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Wall Material Selection Process for CAM200 Low Power Hall Thruster

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Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
1
Wall Material Selection Process for CAM200 Low Power
Hall Thruster
IEPC-2015-103
Presented at Joint Conference of 30th International Symposium on Space Technology and Science
34th International Electric Propulsion Conference and 6th Nano-satellite Symposium,
Hyogo-Kobe, Japan
July 4 10, 2015
Raanan Eytan1, Dan Lev2, Gal Alon3 and Abraham Warshavsky4
Rafael - Advanced Defense Systems Ltd., Haifa, 3102102, Israel
Alexander Kapulkin5 and Maxim Rubanovitz6
Asher Space Research Institute (ASRI), Technion - Israel Institute of Technology, Haifa, 3200003, Israel
Low power Hall Effect Thrusters (HET) under 500 W are characterized by their low efficiency, low specific
impulse and short lifetime relative to their higher power cousins. The CAMILA and Simplified-CAMILA
Hall thrusters are an exception in that regard as their prototypes exhibited outstanding performance when
operated at power levels ranging from 100 W to 300 W. In the process of developing a flight-qualified low
power Hall thruster further investigations were performed on a Simplified CAMILA Development Model
(DM), denoted CAM200-DM. The thruster was designed and manufactured by Rafael, and tested at the
Asher Space Research Institute (ASRI). The present work describes the wall material selection process
campaign of CAM200-DM that includes structural integrity validation, performance tests and discharge
channel erosion measurements of three dielectric materials. Static, dynamic, sine and random vibration tests,
as well as shock tests, were conducted for grade HP-BN, grade M26-BN and ZSBN discharge channel
materials with a CAM200 Structural Model (SM). The thruster successfully passed all tests thus proving
adequate structural integrity of both structural design and wall material. Performance tests were conducted
at power levels in the range 100 W - 300 W. No significant difference in performance or erosion was observed
when comparing the dielectric rings with different materials after 80 hours of operation. CAM200-DM
exhibited exceptional performance when it generated, at 300 W, thrust above 16 mN, anode efficiency above
43% and specific impulse of above 1550 sec. During operation with ZSBN material rings the thruster
exhibited unstable electrical behavior and therefore this material was ruled out. HP-BN rings experienced
some geometry change after the short experimental campaign thus ruling out this material as well. The
chosen dielectric material is grade M26-BN that performed flawlessly.
Nomenclature
Ispa = Anode specific Impulse
P = Electrical Power
T = Thrust
ηa = Anode efficiency
1 Mechanical Research Engineer, Rafael, Manor Division, Space Department, email: eytanr@rafael.co.il.
2 Ph.D., Electric Propulsion R&D Responsible, Rafael, Manor Division, Space Department, danle@rafael.co.il.
3 Mechanical Research Engineer, Rafael, Manor Division, Space Department, email: galalo@rafael.co.il.
4 Mechanical Engineer, Engineering Consulting, email: 123avram@gmail.com.
5 Ph.D., Head of Electric Propulsion Laboratory, Asher Space Research Institute, Technion,
kapulkin@tx.technion.ac.il.
6 Laboratory Engineer, Asher Space Research Institute, Technion, email: aerrumax@tx.technion.ac.il.
IEPC-2015-103ISTS-2015-b-103
IEPC-2015-103ISTS-2015-b-103
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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I. Introduction
ue to their small size micro-satellites (10-100 kg) are power limited and conventionally utilize chemical
propulsion for their manuevering requirements. However, the use of high efficiency low power (<300W)
electric propulsion with Hall Effect Thrusters (HET) can enhance micro-satellites' capabilities and allow them to
perform a veriety of missions, particularly in Low Earth Orbits (LEO). In addition, thanks to the associated high
specific impulse the use of low power HETs allows for propellant mass reduction, thus reducing the overall satellite
mass. This, in turn, enables the enlargement of the mission payload if needed. Although HET technology carries
these advantages at power levels of 50 W to 300 W HETs suffer from low efficiency and short operating lifetime
relative to their higher power cousins1,2.
