- Email: [email protected]
Performance enhancement of an argon-based propellant in a Hall thruster
Accepted Manuscript Performance enhancement of an argon-based propellant in a Hall thruster Junko Yamasaki, Shigeru Yokota, Kohei Shimamura PII:
S0042-207X(18)30244-6
DOI:
10.1016/j.vacuum.2018.09.042
Reference:
VAC 8255
To appear in:
Vacuum
Received Date: 10 February 2018 Revised Date:
20 September 2018
Accepted Date: 21 September 2018
Please cite this article as: Yamasaki J, Yokota S, Shimamura K, Performance enhancement of an argonbased propellant in a Hall thruster, Vacuum (2018), doi: https://doi.org/10.1016/j.vacuum.2018.09.042. This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting proof before it is published in its final form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain.
ACCEPTED MANUSCRIPT
Title:
Performance Enhancement of an Argon-Based Propellant in a Hall
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Thruster Authors:
Junko Yamasaki a *, Shigeru Yokota b , and Kohei Shimamura b
Department of Engineering Mechanics and Energy, University of Tsukuba
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a
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Affiliations:
1-1-1 Tennodai, Tsukuba, Ibaraki 305-0006, Japan b Division
of Engineering Mechanics and Energy, University of Tsukuba
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1-1-1 Tennodai, Tsukuba, Ibaraki 305-0006, Japan
*Corresponding Author: Junko Yamasaki
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Department of Engineering Mechanics and Energy, University of Tsukuba
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1-1-1 Tennodai,
Tsukuba, Ibaraki 305-0006, Japan Tel: 81-29-853-5138
Email: [email protected]
1
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Abstract: Finding alternatives to xenon as a propellant is essential for reducing the launching cost of
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all-electric satellites and for achieving low-cost electric propulsion, especially using Hall thrusters. Argon is a candidate as such a low-cost propellant, but it has never been used practically owing to its poor performance. We attempted two methods to improve the thruster performance: (1) by
SC
mixing xenon in argon propellant and (2) by modifying the thruster structure for the argon propellant. Modifying the thruster structure yielded an increase in the specific impulse from 482 to
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1500 s using pure argon. Also, when we used an argon-xenon mixture gas as propellant, the thruster performance increased under argon-rich conditions using the modified thruster.
Keywords:
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Electric Propulsion, Hall Thruster, Alternative propellant
2
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1. Introduction One of the most effective ways to economize on space transportation is to replace
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conventional chemical propulsion by electric propulsion. For example, a Hall thruster, one of the most promising electric propulsion systems, can save $13 million on the cost of launching a geostationary orbit satellite.[1] Anode-layer-type Hall thrusters are currently
SC
actively researched in Japan.[2][3]
On the other hand, Hall thruster propellants, consisting of xenon gas, are much more
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expensive than those used for chemical propulsion. To address this problem, alternative propellants have been tested. [4-13] Argon is cheap but is not considered to be a strong candidate because of its poor thruster performance [6], which increases the effective launching cost.
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The low thrust performance of argon stems from its higher ionization energy and lighter weight, which make a large amount of input electric power go into ionization rather than thrust power. Argon is therefore rapidly exhausted from the ionization channel
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because of its higher thermal velocity. [6] Therefore, we suggest enhancing argon ionization using small amount of xenon. It is expected that electrons produced by xenon
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ionization collide with neutral argon particles, thereby increasing the mass utilization efficiency of argon.
The lower ionization energy of xenon in the mixture is expected to facilitate argon
ionization. In this study, we measured the thrust performance by applying propellant mixtures with various mixture ratios of argon and xenon. In addition, we tested the thrust performance with an extended discharge channel, which is also expected to enhance
3
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ionization.
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2. Mixture propellant effect 2.1 Experimental procedure
Figures 1 shows a cross section of the sub-kilowatt class anode-layer-type Hall
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thruster developed at the University of Tsukuba.[13] The inner and outer diameters of the discharge channel are 33 and 43 mm, respectively. The discharge channel length is 3 mm.
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The magnetic field in the acceleration channel is applied by one inner coil and three outer coils. To stabilize the operation, the magnetic flux density near the anode was shielded by the magnetic screen. The normalized magnetic field profiles are shown in Fig.2. In this chapter, we used the thruster “without extended channel”.
