroparticles being ejected along similar paths.
(a) (b) (c)
88
Figure 5-15: PIT mode radial density profile at z = 5 cm downstream
Figure 5-16: PIT mode radial density profile at z = 15 cm downstream
Figure 5-15 and Figure 5-16 show radial plasma density profiles that were collected with
the planar Langmuir probe at 5 and 15 cm downstream of the antenna exit plane. 2-D
representations are also shown to clarify the surface plots. The high current PIT
configuration was used to collect this data. This data was collected at 7-9 radial locations;
density in the regions between these stations was interpolated using a cubic spline,
resulting in smooth profiles. The density profile at each (𝑟, 𝑧) coordinate is averaged over
three shots whose breakdown times were all shifted to zero. One trend evident in these
figures is the decrease in density with 𝑟. The density falls off rapidly with radius at 𝑧 = 5
cm, while plasma density decreases more slowly in the 15 cm downstream data. The
density also decreases with 𝑧, falling by an order of magnitude between the 5 and 15 cm
89
downstream stations. These inverse relationships with 𝑟 and 𝑧 are consistent with an
initially collimated plume that expands radially as it travels downstream.
The 2-D plots reveal a slight thrust vector misalignment. The degree of plume collimation
can be quantified by the angle between the plume axis and the radial location where
density falls to 10% of the on-axis value. 5 cm downstream, the plume half angle is 30 ̊,
while at 15 cm the half angle is 23 ̊. These half angles match reasonably well, considering
errors introduced by shot-to-shot density variations. The radial density profiles for the
PPT configuration were also measured and the results are presented below.
Radial PPT mode density profiles at 5 and 15 cm downstream are shown in Figure 5-17
and Figure 5-18, respectively. These plots were constructed in a way similar to that used
for the PIT mode figures, with one exception. The 5 cm data in Figure 5-17 was
Figure 5-17: PPT mode radial density profile at z = 5 cm downstream
Figure 5-18: PPT mode radial density profile at z = 15 cm downstream
90
interpolated linearly instead of with a spline, which led to an unrealistic plume shape due
to slope matching requirements. Like the PIT mode profiles, these show an order of
magnitude drop in peak density between 5 and 15 cm, corresponding to the radial
expansion of the plume.
Also evident in Figure 5-18 is a significant misalignment of the thrust axis. After
pumping down to vacuum, the quartz tube used to match the wall conditions of PIT mode
was visibly askew as some Kapton tape had peeled away. The skewed quartz can explain
this axis misalignment in the 15 cm data, but the 5 cm data appears to be well aligned
with the axis. Since the density trace at each radial location is only averaged over three
shots, the lack of any misalignment in the 5 cm PPT data could be explained by a few
successive shots with above average energy. The 5 cm plume half angle is about 22 ̊,
while 15 cm data indicates a half angle of 27 ̊ with no correction for axis alignment. With
the correction, the half angle is 17 ̊, so either way the half angles for each PPT mode
location are roughly in agreement. The approximate equivalence among the half angles
estimated for the PIT and PPT configurations suggests that the antenna does not have a
significant effect on the degree of plume collimation. It was hypothesized that an induced
azimuthal current density crossed with the axial magnetic field would radially compress
the plasma, potentially leading to a greater degree of collimation and limiting contact
with the walls. However, the lack of a reduced half angle in the PIT mode data suggests
that μPIT is not driving a significant azimuthal current. In fact, the residue found on the
PPT mode quartz tube (shown in Figure 5-13) indicates that the physical barrier
presented by the quartz, rather than the antenna, dictates the plume half angle.
The 5 cm surface plots in both PIT and PPT mode show a sharp transition from zero
density to peak values while the 15 cm data for each mode shows a more gradual rise to
the peak density. This difference stems from the plume’s axial expansion with time due
to the spread out velocity distribution of the arc plasma.
The peak density at 5 cm for PIT mode is about 3 × 1017 𝑚−3 while the peak density in
PPT mode at 5 cm is half that. Similarly, the peak density in PIT mode at 15 cm is
7 × 1016 𝑚−3 while the peak density in PPT mode is just 4 × 1017 𝑚−3. These
91
differences in peak density strongly suggest that the antenna is coupling directly with the
plasma, heating it and increasing the ionization fraction.
A feature that distinguishes the PIT mode density profile at 15 cm from the other three is
the waving pattern visible in Figure 5-16. The wave frequency matches the antenna
frequency and is only visible in 15 cm data, so this effect is clearly related to the RF
power supply. Since the pattern is much more pronounced in the PIT mode data, it is
concluded that the antenna is primarily responsible for the density waves. However, the
oscillating electrodes must also make a contribution as some PPT mode data at 15 cm
also shows a low-level wave pattern. These waves are only visible for early breakdown
times such that plasma is in the antenna when it rings up to its peak current, which occurs
about 15 μs after switching on. Due to greater variation in breakdown time in PPT mode
(discussed further in Section 5.3.3), sufficiently early breakdown times to capture the
wave pattern are uncommon. The averaging process used to generate the radial density
profiles tends to wash out this effect since most of the waveforms contain no
recognizable waving pattern. The end result is that very little of this density wave pattern
appears in the 15 cm PPT mode density profile.