The Coaxial Anode Magneto-Isolated Longitudinal Anode (CAMILA) HET confronts these weaknesses. The
CAMILA HET has been developed and patented by Kapulkin et al3 from the Asher Space Research Institute (ASRI)
at the Technion, Israel institute of Technology. The CAMILA HET consists of an electrically floating gas distributer
and a co-axial anode, where a longitudinal magnetic field is induced by designated anode coils and a radial magnetic
field at the thruster's exit plane, which is induced by a conventional HET magnetic circuit4. This unique
configuration enables almost a full ionization of the propellant in the anode cavity even at low mass flow rates4,5.
Extensive experiments were conducted on CAMILA HET and on Simplified-CAMILA, which is a derivative of the
basic CAMILA HET, yet without the added anodic axial magnetic field. During experimental tests conducted at
ASRI, both configurations exhibited roughly the same performance of high efficiency and specific impulse at power
levels of around 250 W6.
In light of the above and because of its simpler and lighter
mechanical structure, a feature which is relevant to system
engineering considerations, it was decided to continue with the
development of simplified CAMILA, now denoted CAM200.
CAM200 is to be developed and qualified by Rafael to serve as
efficient means of propulsion for power-limited micro-satellites.
Moreover, CAM200 is designated to compete for a place in
the European-Israeli Micro-satellite Electric Propulsion System
(MEPS) project7. The goal of MEPS, which is funded by
European Space Agency (ESA) and the Israeli Space Agency
(ISA), is to design, manufacture and qualify an electric
propulsion system for micro-satellites. MEPS, which is planned
to be qualified for space in three years, is a collaboration
between RAFAEL and the Italian electric propulsion company,
ALTA-SITAEL.
This paper presents an important segment of the CAM200
development program - the material selection process in which the most suitable discharge wall material is chosen.
This development stage was conducted on a Development Model of CAM-200 (CAM200-DM).
II. CAM200 Thruster Description
CAM200-DM, shown in Figure 1, was
developed at Rafael with the participation of ASRI,
where the performance tests were conducted. The
magnetic field pattern and the inner structure of the
thruster are shown in Figure 2. One can observe that
the discharge channel cavity is formed of a co-axial
anode and ceramic dielectric rings attached to the
anode. Also shown is the Gas Distributer (GD)
which is gas-fed symmetrically through a xenon
feed manifold. The GD is electrically isolated from
the anode thus, it is electrically floating with respect
to the Thruster Unit (TU). The design of CAM200
includes key dimensions, which are the same as that
of the laboratory Simplified-CAMILA thruster.
D
Figure 1. Picture of CAM200-DM.
Axi-Symmetry Line
Anode
Discharge
Walls
Cathode
GD
Figure 2. Two dimensional magnetic field pattern and
thruster internal structure schematic of CAM200-DM.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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III. Material Selection Test Campaign
Ceramics are the most commonly used materials in HETs due to their low erosion under ion bombardment, high
temperature resistance and high secondary electron emission relative to other materials. On the other hand ceramic
materials are brittle and therefore require suitable thruster structural design to prevent material cracking and
fracturing. This is especially true during launch phase when the highest structural stresses are experienced by the
propulsion system. Of the available ceramics Boron-Nitride (BN) based ceramic materials exhibit highest HET
performance and are the natural candidates to be used in future HETs and specifically in CAM200.
The aim of the material selection program is to identify the most suitable discharge channel wall material that
can satisfy both structural integrity requirements and exhibits highest performance, both in terms of thrust efficiency
and low erosion. To do so three material candidates were examined: (1) High Purity BN (HP-BN) (2) M26-BN and
(3) Zirconia-based BN (ZSBN). All materials and their key properties are presented in Table 1.
Property
M26-
BN
ZSBN
Crystalline Phase
BN-60%,
SiO2-40%
BN-45%,
Zr2O3-45%
Directionality
Flexural Strength [MPa]
59
45
62
34
144
107
Thermal Conductivity (at 25°C) [W/mK]
27
29
11
29
24
34
Dielectric Strength [kV/mm]
66
>9
Table 1. Key properties of the three BN-based materials examined in the CAM200 material selection test plan
It can be seen that while pure hexagonal BN is a structurally weak ceramic material it has a relatively high
thermal conductivity, an important feature that helps maintain low wall temperature thus reduce channel erosion8.