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Figure 3 shows a schematic diagram of the experimental setup. The Hall thruster was operated in a vacuum chamber whose diameter and length were both 1.0 m. The vacuum chamber was evacuated by a turbo molecular pump (3,000 L/s) backed by a rotary pump
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(1,500 L/min). The background pressure varied depending on the xenon mixture ratio and the backpressure increased gradually with the xenon fraction. The backpressure reached a
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maximum of 3.8×10-2 Pa in pure xenon, and a minimum of 2.6 ×10-2 Pa in pure argon. The thrust was measured by a pendulum-type thrust stand, also developed at the
University of Tsukuba. Only knife-edge joints were used between the components. To measure the discharge current Id, the voltage between both ends of a 0.5 Ω metal-film resistor (inserted between the anode and discharge power supply) was measured with an oscilloscope using a differential probe. The ion beam current Ib was measured using the ion
4
Figure 1 Figure 2 Table 1
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collector, a 250×300 mm copper plate biased at -20 V with respect to the cathode. The collector was placed 30 cm downstream from the thruster. Another 0.5 Ω metal-film resistor
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was used for the current measurement. The collected ion current must be corrected to eliminate charge-exchange (CEX) collisions between the thruster and the ion collector. Owing to a lack of Ar-Xe+ and Xe-Ar+ CEX collision cross-section data, Ib was corrected
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using only the Ar-Ar+ collision cross section data or only the Xe-Xe+ collision
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． cross-section data. From Ib, the mass utilization efficiency (ηu=miIb/em) was deduced. Here, ． mi, e, and m are the ion mass, elementary charge, and mass flow rate, respectively. The results for ηu, corrected using the Ar-Ar+ data and Xe-Xe+ data separately, are plotted in Fig. 4.
The operation conditions are tabulated in Table 1. The total mass flow rate was 27.8
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sccm. The xenon fraction α was varied from 0 to 100%. The discharge voltage was fixed at 200 V. The magnetic flux density was set to maximize the thrust efficiency ηt for each xenon fraction. For example, a peak value of 60 mT was applied in the case of pure xenon,
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and 10 mT was applied in the case of argon. Here, the thrust efficiency of an electrically
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． powered thruster is defined as ηt = F2/mP, where, F and P are the thrust and input electrical ． power, respectively. [15] Also, we deduced the specific impulse Isp = F/mg using the measured thrust in order to discuss the “propellant consumption efficiency”. 2.2 Result and Discussion The thrust efficiency and the mass utilization efficiency are shown in Figs. 4 and 5, respectively, as functions of the xenon fraction. The error bars represent the standard deviations in the data. This chapter discusses only the normal case (denoted by circles); the
5
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extended discharge channel (square symbols) is explained in the next chapter. The thrust efficiency increases with the xenon fraction (Fig. 4). The thrust efficiency
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was less than 5% for pure argon but 48 % for the pure xenon case, consistent with earlier studies [12][14][16]. The dominant factor behind the ηt increase is the increase in ηu with increasing xenon fraction (Fig. 5). Thus, a low thrust efficiency when using argon
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propellant can be improved by enhancing ionization via a mixture with a low-ionization-energy substance like xenon. For simplicity, we here consider only singly
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charged ions. According to reference [17], the fraction of multiply charged ions is relatively small: the Xe2+ and Xe3+ fractions are 3% and 1%, respectively, for a discharge voltage of 300 V. Therefore, the contribution of multiply charged ions to ionization enhancement by xenon is small.
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The specific impulse Isp is shown in Fig. 6 as a function of the xenon fraction, displaying an increase with the xenon fraction and attaining a maximum when the xenon ． ． fraction is 70%. As mentioned above, Isp is defined as F/mg. Here, obviously m is
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monotonically increasing functions of α. On the other hand, measured F was also monotonically increasing functions of a (see Fig. 7). As a consequence, Isp is a upward-convex function of α.
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Figure 3 Figure 4
We note that Isp for pure xenon reaches 1500 s, approximately 100 s more than with
other anode-layer-type Hall thrusters [14], so that Isp in the xenon-rich case may be overestimated because of the effect of the background pressure. On the other hand, argon is assumed not to contribute significantly to an increase in background pressure because of its higher pumping speed. Therefore, the estimate of Isp in the argon-rich case is likely to be
6
Figure 5
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3. Ionization enhancement by an extended discharge channel 3.1 Hall thruster with an extended discharge channel
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more correct than under xenon-rich conditions.
As discussed above, a greater Isp is achieved using argon gas rather than pure xenon.