The waving signal is not visible in the 5 cm data because the arc plasma density is
sufficiently large to wash out this relatively small feature. Two possible mechanisms
could create this waving signal. The waving pattern could be real density waves driven in
the arc plasma by the antenna, accelerated in pulses by 𝐽𝜃 × 𝐵𝑟 . Alternatively, it could be
purely EMI, resulting from the electrodes and antenna broadcasting fields and driving an
additional oscillating current in the planar Langmuir probe.
92
Figure 5-19: Axial density profile comparing the 8:185 PIT and PPT modes
Figure 5-19 shows the on-axis peak plasma density profiles for the high current PIT and
PPT modes between 5 and 18 cm downstream. Like the radial density profiles, three
shots were collected at each data point and averaged to form the plot. Due to the scatter
inherent in the operation of PPTs, exponentially decaying curve fits were applied to the
data. In addition, the averaged data points suggest higher peak plasma density in PIT
mode, and these curve fits allow the difference in density to be quantified. Based on the
curve fits, operation in PIT mode results in a 10% to 40% higher peak plasma density
than in PPT mode. This density increase is directly attributable to coupling between the
antenna and the arc plasma. Additional differences in performance due to the antenna are
discussed in the following section.
5.3.3 Specific Antenna Effects
As was pointed out earlier, the presence of the antenna has a positive impact on μPIT’s
shot-to-shot reliability. Figure 5-20 shows a histogram of breakdown times that compares
operation in PIT and PPT modes. Each set contains about 100 data points.
93
Figure 5-20: Breakdown time histogram for 8:185 PIT and PPT modes
Antenna shutoff occurred at 145 μs in PIT mode and at 158 μs in PPT mode. The shots in
PPT mode shown to occur at ~170 μs actually represent the total collection of shots that
broke down an unknown time after the antenna shut off or failed to break down at all.
Even with a peak igniter voltage of 7.5 kV, the PPT configuration did not successfully
develop an arc on every shot. About 8% of the shots in PPT mode either fired late or not
at all. Furthermore, the possible range of breakdown times in PPT mode are spread out
over the whole time the antenna is ringing. In contrast, the breakdown time histogram in
PIT mode is much more peaked, with every arc occurring within 42 μs of the time the
antenna first switches on. Additionally, 97% of shots in PIT mode occurred before 33 μs.
Virtually all PIT mode shots develop an arc quickly enough for plasma to enter the
antenna close to the time that it reaches peak current. This histogram illustrates
significant gains in reliability and shot-to-shot repeatability, and these gains are directly
attributable to the antenna.
Many time of flight measurements were performed to find the bulk velocity of the arc
plasma. TOF results for the high and low current PIT modes as well as the PPT
94
Table 5-4: Exhaust velocities at 350 V charge for the three 8:185 operating modes
Configuration PIT (high I) PIT (low I) PPT Exhaust velocity [km/s] 32±3 21±3 20±3
configuration are shown in Table 5-4. Note that all these measurements were collected
with the standard storage capacitor charge of 350 V.
At 20 km/s, the arc plasma velocity in PPT mode is by itself quite high. This velocity
falls within the range measured by other researchers who examined the plumes of
primarily electromagnetic PPTs, suggesting that electromagnetic acceleration dominates
the operation of μPIT as well. With an ion temperature estimated to be on the order of 1
eV, the ion thermal speed is only a few km/s. This ion thermal speed, combined with the
lack of a solid nozzle strongly indicates that μPIT is primarily an electromagnetic
accelerator.
Table 5-4 clearly shows the impact of the antenna on the arc plasma bulk velocity. The
low-current PIT mode evidently has too little current in the antenna to affect the exhaust
velocity at all. On the other hand, operating in the high current PIT mode was found to
increase the bulk velocity of the arc plasma by about 50%, to 32 km/s. The acceleration
mechanism that produces this dramatic increase in exhaust velocity may be either
electrothermal or electromagnetic. In the former case, the antenna may drive currents in
the plasma, transferring energy to the electrons which then redistribute their energy via
collisions. The second possible acceleration mechanism is the same 𝐽𝜃 × 𝐵𝑟 force found
in large scale pulsed inductive thrusters. For the same reasons described earlier, the
electromagnetic mechanism seems more plausible.
In order to improve confidence in the TOF measurements, they were performed in greater
detail for the low current PIT configuration. The bulk velocity of the arc plasma was
measured and averaged over ten shots with probe separations of 5, 10, and 15 cm. The
results are listed in Table 5-5.
95
Table 5-5: Exhaust velocity measurements for the low current PIT configuration at a 350 V charge
Δz [cm] Exhaust velocity [km/s]
5 20±4 10 23±3 15 22±2
Keeping in mind the estimated uncertainty, the table shows that the plume velocity
remains constant downstream of the antenna. These measurements serve to support the
TOF results presented in Table 5-4.