Zirconia-based BN is the strongest material and has an excellent wear resistance yet it carries a possible drawback -
the risk of increased electrical conductivity at temperatures over 300°C9. Since ZSBN contains a high fraction of
zirconia (Zr2O3) it was acknowledged that there is the possibility that during HET operation the ZSBN ceramic wall
will start conducting the electrons trapped in the acceleration and ionization regions and essentially short-circuiting
the thruster. Still, due to the lack of existing literature on ZSBN electrical properties and its interaction with HET
plasma it was decided to keep this material in the material selection program. Lastly, M26-BN is a common
discharge channel wall material used in HETs10 exhibits properties similar, yet not identical, to HP-BN and therefore
is investigated as part of the test plan.
CAM200-DM ceramic wall material selection test program flow chart is presented in Figure 3.
The test program commences with X-ray computed tomography (CT) to verify initial ceramic wall structural
integrity. Subsequently a set of environmental tests are applied to verify structural integrity under expected launch
conditions. These include vibration and shock tests as well as frequency response test to measure the thruster's
structural resonanse frequency. The vibration and shock tests conditions are described in the subsequent sections.
Following the environmental tests an additional CT scan was performed to assure that no fractures were created
during the tests. Finally, an 80 hours thruster firing test is conducted to characterize thruster's performance, compare
Structural
Frequency
Response Testing
Sine and Random
Vibration Testing
Structural
Frequency
Response Testing
X-Ray Computed
Tomography (CT)
Shock Test X-Ray Computed
Tomography (CT) Performance Test Erosion
Measurement
Figure 3. Ceramic wall material selection test program flow chart.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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between CAM200-DM operating under the same conditions with different ceramic wall materials and compare the
the channel erosion of the different wall materials.
IV. Structural Integrity Tests
Modal analysis of CAM200 was performed using
'SolidWorks Simulation' final element software and
was part of CAM200-DM design and fabrication
process. To validate the design and mechanical
integrity of CAM200-DM dynamic environment
tests were performed on a Structural Model (SM) of
the thruster. As a test case the dynamic environment
specification requirements of the Venus satellite
were implemented11. These requirements consider
load amplifications on a small satellite due to launch
loads such as those exerted on the satellite platform
in various launchers. All tested vibration and shock
tests correspond to estimated worst case scenario
levels among all known commercial launchers. The
tests included sine and random vibration and shock
tests. Random vibration and shock test levels are
presented in Figure 4 and in Table 2 respectively.
Figure 5 depicts the x-axis random vibrations test setup along with an illustration of the Hall thruster structural
model.
All three material rings assembled in the SM successfully passed all the dynamic tests, therefore proving
adequate structural integrity and paving the way for the next stage of the material selection program.
V. Performance Testing
A. Test Facility
Performance tests were conducted at the
Asher Space Research Institute (ASRI) electric
propulsion laboratory at the Technion (Figure 6).
The test facility includes a 1.22m Ø × 2.7m
stainless steel vacuum chamber with inner
volume of approximately 3.2 m3. The chamber is
equipped with a 'RUTA WAU 501/D65B/A'
forevacuum pump with pumping speed of 230
liter/s of air. Three 'Sumitomo' CP-22 cryopumps
are located on the vacuum chamber's opposite
end from the thruster. The combined measured
pumping speed achieved is approximately 9,000
liter/s on xenon. The residual pressure of the
vacuum chamber is lower than 9×10-8 Torr.
During thruster operation at a total flow rate of
1.2 mg/s of xenon, chamber pressure was lower
than 2×10-5 Torr.
10
2
10
3
10
-2
10
-1
10
0
Frequency [Hz]
Level [g
2
/Hz]
X-Y Axes
Z Axis
Figure 4. Vibrat
ion test levels performed on a structural
model of CAM200 with all three types of wall channel
materials.
Frequency [Hz]
Level [g]
100
150
2000
3000
10000
3000
Table 2. Shock test levels.
Z
YX
Figure
5.
Left: CAM200 during random vibration test along
the thruster's X
-
axis. Right: CAD model of CAM200
structural model.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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Pressure is measured using an 'Ionivac' sensor, model ITR90 from 'Leybold', located behind the thruster. The
'Ionivac' pressure gauge is calibrated by the manufacturer and has a declared precision of 15%. A commercial
0÷1.96 mg/s flow controller, model M100b from MKS, controls xenon flow rate to the thruster’s gas distributor.