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However, a higher ηt under argon-rich conditions is needed to reduce the propellant cost. Therefore, the discharge channel was extended to enhance argon ionization. The
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ionization-collision probability is thus expected to be higher because the propellant channel transit time becomes longer. Figure 8 shows the Hall thruster with the extended discharge channel. The distance between the anode tip and thruster exit plane is 9 mm. The magnetic flux density distribution is the “with extended channel” case in Fig.3. The experimental
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procedure is described above.
3.2 Result and discussion
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The open squares in Figs. 4 and 5 display ηt and ηu, respectively, as functions of
α when using the extended discharge channel. As shown in Fig. 4, ηt is improved using the
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extended discharge channel under argon-rich conditions (α < 50%) thanks to the enhancement of ionization resulting from the extended discharge channel, as intended. There is also an improvement in ηu when using the extended discharge channel (Fig. 5). On the other hand, as a result of the discharge channel extension, ηt decreases under xenon-rich conditions (α > 50%) despite ηu increasing in all the mixtures. This result indicates that the extended discharge channel is too long for the xenon propellant; a 7
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proportion of ionized xenon collides with the long-channel wall before exhaust release, so that the extra energy is used for the re-ionization of the recombined xenon atoms.[6]
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Figure 6 plots Isp as a function of α. Especially under argon-rich conditions, Isp increases to 1500 s over the entire α range when the discharge channel is extended. Because argon is lighter than xenon (39.95 amu, compared to 131.29 amu for xenon), the
SC
greater the increase in α, the higher the atomic mass and the lower the exhaust speed. Therefore, Isp increases much more than the efficiency in the case of an extended discharge
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channel because Isp is also defined as v/g, where g is the acceleration due to gravity and v is the exhaust speed. Moreover, considering the back-pressure effect mentioned above, it is expected that Isp under argon-rich conditions is greater than under xenon-rich conditions. This means that an argon or an argon-rich mixture is more advantageous than xenon in
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terms of cost for space transportation.
Whereas the extended discharge channel is effective for achieving ionization enhancement, it is likely to shorten the thruster lifetime owing to wall erosion. Optimizing
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future work.
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the channel length for low-ionization-energy gas propellants thus remains an objective for
5. Conclusion
(1) The thrust performance of the Hall thruster using an argon-xenon gas mixture was investigated. The thrust efficiency and the mass utilization efficiency increased with increasing xenon fraction in the gas mixture. The specific impulse reached a maximum for a
8
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xenon fraction of approximately 70 % as a result of the trade-off between ion mass and thrust efficiency.
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(2) We tested the thrust performance using an extended discharge channel, which is expected to enhance ionization. The thrust efficiency increased by up to 300 % under argon-rich conditions, but decreased under xenon-rich conditions. This result indicates that
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the longer discharge channel increases not only the ionization rate but also the rate of ion loss to the wall. As the result of the improvement in the thrust efficiency, the specific
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impulse reached approximately 1500 s regardless of the xenon fraction.
Acknowledgements
References [1]
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This work was supported by JSPS KAKENHI Grand Number 15H0558.
K. Kinefuchi, N. Nagao, Y. Saito, K. Okita, K. Kuninaka. Test Facility Concept for
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High Power Electric Propulsion, STEP-2013-024, Space Transportation Symposium FY2013, 2014. (in Japanese) Y. Hamada et al. Hall thruster development for Japanese space propulsion programs,
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[2]
Transactions of JSASS, 60 (2017) 320-326.
[3]
K. Komurasaki, Y. Arakawa, Definition of Hall Current Ion-Thruster Performance, Journal of Propulsion and Power 8 (1992) 1212-1216.
[4]
J. A Linnell, A. D. Gallimore, Efficiency Analysis of a Hall Thruster Operating with Krypton and Xenon. J. Propul. Power 22 (2006) 1402–1412.
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[5]
Lorand, O. Duchemin, S. Zurbach, D. LE. Mehaute, N. Cornu, Alternate propellants for PPS Hall-Effect Plasma Thruster. IEPC-2013-074, 33rd International Electric
[6]
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Propulsion Conference, 2013.
Eunsun Cha1, David B. Scharfe, Michelle K. Scharfe, and Mark A. Cappelli, Hybrid Simulations of Hall Thrusters Operating on Various Propellants.
[7]
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IEPC-2009-075, 29th International Electric Propulsion Conference, 2009.