In addition to tests of the different 8:185 configurations at the maximum 350 V storage
capacitor charge, a series of measurements were performed to identify the minimum
breakdown voltage in the PIT and PPT modes. In addition, the data collected allowed for
a characterization of the dependence of exhaust velocity and density upon charge voltage.
These results were collected using the following procedure. The test series began with
one shot at full voltage (350 V). The charge voltage was then reduced to around 100 V
and a fire command was sent to the power supply. This is too low for μPIT to fire in
vacuum in either PIT or PPT mode, and so the charge voltage was increased by about 10
V and another fire command was sent. This process of increasing the charge and
attempting to fire was repeated until μPIT fired successfully, yielding an approximate
knowledge of the minimum required breakdown voltage for each shot. Then μPIT was
fired at 350 V in an effort to reset the system and minimize any carbon deposition on the
thruster face due to the low discharge energy. This process of alternating between the
maximum and minimum charge voltages was repeated about 25 times in high current PIT
mode and again in PPT mode.
96
Figure 5-21: Minimum voltage required for breakdown for the PPT and high current PIT
configurations
A comparison between PIT and PPT mode minimum breakdown voltages is shown in
Figure 5-21. An obvious trend is the reduced voltage required for breakdown in PIT
mode. As described in Section 4.1, the radial component of the magnetic field near the
thruster face provides a path for electrons, effectively increasing the radial conductivity.
The average igniter voltage required for an arc in PPT mode is 4.0 kV, while the average
igniter voltage required in high-current PIT mode is just 2.7 kV, a 30% reduction.
Reduced breakdown voltage could be an advantage for PPTs, particularly those like μPIT
that initiate the main discharge with high voltage instead of a separate igniter.
Specifically, a reduced voltage requirement could lead to mass savings in shielding,
insulation, and power supply mass.
Figure 5-21 also demonstrates that the average breakdown voltage remains relatively
constant throughout the test series. There was a concern based on work by Keidar, et al.
[2004] that the very low discharge energy associated with the minimum breakdown
voltage could lead to carbon deposits on the Teflon surface. Experience with the AFRL
μPPT showed that such deposits can increase the electrical conductivity of the Teflon
97
Figure 5-22: TOF-derived exhaust velocities at minimum breakdown voltages for the PIT and PPT
modes
face and ultimately short-circuit the PPT. However, this plot of sequential firings shows
that at least on the scale of tens of shots, the average breakdown voltage does not change
significantly.
Several exhaust velocity measurements from the minimum breakdown voltage test series
are shown in Figure 5-22. For all of these shots, TOF was measured with a separation of
15 cm between the symmetric double and planar Langmuir probes. As was the case for
the shots with a 350 V charge, the exhaust velocity in PIT mode at minimum voltage
tends to be much greater than in PPT mode. The average exhaust velocity is 28±2 km/s in
PIT mode and 20±2 km/s in PPT mode. These figures approximately match the full-
voltage results listed in Table 5-4. Also note that exhaust velocity does not appear to
increase with charge voltage, but remains relatively flat. In the case of PPT mode,
exhaust velocity may even decrease slightly with increasing charge voltage. The fact that
velocity is rough independent of charge voltage suggests that any additional energy
98
Figure 5-23: Peak plasma density dependence on charge voltage for PPT and high current PIT
modes
transferred to the plasma at 350 V charge, either inductively or directly through a quasi-
DC arc, does not go toward increasing the bulk velocity.
The method of alternating between maximum and minimum igniter voltage also allowed
for a continuous check on the full charge exhaust velocity to see if it remained the same
as the figures reported in Table 5-4. The TOF measurements at a charge of 350 V yielded
exhaust velocities that match the results reported earlier to less than 1 km/s.
Figure 5-23 shows how the peak plasma density changes with charge voltage. The lower
breakdown voltage requirement of the PIT configuration is reflected in this plot by the
absence of any PPT mode data around the 100 V level. For a fixed charge voltage, the
peak density in PIT mode is generally higher than the corresponding PPT mode peak
density. The average increase in density for a given charge voltage in PIT mode is 31%,
which falls within the 10-40% range predicted from the peak density curve fits in Figure
5-19.
99
Taken together, the exhaust velocity and peak density dependencies on charge voltage
show that the increased energy associated with higher charge voltage does not
significantly affect the exhaust velocity, but it does show up as increased plasma density.
As was mentioned earlier, the usual linear scaling between energy and peak density that
appears in larger PPTs is not appropriate for the tens of mJ energy range encountered in
the operation of μPIT. Instead, Figure 5-23 shows that doubling the charge voltage leads
to a little over twice the peak density, resulting in a scaling with energy like 𝑛 ∝ 𝐸1/2.