Similarly, a 0÷0.98 mg/s flow controller controls flow rate to the cathode which set to 0.15 mg/s during all tests. The
mass flow controllers are calibrated by the manufacturer, with a declared precession of 1% of full scale. The thrust
stand used is a feedback-controlled inverted pendulum type, manufactured for ASRI by PLATAR LTD, Moscow,
Russia. It is capable of thrust measurements in the range 1÷200 mN. Measurement accuracies, as declared and
verified by the manufacturer, are ±3.5% between 5÷10 mN and ±3% between 10÷20 mN. During tests the thrust
stand is calibrated using weights once in every two hours.
The experiments were conducted on CAM200-DM, using a
low-current heaterless hollow cathode procured from Kharkiv
Aviation Institute (KhAI) of Kharkiv, Ukraine. A dedicated
Ignition Block (I/B) supplies the high voltage pulse necessary to
ignite the cathode as well as the keeper holding voltage required
to sustain emitter-keeper plasma after cathode discharge is set.
The thruster discharge voltage is supplied by the main P/S that
connects between the anode and cathode. Two power supplies
feed current to the magnetic circuit; one to the outer coils
assembly and the other to the centeral coil. This arrangement
enables small magnetic field pattern adjustments during tests in
case unexpected thruster behavior is encountered. The thruster
electrical setup schematic is depicted in Figure 7.
B. Operation Test Plan
The wall material selection criteria is based on (a)
performance (b) erosion level and (c) material dimensional
stability. For each material examined the test plan commenced
with 15 hours performance investigation at three power levels,
100 W, 200 W and 300 W, followed by 50 hours of operation at
200 W and concluded with a second performance investigation at
Ion Gauge
Thrust
Stand
Thruster
MFC
MFC
CryoPump
CryoPump CryoPump
Xe
To Cathode
To Anode
Forvacuum
Pump
2.7 m
1.22mØ
Air Vent
Figure
6. Schematic of the Asher Space Research Institute (ASRI)
electric propulsion laboratory
experimental facility
(not to scale).
Outer Coil P/S
Inner Coil P/S
Ignition BlockAnode P/S
++
-
Figure 7. Left: CAM200-DM electrical
experimentation scheme.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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the above three power levels. The entire operation duration for each wall material was 80 hours. After thruster
operation concluded, for each wall material test run, the discharge channel rings were visually examined and
material erosion measured for comparison.
To measure discharge rings erosion their profiles were scanned prior and post operation. Scanning was
performed with a FARO-GAGE-1001-0411, four arms, 3D measuring system with a measurement resolution of 0.02
mm. Although Hall thruster erosion is usually not significant after less than 100 hours of operation thanks to the
high accuracy of the measurement equipment minor variations in erosion should be possible to detect. Moreover,
wall erosion is expected to be significantly larger in low-power Hall thrusters in general, and at the thruster's
beginning of life12,13 in particular, which is the case of the CAM200-DM performance testing.
C. Test Results and Discussion
The first discharge wall material to be tested was ZSBN. Soon after ignition it was found that the ZSBN
discharge rings became electrically conductive to a detrimental level. This behavior manifested by a sharp increase
in discharge current accompanied by a visual change in the shape of the main discharge plasma at the channel exit
plane. The thruster eventually shut down spontaniously. Therefore, it was concluded that thruster operation with
ZSBN discharge rings could not be properly executed and the material was disqualified. Subsequently the discharge
rings were replaced with HP-BN discharge rings followed by M26-BN rings. A picture of CAM200-DM during
operation is shown in Figure 8(d).
The performance of CAM200-EM, in terms of specific impulse, thrust and anode efficiency, assembled with HP-
BN and M26-BN discharge ring are presented in Figure 8(a)-(c). It can be observed from the figure that at all
presented anode power both discharge ring types manifested comparable performance within errorbars. Therefore,
there is no particular reason to prefer one material over the other based on thruster performance.