J. Szabo, M. Robin, V. Hruby, Bithmath Vapor Hall Effect Thruster Performance
Conference, 2017. [8]
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and Plum Experiments, IEPC-2017-25, 35th International Electric Propulsion
M. A. Hopkins and L. B. King, Performance Comparison Between a Magnesiumand Xenon-Fueled 2 Kilowatt Hall Thruster, J. Propul. Power 32 (2016) 1015-1021. J. Szabo, M. Robin, J. Duggan, Light Metal Propellant Hall Thrusters,
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[9]
IEPC-2009-138, 31st International Electric Propulsion Conference, 2009. [10]
Adam Shabshelowitz, Alec D. Gallimore, Peter Y. Peterson, Alternative propellant
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Performance of a Helicon Hall Thruster Operating with Xenon, Argon, and Nitrogen, Journal of Propulsion and Power, 2014, Vol.30: 664-671. James Szabo, Bruce Pote, Surjeet Paintal, Mike Robin, Adam Hillier, Richard D.
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[11]
Branam, Richard E. Huffmann, Performance Evaluation of an Iodine-Vapor Hall Thruster, Journal of Propulsion and Power, 2012, Vol.28: 848-857
[12]
A.Lorand, O.Duchemin, S. Zurbach, D.Mehaute , N. Cornu, Alternate propellants
for PPS® Hall-Effect Plasma Thruster, IEPC-2013-074, 2013 [13]
J. Yamasaki, S. Yokota, K. Shimamura, Operation Characteristics of Anode Layer
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Hall Thruster Using Alternative Propellant, ISTS-2017-b-13, 31st International Symposium on Space Technology and Science. 2017. D. Manzella et al. Evaluation of Low Power Hall Thruster Propulsion,
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[14]
AIAA-96-2736, 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 1996. [15]
D. M. Goebel, I. Katz, Fundamentals of Electric Propulsion: Ion and Hall Thrusters,
[16]
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first ed., Wiley, 2008.
T. Shoenherr, R. Kawashima, H. Koizumi, K. Komurasaki, D. Fujita, Y. Ito: Design
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and Performance Evaluation of Thruster with Anode Layer UT-58 for High-Power Application, IE`PC-2013-242, 33rd International Electric Propulsion Conference, 2013. [17]
R.R. Hofer, A. D. Gallimore, Efficiency Analysis of a High-Specific Impulse Hall
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Thruster, AIAA-2004-3602, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 2004.
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Figure Captions:
Fig. 1 Cross section of the anode-layer-type Hall thruster at the University of Tsukuba.
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Fig. 2 Normalized magnetic flux density Fig. 3 Schematic of the measurement system. Fig. 4 Thrust efficiency as a function of the xenon fraction. Fig. 5 Mass utilization efficiency as a function of the xenon fraction. Fig. 6 Specific impulse as a function of the xenon fraction. Fig.7 Thrust as a function of the xenon fraction (without extended channel).
11
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Fig. 8 Cross section of the anode-layer-type Hall thruster with an extended discharge
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channel, at the University of Tsukuba.
Table captions:
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Table 1 Operation conditions.
12
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Fig. 1 Cross-section of the anode-layer-type Hall thruster at the University of Tsukuba.
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Thruster exit plane (with extended channel)
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1 without Extended channel with Extended channel
0.8
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Normalized magnetic flux density
Thruster exit plane (without extended channel) Anode tip 1.2
0.6 0.4 0.2
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0 -15
-10
-5
0
5
10
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Axial potition, mm
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Fig. 2 Normalized magnetic flux density
15
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Fig. 3 Schematic of the measurement system.
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60
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40 30 20 10
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Thrust efficiency, %
50
without Extended channel with Extended channel
0
20
40 60 Xenon fraction, %
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0
80
100
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Fig. 4 Thrust efficiency as a function of the xenon fraction.
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80
60
40
without Extended channel (Xe correction) without Extended channel (Ar correction) with Extended channel (Xe correction) with Extended channel (Ar correction)
20
0
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Mass Utilization Efficiency, %
100
0
20
40
60
80
100
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Xenon fraction, %
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Fig. 5 Mass utilization efficiency as a function of the xenon fraction.
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2500
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1500
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1000
without Extended channel with Extended channel
500 0
0
20
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Specific Impulse, s
2000
40 60 80 Xenon fraction, %
100
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Fig. 6 Specific impulse as a function of the xenon fraction.
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40
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35
25
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20 15 10 5 0
0
20
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Thrust, mN
30
40
60
80
100
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Xenon fraction, %
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Fig. 7 Thrust as a function of the xenon fraction (without extended channel).
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9 mm
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Fig. 8 Cross-section of the anode-layer-type Hall thruster with an extended discharge channel,
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at the University of Tsukuba.