5.3.4 μPIT Performance Estimates
An estimated impulse bit for the PPT and high current PIT configurations were calculated
based on Equation (4.24). The function 𝑛(𝑟, 𝑡) was measured at 5 cm and 15 cm
downstream for both configurations, shown in Figure 5-15 through Figure 5-18. The
plume was assumed to be axisymmetric, but the radial density profiles showed the PIT
mode plume was slightly skewed, while the PPT mode plume was misaligned more
severely. The effect of axis misalignment was corrected for by adjusting the bounds in
integration over 𝑑𝑟 such that integration began at the location of peak density, assumed to
be the true axis. This method does not account for out-of-plane alignment error; however,
axis misalignment was very small for the PIT mode data and in PPT mode, the quartz
tube was visibly tilted to one side such that the radial Langmuir probe measurements
were roughly in-plane. The resulting impulse bit estimates are listed in Table 5-6.
As indicated by the table, μPIT can generate extremely small impulse bits appropriate for
very fine pointing maneuvers. Although these impulse bits may be too small for some
current microspacecraft requirements, μPIT can easily be scaled up to a larger electrode
gap and arc energy, generating higher thrust. Also, the pulsed nature of the device could
allow for operation at several Hz, given appropriate thermal management and enough
power. It is important to note that these impulse bit calculations are underestimates since
the momentum contribution from neutral particles is taken to be zero. However, the
Table 5-6: Estimated impulse bits for PPT and high current PIT modes at 350 V charge
Mode 𝑰𝒃𝒊𝒕 at 5 cm [nN-s] 𝑰𝒃𝒊𝒕 at 15 cm [nN-s] PIT 54 120 PPT 35 90
100
velocity of these neutrals is thought to be on the order of 1 km/s, so the correction is
probably small.
Large discrepancies exist between the impulse bits calculated at 5 cm and at 15 cm. It
seems likely that the measurements 5 cm downstream missed a significant fraction of the
total shot mass. For example, Figure 5-17 shows that the density had not dropped to zero
at the maximum r coordinate where data was collected. The broader range of the 15 cm
measurements and the improved resolution probably contributed to a better estimate of
the impulse bit.
Despite these discrepancies, Table 5-6 offers an order of magnitude estimate of the
impulse bit generated by μPIT. More importantly, these estimates demonstrate that the
RF antenna drives up the impulse bit by 30-50%. This larger impulse bit is due in part to
the increased plasma density generated by the antenna, but the majority of this increase
comes from the greater exhaust velocity imparted by the antenna.
The increased impulse bit in PIT mode comes at the cost of additional energy drawn from
the storage capacitor. The total energy drawn from the storage was calculated using data
from a voltage divider. Figure 5-24 shows a typical storage capacitor voltage trace along
with a planar probe density trace for comparison. The voltage begins at 350 V, decreases
rapidly after the antenna is switched on at 𝑡 = 0. Breakdown occurs at about 10 μs and
the plasma density peak flows past a Langmuir probe 5 cm downstream before 20 μs. The
storage capacitor voltage trace changes slope at about 20 μs. At earlier times, the power
supply is ringing up the antenna, which is charging the igniter capacitors, then coupling
power directly into the plasma. The change in slope is believed to signal an end to
significant coupling with the plasma, so this point is used to define the ending voltage in
the shot input energy calculation. Current continues in the antenna for until the switch off
time at 145 μs, but the energy expended does not appear to perform any useful function
with respect to LTPP. As a result, the most efficient mode of operation for the 8:185
μPIT seems to be to concentrate on adding energy to the arc plasma, then to shut off once
the arc plasma is ejected.
101
Figure 5-24: Representative storage capacitor voltage waveform from a high current PIT mode shot
Table 5-7 shows the shot efficiency and energy. The shot energy was calculated based on
the arguments above for eight shots each in PIT and PPT mode. The averaged results for
each configuration are listed in the table. The shot efficiency is calculated using Equation
(4.25), which accounts for all the energy losses between the RF power supply the kinetic
energy in the plume. In order to find the overall efficiency, the shot efficiency should be
multiplied by the electrical efficiency of the 350+ V supply used to charge the storage
capacitors, which should be quite high in any case (perhaps 90%).
The shot energies listed in Table 5-7 are about the same, but the efficiency in PIT mode is
double that in PPT mode. These results indicated that some portion of the energy stored
in the tank circuit’s magnetic field is transferred to the plasma in PIT mode, while in PPT
mode the stored energy is resistively dissipated in the coils as the field decays away.
Table 5-7: Energy and shot efficiency comparison for PIT and PPT modes
Mode Shot Energy [J] Efficiency PIT 3.3 0.11 % PPT 3.4 0.050 %
102
While these efficiencies are quite low compare to flight PPTs, this design is a prototype
lacking optimization. Additionally, the doubled efficiency in PIT mode illustrates the
potential of this inductive coupling strategy to surpass standard PPT performance when
better coupling is achieved.