It can also be observed from the figure that at low power of 100 W CAM200-DM generates thrust higher than 5
100 200 300 400
4
6
8
10
12
14
16
18
20
22
Anode Power [W]
Thrust [mN]
M-26 (60%BN-40%SiO
2
)
HPBN (High Purity BN)
100 200 300 400
800
1000
1200
1400
1600
1800
Anode Power [W]
Isp
a
[sec]
M-26 (60%BN-40%SiO
2
)
HPBN (High Purity BN)
100 200 300 400
20
25
30
35
40
45
50
55
Anode Power [W]
K
a [%]
M-26 (60%BN-40%SiO
2
)
HPBN (High Purity BN)
(a) (b)
(c) (d)
Figure
8. Measured performance of CAM200-DM. (a) Thrust, (b) Specific Impulse, (c) Anod
e Efficiency and (d)
Picture of CAM200
-DM during operating.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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mN, specific impulse around 1000 sec and anode efficiency higher than 25%. Additionally, at higher power levels of
about 300 W CAM200-DM generates thrust higher than 16 mN, specific impulse around 1600 sec and anode
efficiency higher than 42%. All performance results are quite satisfactory and clearly indicate on the advantage of
using the CAM200 Hall thruster at low power under 300 W.
After thruster operation was concluded the thruster was taken out of the vacuum chamber and discharge rings
visually examined. After taking out the thruster with the HP-BN discharge rings it was noticed that the rings shrunk
in size and is not pressure-locked as it was prior to thruster operation. Albeit the reduction in HP-BN ceramic
channel geometry similar behavior was not observed for the case of M26-BN discharge channel rings. Be that as it
may the ceramic discharge channel looked sound and without any visible damage.
It is speculated that the change in HP-BN dimensions during thruster operation can be attributed to the sintering
process of the ceramic. It is possible that the combination of temperatures of over 300°C and the pressure exerted by
the thruster mechanical configuration led to ceramic material density increase. It was decided to investigate this
behavior in more detail in the 'Erosion Measurement' phase where more accurate dimensional analysis is conducted.
D. Relative Erosion Estimation
According to the test plan after the thruster was disassembled and discharge
ceramic rings taken out their dimensions were measured. The purpose of the
measurements is to examine possible differences in erosion between the thruster
with two discharge channel materials - HP-BN and M26-BN.
The dimensional measurements are presented in Figure 9. It can be seen from
the figure that both channel type materials were eroded more on the inner
discharge ring, close to the thruster center, than on the outer discharge ring. The
highest erosion measured is about 0.5 mm on the inner discharge ring at the
thruster exit plane. It should be noted that erosion rate in Hall thrusters is highly
non-linear13-15 and diminished significantly during the thruster lifetime. It can also
be observed from the erosion measurement results that the HP-BN channel
experienced erosion similar to the M26-BN discharge channel. Therefore no
obvious advantage in erosion-resistance ability can be attributed to any one of the
tested materials.
The most important finding detected during the channel dimension
examination was a minor dimension reduction in the HP-BN discharge rings of the
order of ~0.1 mm. Due to the small dimensional change it is hard to discern it in
the corresponding figure. Simply put, the discharge channel rings experienced a
minor shrinkage that made them loose within the thruster configuration. This
phenomenon is not surprising as it was previously observed in Hall thruster
discharge channel materials16 when using HP-BN purchased from the same
vendor. It was also found in the same research that M26-BN discharge channel
materials do not shrink at temperatures under 1000°C such as in the CAM200 Hall
thruster.
Due to the fact that future acceptance test procedures, on future CAM200
thrusters, will include a short thruster operation segment similar phenomenon is
expected if the thruster is to be used with HP-BN. Discharge channel shrinkage
prior to launch is unacceptable because of possible jilts that might occur to the ceramic channel. For this reason, and
because both HP-BN and M26-BN materials exhibited similar performance, it was decided to disqualify the HP-BN
material.
The material chosen for the CAM200 Hall thruster is M26-BN which is also the material used in other
conventional Hall thrusters.
VI. Conclusion
The material selection program of CAM200 was concluded. The program included a structural integrity test,
performance test and erosion test. Three discharge channel materials were examined - HP-BN, M26-BN and ZSBN.
All three discharge channel materials were tested and approved for high launching loads, therefore validating
both the thruster structural configuration and discharge ring materials.
Thruster performance was tested with all three discharge channel materials. During this test the ZSBN-made
discharge channel rings became electrically conductive to a detrimental level. For this reason ZSBN was
Original Outline (0 Hr)
M26 BN (80 Hr)
HP BN (80 Hr)
1mm
Inner Channel
Outer Channel
Figure 9
. Ceramic channel
dimensions for HP-BN
and
M26-BN
ceramic channel
materials before and after
thruster operation.
Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Kobe-Hyogo, Japan
July 4 10, 2015
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disqualified and performance tests continued only with HP-BN and M26-BN channel materials. Thruster
performance with HP-BN and M26-BN rings was similar without any obvious advantage to neither of them.
Erosion measurements were performed on the discharge rings after 80 hours of thruster operation and showed no
significant difference in erosion between the two materials. On the other hand a slight reduction in dimensions was
observed in the HP-BN rings that made them loose within the thruster configuration. To avoid any possible future
complications it was decided to continue with the M26-BN material.
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... Those materials offer similar thruster performances. 3,4 However, additives such as silicon nitride and aluminum nitride can degrade performance by modifying the secondary electron emission (SEE) behavior of the walls. 5 Data on the performance of Hall thrusters with alternative materials are less common. ...
Full-text available
Article
Traditional Hall thrusters are fitted with boron nitride dielectric discharge channels that confine the plasma discharge. Wall properties have significant effects on the performances and stability of the thrusters. In magnetically shielded thrusters, interactions between the plasma and the walls are greatly reduced, and the potential drop responsible for ion acceleration is situated outside the channel. This opens the way to the utilization of alternative materials for the discharge channel. In this work, graphite walls are compared to BN-SiO2 walls in the 200 W magnetically shielded ISCT200-MS and the unshielded ISCT200-US Hall thrusters. The magnetically shielded thruster shows no significant change in the discharge current mean value and oscillations, while the unshielded thruster's discharge current increases by 25% and becomes noticeably less stable. The electric field profile is also investigated through laser spectroscopy, and no significant difference is recorded between the ceramic and graphite cases for the shielded thruster. The unshielded thruster, on the other hand, has its acceleration region shifted 15% of the channel length downstream. Lastly, the plume profile is measured with planar probes fitted with guard rings. Once again the material wall has little influence on the plume characteristics in the shielded thruster, while the unshielded one is significantly affected.
... The Russian SPT-50 employs a 5 cm discharge channel outer diameter and demonstrates a maximum anode efficiency of nearly 40% with a thrust of between 20 and 30 mN and a specific impulse of between 1300 and 2000 s over discharge powers of 350-500 W; a flight- demonstrated lifetime of approximately 2500 h has been reported [6,7]. Recently, thrusters such as the Plas-40, CAM200, and HT100 have been developed and tested, yet no published data have shown operational lifetimes greater than a few thousand hours [8][9][10]. In addition to operational lifetime, thruster efficiency is a critical performance metric for small satellites. ...
Article
The successful application of a fully shielding magnetic field topology in a low-power Hall thruster is demonstrated through the testing of the MaSMi-60 Hall thruster (an improved variant of the original Magnetically Shielded Miniature Hall thruster). The device was operated at discharge powers from 160 to 750 W at discharge voltages ranging from 200 to 400 V. Several techniques were used to determine the effectiveness of magnetic shielding achieved by the MaSMi-60 and to estimate the reduction in discharge channel erosion rate enabled by the shielding field topology. This ultimately suggested an improvement in discharge channel life by a factor of at least 10 times, and likely greater than 100 times, when compared to unshielded devices. Thruster performance, measured both directly by a thrust stand and indirectly by plume diagnostics, was lower than expected when compared to results from high-power magnetically shielded Hall thrusters. However, the plume diagnostic measurements enabled the identification of the primary causes for the MaSMi-60's moderate performance. © 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
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Thesis
Hall thrusters are one of the most used rocket electric propulsion technology. They combine moderate specific impulse with high thrust to power ratio which makes them ideal for a wide range of practical commercial and scientific applications. One of their limitations is the erosion of the thruster walls which reduces their lifespan.The magnetic shielding topology is a proposed solution to prolong the lifespan. It is implemented on a small200W Hall thruster.In this thesis the scaling of classical unshielded Hall thrusters down to 200 and 100W is discussed. A 200W low power magnetically shielded Hall thruster is compared with an identically sized unshielded one. The ion behavior inside the thruster is measured and significant differences are found across the discharge channel.Both thrusters are tested with classical BN-SiO2 and graphite walls. The magnetically shielded thruster is not sensitive to the material change while the discharge current increase by 25% in the unshielded one. The result is a maximum efficiency of 38% for boron nitride in the unshielded thruster but only 31% with graphite.The shielded thruster achieves a significantly lower efficiency with only 25% efficiency with both materials.Analysis of the experimental results as well as simulations of the thrusters reveal that the performance difference is mostly caused by low propellant utilization. This low propellant utilization comes from the fact that the ionization region doesn’t cover all of the discharge channel. A new magnetically shielded thruster is designed to solve this issue.