103
6 Summary and Recommendations
Pulsed plasma thrusters have several advantages derived from their solid fuel and pulsed
operation including easy vehicle integration, low dry mass, and the ability to vary thrust
at constant specific impulse. PPTs can bring these advantages to bear in the growing field
of micropropulsion, but their low efficiency is a significant drawback. A large neutral
population is generated by each shot and very little of this neutral gas can be ionized and
accelerated by the arc, limiting efficiency. The μPIT prototype was constructed with the
goal of improving upon state of the art PPT technology by inductively ionizing and
heating into these neutrals, enhancing mass utilization and overall thruster efficiency. The
specific goals of this thesis were to develop a μPIT prototype that would reliably break
down, to examine the effects of the antenna on the arc plasma and late-time neutrals, and
to estimate the performance parameters of the prototype thruster.
An early iteration of μPIT using an 8:120 transformer demonstrated late-time plasma
production. After the arc plasma passed by, two separate Langmuir probes measured a
low-level DC discharge for the duration of the shot. The signal decayed to zero shortly
after the antenna switched off, demonstrating a strong correlation with antenna operation.
The late-time plasma density was about two orders of magnitude lower than the arc
plasma at the thruster exit plane; ionization rate estimates suggest that a higher energy
discharge would increase the density of the late-time plasma.
A subsequent transformer design applied a peak voltage of 10.5 kV across the electrodes.
This 8:255 μPIT demonstrated repeatable breakdown but suffered from low energy
discharges and power supply arcs. An 8:185 transformer later demonstrated a high level
of reliability without the severe drawbacks of the 8:255 transformer. Additionally,
presence of the antenna was found to reduce the minimum breakdown voltage. Radial
magnetic field components from the antenna provide a lower resistance path for
electrons, reducing the voltage requirement.
The radial and axial structure of the 8:185 plume was mapped out with Langmuir probes
and the arc plasma bulk velocity was measured using time of flight. Additionally, a fast-
framing camera was used to capture 20 μs exposures of the plume as it evolved in time.
104
These density and velocity measurements were used to estimate the impulse bit. The
calculation yielded an impulse bit of approximately 0.1 μN-s. Firing at 1 Hz, the 8:185
μPIT would consume about 3 W.
Some effects specifically attributable to the antenna were investigated by comparing
density and velocity measurements with and without the antenna located in a position that
enabled inductive coupling with the plasma. The most striking result of these tests was
the increased exhaust velocity. In PPT mode, the bulk plasma velocity was about 20
km/s, but it was repeatedly measured at 32 km/s in PIT mode. The high exhaust velocities
indicate that μPIT accelerates propellant electromagnetically. In addition, the PIT
configuration increased the peak plasma density over the PPT mode by about 30%. A
series of experiments involving dependencies on charge voltage showed that exhaust
velocity was roughly independent of the storage capacitor charge, while the peak plasma
density increased linearly with charge voltage. The additional energy associated with
increased charge voltage shows up in the plume not as higher exhaust velocity but as a
larger quantity of plasma.
An important next step for μPIT is a direct measurement of the impulse bit, which will be
difficult due to the extremely low forces involved. The method employed by Emhoff et
al. [2007] may be useful here.
The inductive coupling also needs to be improved in order for μPIT to outperform the
efficiency of flight-model PPTs. Measurements by Spanjers [1996] and Koizumi [2007]
indicate that a great deal of neutral particles are present around the plume and for
hundreds of microseconds after the arc. The ionization rate estimates in Chapter 4 show
an 𝑛2 scaling, indicating that a scaled-up version of μPIT (perhaps macroPIT?) could
develop a much more substantial late-time plasma. Increased antenna current would also
help improve inductive coupling due to the 𝐼2 power transfer scaling.
A study of μPIT performance at higher frequencies would likely yield interesting results.
For example, the higher B would stronger electric fields which drive 𝐽𝜃 . Increased
azimuthal current density is valuable from both from an acceleration and a heating
standpoint. In addition, a strong 𝐽𝜃 × 𝐵𝑟 force inside the antenna would compress the
105
plasma, reducing wall losses and improving plume collimation. More rapid magnetic
field inflation would improve the antenna impedance matching, providing an increase in
the arc plasma exhaust velocity.
The development effort so far has shown μPIT to be a promising micropropulsion
concept. The increased arc plasma velocity, higher peak plasma density, and the late-time
plasma are new contributions to PPT technology. As this and other micropropulsion
projects continue forward, the advantages of spacecraft formation flight will be more
fully realized. As an enabling technology, micropropulsion will help make space systems
more reliable, more flexible, and less expensive. The inductive coupling employed in
μPIT also suggests future work in the field of larger-scale PPTs.
106
Bibliography
Air Force Research Laboratory. ―XSS-11 Micro-Satellite Fact Sheet.‖ 2005.
Antonsen, ―Herriott Cell Interferometer for Unsteady Density Measurements in Small
PPT Plasmas.‖ 36th AlAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,
July 16-19, 2000, Huntsville, AL. AIAA paper 2000-3431.
Bartone, M. ―A Gas-fed Micro Pulsed Inductive Thruster.‖ MS Thesis, Dept. of
Aeronautics and Astronautics, Univerisity of Washington. 2007.
Bohm, D. Characteristics of Electrical Discharges in Magnetic Fields. New York,
McGraw-Hill, 1949.
Burke, J.D. ―Micro-Spacecraft‖ Jet Propulsion Lab. Publication 715-87, Pasadena, CA.
15 Oct. 1981.
Burton, R. L. and S. S. Bushman. ―Probe Measurements in a Coaxial Gasdynamic PPT.‖
35th ‘AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. June 20,23,1999,
Los Angeles, CA. AIAA paper 99-2288.
Burton. R. L, F. Rysanek, E. Atonsen, M. Wilson, and S. Bushman. ―Pulsed Plasma
Thruster Performance for Microspacecraft Propulsion.‖ Progress in Astronautics and
Aeronautics. AIAA Vol. 187. 2000. p 337-352.
Burton, R. L, M. J. Wilson, and S. S. Bushman. ―Energy Balance and Efficiency of the
Pulsed Plasma Thruster.‖ AIAA/ASME/SAE/ASEE Joint Propulsion Conference and
Exhibit, 34th, Cleveland, OH, July 13-15. AIAA paper 1998-3808.
Burton, R. L. and P. J. Turchi. ―Pulsed Plasma Thruster.‖ Journal of Propulsion and
Power. 14, 5, 1998.
Busek Co. ―Colloid Thruster Systems.‖ 2007a.
Busek Co. ―Micro Pulsed Plasma Thrusters.‖ 2007b.
Bushman, S. and R. L Burton. ―Heating and Plasma Properties in a Gasdynamic PPT.‖
Journal of Propulsion and Power, 17, 5, 2001.
Byrne, L. T. ―Langmuir Probe Measurements in the Plume of a Pulsed Plasma Thruster.‖
MS Thesis. Dept. of Mechanical Engineering, Worchester Polytechnic Institute, 2002.
Cardiff, E. H. and G. C. Marr. ―Propulsion Options for the LISA Mission.‖ 40th
AIAA
Joint Propulsion Conference, Fort Lauderdale, FL, 2004. AIAA-2004-3340.
Chung, P. M, L. Talbot, K. J. Touryan. Electric Probes in Stationary and Flowing
Plasmas: Theory and Application. Springer-Verlag, New York, 1975.
Dept. of Defense. RDT&E Budget Item Justification Sheet, Space Programs and
Technology. 2007.
107
Dipprey, N. F. and S. J. Rotenberger. ―Orbital Express Propellant Resupply Servicing.‖
39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. 20-23 July 2003,
Huntsville, Alabama. AIAA paper 2003-4898.
Ebert, W. L, S. J. Kowel, R. F. Sloan. ―Operational NOVA Spacecraft Teflon Pulsed
Plasma Thruster System.‖ ASME, SAE, and ASEE, Joint Propulsion Conference, 25th,
Monterey, CA, July 10-13, 1989. AIAA paper 1989-2497.
Eckman, R, N. A. Gatsonis, R. M. Meyers, E. J. Pencil. Experimental Investigation of the
LES 8/9 Pulsed Plasma Thruster Plume. 25th International Electric Propulsion
Conference, Oct, 1997, Cleveland, Ohio. 797.
Emhoff, J. W, D. H. Simon, H, B. Land ―Progress in Thrust Measurements from Micro
Pulsed Plasma Discharges.‖ 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference
& Exhibit, 8 - 11 July 2007, Cincinnati, OH. AIAA paper 2007-5299.
Gatsonis, N. A, J. Zwahlen, A. Wheelock, E. Pencil, H. Kamhawi. ―Pulsed Plasma
Thruster Plume Investigation Using a Current-Mode Quadruple Probe Method.‖ Journal
of Propulsion and Power, 20, 2, 2004.
Genovese, A., W. Steiger, and M. Tajmar. ―Experimental Characterization of Indium
FEEPs.‖ 2001. AIAA paper 2001-3788.
Hirata, M. and H. Murakami. ―Impulse Measurement of a Pulsed Plasma Engine on
Engineering Test Satellite IV.‖ 16th International Electric Propulsion Conference, New
Orleans, LA, Nov 17-19, 1982. AIAA paper 1982-1875.
Hutchinson, I. Principles of Plasma Diagnostics. Cambridge University Press, New
York, 1987.
Jahn R.G, Physics of Electric Propulsion, McGraw-Hill, New York, 1968.
Johonson, B. H, and D. L. Murphree. ―Plasma Velocity Determination by Electrostatic
Probes.‖ AIAA Journal, 7, 10, 1969.
Johnson E. O. and L. Malter. ―A Floating Double Probe Method for Measurements in
Gas Discharges.‖ Physical Review, 80, 1, 1950.
Kanal, M. ―Theory of Current Collection of Moving Cylindrical Probes.‖ Journal of
Applied Physics, 35, 6, 1964.
Katz, I. ―Electric Propulsion for JPL Missions.‖ 41st AIAA/ASME/SAE/ASEE Joint
Propulsion Conference & Exhibit. 10 - 13 July 2005, Tucson, Arizona. AIAA paper
2005-3674.