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Article
The operation of a laboratory version of the flight-qualified SPT-100 stationary plasma thruster is compared for four different discharge chamber wall materials: a boron nitride–silica mixture (borosil), alumina, silicon carbide, and graphite. The discharge is found to be significantly affected by the nature of the walls: changes in operating regimes, up to 25% variations of the mean discharge current, and over 100% variations of the discharge current fluctuation amplitude are observed between materials. Thrust, however, is only moderately affected. Borosil is the only material tested that allows operating the thruster at a low mean current, low fluctuation level and high thrust efficiency regime. It is suggested that secondary electron emission under electron bombardment is the main cause of the observed differences in discharge operation, except for graphite, where the short-circuit current inside the walls is believed to play a major role. It is also suggested that the photoelectric effect, which has apparently not been given attention before in the Hall thruster literature, could increase the cross-field electron current. © 2003 American Institute of Physics.
Article
The low-energy sputtering of boron nitride, magnesium oxide, boron nitride and aluminum nitride (BNAlN), and boron nitride and silicon oxide (BNSiO2) by xenon ions of bombarding energies 350, 500 eV, and 1 keV was studied experimentally. In order to measure the ion current without being significantly disturbed by slow ions, only planar probes were used during short duration sputtering experiments (of the order of 10 h). Moreover, slow ion current contribution was estimated by numerical simulations and subtracted from each ion current measurement. It was found that the ion-beam incidence effect on sputtering yields was not as important as for theoretical results or experimental results on quasinonrough solid surfaces, for which it is possible to observe a more pronounced angular dependence of the sputtering yield. This phenomenon is due to surface irregularities of ceramic materials and because of surface roughness the macroscopic sputtering yield should actually result from the convolution of the microscopic sputtering yield by the angular distribution of surface facet incidences. The irregular surface structure of ceramics like BNSiO2 or BNAlN seems to be sputtered differently due to a differential erosion from grain to grain and from grain to surrounding matrix. This uneven erosion may be explained by the wide angular distribution of facet incidences of surface microprofiles and the various binding energies in the selvage of material. Finally, the dependence of sputtering yield at normal incidence on ion energy, in the range 0.35–1 keV, is almost a linear one. © 1999 American Vacuum Society.
Article
The purpose of the present work is to compare data obtained by impedance spectroscopy in several laboratories on the same ionically conductive materials. Yttria-doped zirconia, stabilized under the cubic form, was selected. Measurements were done in the temperature range 200-600°C to get the bulk and the grain boundary contributions to the material impedance. The data were analyzed centrally. The dc conductivity could be obtained with a fair accuracy. Problems concerning the determination of ac properties are clearly pointed out. Some recommendations are made to improve the general reliability of measurements made with the aid of impedance spectroscopy.