Keidar, M, I. D. Boyd, E. L. Antonsen, F. S. Gulczinski, G. G. Spanjers. ―Propellant
Charring in MicroPPTs.‖ Journal of Propulsion and Power, 20, 6, 2004.
Keidar, M, I. D. Boyd, E. Antonsen, R. Burton, and G. Spanjers. ―Optimization Issues for
a Micropulsed Plasma Thruster.‖ Journal of Propulsion and Power, 22, 1, 2006.
108
Ketsdever, A. D. ―System Considerations and Design Options for Microspacecraft
Propulsion Systems.‖ Progress in Astronautics and Aeronautics. AIAA Vol. 187. 2000. p
139-163.
Khayms, V. and Martinez-Sanchez, M. ―50 Watt Hall Thruster for Micropropulsion.‖
Progress in Astronautics and Aeronautics. AIAA Vol. 187. 2000. p 233-254.
Koizumi, H, R. Noji, K. Komurasaki, and Y. Arakawa. ―Plasma Acceleration Processes
in an Ablative Pulsed Plasma Thruster.‖ Physics of Plasmas. 14, 033506. 2007.
Kristiansen, R. ―Dynamic Synchronization of Spacecraft: Modeling and Coordinated
Control of Leader-Follower Spacecraft Formations.‖ PhD Thesis. Dept. of Engineering
Cybernetics, Norwegian University of Science and Technology. 2008
Laframboise, J. G. Theory of Cylindrical and Spherical Langmuir Probes in a
Collisionless, Maxwellian Plasma at Rest. UTIAS Report, 1966.
Laure, S, J. Heiermann, M. Aueter-Kurtz, H. Kurtz, ―ATILLA—Adjustable Throttle
Inductively Afterburning Arcjet.‖ 27th IEPC, Oct, 2001, Pasadena, California, USA.
IEPC 01-142.
Lieberman, M.A, and A.J. Lichtenberg. Principles of Plasma Discharges and Materials
Processing. John Wiley & Sons, New York, 1994.
Mitterauer, J. ―Indium: An Alternative Propellant for FEEP-Thrusters.‖ 37th
AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. 8-11 July 2001 Salt
Lake City, Utah, USA. AIAA paper 2001-3792.
Mueller, J. "Thruster Options for Microspacecraft: A Review and Evaluation of State-of-
the-Art and Emerging Technologies." Progress in Astronautics and Aeronautics. AIAA
Vol. 187. 2000. p 45-138.
Rayburn, C. D, M. E. Campbell, and A. T. Mattick. ―Pulsed Plasma Thruster System for
Microsatellites.‖ Journal of Spacecraft and Rockets, 42, 1, 2005.
Ruchti, C. B, and L. Niemeyer. ―Ablation Controlled Arcs.‖ IEEE Transactions on
Plasma Science. 14, 4, 1986.
Russel, D, J. Poylio, W. Goldstein, B. Jackson, R. Lovberg, and C. Dailey. ―The Mark VI
Pulsed Inductive Thruster.‖ Space 2004 Conference and Exhibit, 28-30 Sept. 2004, San
Diego, CA. AIAA paper 2004-6054.
Rysanek, F. and R. L. Burton. ―Effects of Geometry and Energy on a Coaxial
Gasdynamic PPT.‖ 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and
Exhibit, July 16-19, 2000. Huntsville, AL. AIAA paper 2000-3429.
Scharlemann, C. and M. Tajmar. ―Development of Propulsion Means for Microsatellites‖
43rd
Joint Propulsion Conference & Exhibit. 8-11 July 2007, Cincinnati, Ohio.
109
Schilling, J. H, R.A. Spores, and G. G. Spanjers. ―Micropropulsion Options for
Techsat21 Space-Based Radar Flight.‖ Progress in Astronautics and Aeronautics. AIAA
Vol. 187. 2000. p 3-23.
Shumlak, U, and T. Maruo. ―Plasma Plume Mass Characterization of a Mini-Pulsed
Plasma Thruster.‖ 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit,
9 - 12 July 2006, Sacramento, California. AIAA paper 2006-4333.
Simon, D. H. and H. B. Land III. ―Micro Pulsed Plasma Thruster Technology
Development.‖ 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,
11 - 14 July 2004, Fort Lauderdale, Florida. AIAA paper 2004-3522.
Spanjers, G. G, K. A. McFall, F. S. Gulczinski III, R. A. Spores. ―Investigation of
Propellant Inefficiencies in a Pulsed Plasma Thruster.‖ 32nd AIAA, ASME, SAE, and
ASEE, Joint Propulsion Conference and Exhibit., Lake Buena Vista, FL, July 1-3, 1996.
AIAA paper 96-2723.
Spanjers, G. G, J. S. Lotspeich, K. A. McFall, R. A. Spores. ―Propellant Losses Because
of Particulate Emission in a Pulsed Plasma Thruster.‖ Journal of Propulsion and Power,
14, 4, 1998.