Evaluation of Low Power Hall Thruster Propulsion AIAA-96- 2736. 3 United States, Patent Application PublicationTheoretical Modeling of Ionization Processes in Anode Cavity of CAMILA Hall Thruster
  • J M Ahedo
  • D Gallardo
  • S Manzella
  • J Oleson
  • T Sankovic
  • A Haag
  • V Semenkin
  • Lake Buena Kim
  • Vista
  • Fl
  • A Usa Kapulkin
  • M Guelman
  • V Balabanov
  • B Rubin
  • M Pub
  • I E Guelman Kronhaus
  • L Behar
  • A Rabinovich
  • Warshavsky
Ahedo and J. M. Gallardo. "Scaling Down Hall Thrusters". Proceedings of the 29 th International Electric Propulsion Conference (IEPC), 17-21 March, 2003, Toulouse, France, IEPC-03-0104. 2 D. Manzella, S. Oleson, J. Sankovic, T. Haag, A. Semenkin and V. Kim. "Evaluation of Low Power Hall Thruster Propulsion". Proceedings of the 32 nd Joint Propulsion Conference (JPC), 1-3 July, 1996, Lake Buena Vista, FL, USA, AIAA-96- 2736. 3 United States, Patent Application Publication, Kapulkin A., Guelman M., Balabanov V. and Rubin B., Pub. No.: US 2010/0107596 A1, Pub. Date: May 6, 2010 4 A. Kapulkin and M. Guelman. "Theoretical Modeling of Ionization Processes in Anode Cavity of CAMILA Hall Thruster". Proceedings of the 31st International Electric Propulsion Conference (IEPC), 20-24 September, 2009, Ann Arbor, Michigan, USA, IEPC-2009-068, 2003. 5 Kronhaus, I., " Experimental and Numerical Investigations of the Physical Processes in a Co-Axial Magneto-Isolated Longitudinal Anode Hall Thruster ". Ph.D. Dissertation, Aerospace Dept., Technion -Israel Institute of Technology, Haifa, Israel, 2012. 6 A. Kapulkin, V. Balabanov, M. Rubanovich, E. Behar, L. Rabinovich and A. Warshavsky. "CAMILA Hall Thruster: New Results". Proceedings of the 32 nd International Electric Propulsion Conference (IEPC), October 11-15, 2011, Wiesbaden, Germany, IEPC-2011-041, 2011. 7 T. Misuri, et al. " MEPS Programme -Development of a Low Power, Low Cost HET for Small Satellites ", SP2014_ 2969515, Space Propulsion Conference, Cologne, May 2015 8 A. N. Raneev, A. A. Semenov and O. B. Solovyev. "Surface Temperature Influence onto Ceramic Sputtering Rate". Proceedings of the 13 th International Conference about Ion Interaction with a Surface, Moscow, Russia, 1997. 9
Scaling Down Hall Thrusters
  • E Ahedo
  • J M Gallardo
E. Ahedo and J. M. Gallardo. "Scaling Down Hall Thrusters". Proceedings of the 29 th International Electric Propulsion Conference (IEPC), 17-21 March, 2003, Toulouse, France, IEPC-03-0104.
Evaluation of Low Power Hall Thruster Propulsion
  • D Manzella
  • S Oleson
  • J Sankovic
  • T Haag
  • A Semenkin
  • V Kim
D. Manzella, S. Oleson, J. Sankovic, T. Haag, A. Semenkin and V. Kim. "Evaluation of Low Power Hall Thruster Propulsion". Proceedings of the 32 nd Joint Propulsion Conference (JPC), 1-3 July, 1996, Lake Buena Vista, FL, USA, AIAA-96-2736.
Theoretical Modeling of Ionization Processes in Anode Cavity of CAMILA Hall Thruster
  • A Kapulkin
  • M Guelman
A. Kapulkin and M. Guelman. "Theoretical Modeling of Ionization Processes in Anode Cavity of CAMILA Hall Thruster". Proceedings of the 31st International Electric Propulsion Conference (IEPC), 20-24 September, 2009, Ann Arbor, Michigan, USA, IEPC-2009-068, 2003.
Experimental and Numerical Investigations of the Physical Processes in a Co-Axial Magneto-Isolated Longitudinal Anode Hall Thruster
  • I Kronhaus
Kronhaus, I., "Experimental and Numerical Investigations of the Physical Processes in a Co-Axial Magneto-Isolated Longitudinal Anode Hall Thruster". Ph.D. Dissertation, Aerospace Dept., Technion -Israel Institute of Technology, Haifa, Israel, 2012.
CAMILA Hall Thruster: New Results
  • A Kapulkin
  • V Balabanov
  • M Rubanovich
  • E Behar
  • L Rabinovich
  • A Warshavsky
A. Kapulkin, V. Balabanov, M. Rubanovich, E. Behar, L. Rabinovich and A. Warshavsky. "CAMILA Hall Thruster: New Results". Proceedings of the 32 nd International Electric Propulsion Conference (IEPC), October 11-15, 2011, Wiesbaden, Germany, IEPC-2011-041, 2011.