Spanjers, G. G, E. L. Antonsen, M. Keidar, S. Bushman. ―Advanced Diagnostics for
Millimeter Scale MicroPPTs.‖ 33rd Plasmadynamics and Lasers Conference, 20-23 May
2002, Maui, Hawaii. AIAA paper 2002-2119.
Spores, R. A, M. Birkan. ―The USAF Electric Propulsion Program.‖ 35th
AIAA/ASME/ASEE Joint Propulsion Conference and Exhibit, 20-24 June 1999 Los
Angeles, CA. AIAA paper 1999-2162.
Stebbins, R. T. ―Recent Progress at NASA in LISA Formulation and Technology
Development.‖ Classical and Quantum Gravity. 25, 11. 2008.
Tajmar, M., A. Genovese, and W. Steiger. ―Indium Field Emission Electric Propulsion
Microthruster Experimental Characterization.‖ Journal of Propulsion and Power, 20,
211. 2004.
Thomassen, K. I. and R. J. Vondra. ‖Exhaust Velocity Studies of a Solid Teflon Pulsed
Plasma Thruster.‖ Journal of Spacecraft and Rockets, 9, 61, 1972.
Tilley, D. L, A. J. Kelly, and R. G. Jahn. ―The Application of the Triple Probe Method to
the MPD Thruster.‖ 21st Internationational Electric Propulsion Conference, July 18-20,
1990, Orlando, FL. AIAA paper 1990-2667.
Turchi, P. J, I. G. Mikellides, P. G. Mikellides, and H. Kamhawi. ―Pulsed Plasma
Thrusters for Microsatellite Propulsion: Techniques for Optimization.‖ Progress in
Astronautics and Aeronautics. AIAA Vol. 187. 2000. p 353-368.
Velasqez-Garcia, L. F, M. Martinez-Sanchez, and A. I. Akinwande. ―Two-Dimensional
Microfabricated Colloid Thruster Arrays.‖ 40th AIAA/ASME/SAE/ASEE Joint Propulsion
110
Conference and Exhibit, 11 - 14 July 2004, Fort Lauderdale, Florida. AIAA paper 2004-
3595.
White, D, J. H. Schilling, S. Bushman, G. G. Spanjers, D. R. Bromaghim, J. Lake, M.
Dulligan. ―AFRL MicroPPT Development for Small Spacecraft Propulsion.‖ 33rd
Plasmadynamics and Lasers Conference, 20-23 May 2002, Maui, Hawaii. AIAA paper
2002-2120.
Zakrzwski, C, S. Benson, P. Sanneman, and A. Hoskins. ―On-Orbit Testing of the EO-1
Pulsed Plasma Thruster.‖ 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &
Exhibit, 7-10 July 2002, Indianapolis, Indiana. AIAA paper 2002-3973.
Zakrzwski, C, S. Benson, ―Section 25: Pulsed Plasma Thruster.‖ EO-1 Mission
Technology Forum. 2001
Zeuner, M, F. Scholze, H. J. Leiter, R. Kukies, D. Feili, M. Tartz, and H. Neumann. ―Ion
Beam Characteristics of the RIT 10 Ion Thruster.‖ 39th AIAA/ASME/SAE/ASEE Joint
Propulsion Conference and Exhibit, 20-23 July 2003, Huntsville, Alabama. AIAA paper
2003-5009.
Ziemba, T. M. ―Experimental Investigation of the Mini-Magnetospheric Plasma
Propulsion Prototype.‖ PhD Thesis, Dept. of Aeronautics and Astronautics, Univerisity of
Washington. 2003.
Ziemba, T, P. Euripides, J. Slough, R. Winglee, L. Giersch, J. Carscadden, T.
Schnackenberg, and S. Isley. ―Plasma Characteristics of a High Power Helicon
Discharge.‖ Plasma Sources Sci. Technol. 15, 517, 2006.
Ziemer, J. K, M. Gamero-Castano, V. Hruby, D. Spence, N. Demmons, R. McCormick,
T. Roy, C. Gasdaska, J. Young, and B. Connolly. ―Colloid Micro-Newton Thruster
Development for the ST7-DRS and LISA Missions.‖ 41st AIAA/ASME/SAE/ASEE
Joint Propulsion Conference and Exhibit, 10-13 July 2005, Tuscon, Arizona. AIAA paper
2005-4265.
Ziemer, J. K, T. Randolph. V. Hruby, D. Spence, N. Demmons, T. Roy, W. Connolly, E.
Ehrbar, J. Zwahlen, and R. Martin. ―Colloid Microthrust Propulsion for the for the Space
Technology 7 and LISA Missions.‖ Laser Interferometer Space Antenna—6th
International LISA Symposium. p548. 2006
Ziemer, J. K. and S.M. Merkowitz. ―Microthrust Propulsion for the LISA Mission.‖ 40th
AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, FL, July, 2004. AIAA
paper 2004-3439.