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Values
Propellant
argon-xenon gas mixture
Total flow rate
27.9 sccm
Xenon fraction
0 ~ 100 %
Discharge voltage
200 V
Magnetic flux density
10 ~ 50 mT
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Parameters
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Table 1 Experimental Conditions
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Thrust efficiency of argon propellant Hall thruster can be improved by ionization enhancement by mixing of xenon.
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Longer discharge channel is effective to promote argon ionization.
S0042-207X(18)30244-6
DOI:
10.1016/j.vacuum.2018.09.042
Reference:
VAC 8255
To appear in:
Vacuum
Received Date: 10 February 2018 Revised Date:
20 September 2018
Accepted Date: 21 September 2018
Please cite this article as: Yamasaki J, Yokota S, Shimamura K, Performance enhancement of an argonbased propellant in a Hall thruster, Vacuum (2018), doi: https://doi.org/10.1016/j.vacuum.2018.09.042. This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting proof before it is published in its final form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain.
ACCEPTED MANUSCRIPT
Title:
Performance Enhancement of an Argon-Based Propellant in a Hall
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Thruster Authors:
Junko Yamasaki a *, Shigeru Yokota b , and Kohei Shimamura b
Department of Engineering Mechanics and Energy, University of Tsukuba
M AN U
a
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Affiliations:
1-1-1 Tennodai, Tsukuba, Ibaraki 305-0006, Japan b Division
of Engineering Mechanics and Energy, University of Tsukuba
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1-1-1 Tennodai, Tsukuba, Ibaraki 305-0006, Japan
*Corresponding Author: Junko Yamasaki
EP
Department of Engineering Mechanics and Energy, University of Tsukuba
AC C
1-1-1 Tennodai,
Tsukuba, Ibaraki 305-0006, Japan Tel: 81-29-853-5138
Email: [email protected]
1
ACCEPTED MANUSCRIPT
Abstract: Finding alternatives to xenon as a propellant is essential for reducing the launching cost of
RI PT
all-electric satellites and for achieving low-cost electric propulsion, especially using Hall thrusters. Argon is a candidate as such a low-cost propellant, but it has never been used practically owing to its poor performance. We attempted two methods to improve the thruster performance: (1) by
SC
mixing xenon in argon propellant and (2) by modifying the thruster structure for the argon propellant. Modifying the thruster structure yielded an increase in the specific impulse from 482 to
M AN U
1500 s using pure argon. Also, when we used an argon-xenon mixture gas as propellant, the thruster performance increased under argon-rich conditions using the modified thruster.
Keywords:
AC C
EP
TE D
Electric Propulsion, Hall Thruster, Alternative propellant
2
ACCEPTED MANUSCRIPT
1. Introduction One of the most effective ways to economize on space transportation is to replace
RI PT
conventional chemical propulsion by electric propulsion. For example, a Hall thruster, one of the most promising electric propulsion systems, can save $13 million on the cost of launching a geostationary orbit satellite.[1] Anode-layer-type Hall thrusters are currently
SC
actively researched in Japan.[2][3]
On the other hand, Hall thruster propellants, consisting of xenon gas, are much more
M AN U
expensive than those used for chemical propulsion. To address this problem, alternative propellants have been tested. [4-13] Argon is cheap but is not considered to be a strong candidate because of its poor thruster performance [6], which increases the effective launching cost.
TE D
The low thrust performance of argon stems from its higher ionization energy and lighter weight, which make a large amount of input electric power go into ionization rather than thrust power. Argon is therefore rapidly exhausted from the ionization channel
EP
because of its higher thermal velocity. [6] Therefore, we suggest enhancing argon ionization using small amount of xenon. It is expected that electrons produced by xenon
AC C
ionization collide with neutral argon particles, thereby increasing the mass utilization efficiency of argon.
The lower ionization energy of xenon in the mixture is expected to facilitate argon
ionization. In this study, we measured the thrust performance by applying propellant mixtures with various mixture ratios of argon and xenon. In addition, we tested the thrust performance with an extended discharge channel, which is also expected to enhance
3
ACCEPTED MANUSCRIPT
ionization.
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2. Mixture propellant effect 2.1 Experimental procedure
Figures 1 shows a cross section of the sub-kilowatt class anode-layer-type Hall
SC
thruster developed at the University of Tsukuba.[13] The inner and outer diameters of the discharge channel are 33 and 43 mm, respectively. The discharge channel length is 3 mm.
M AN U
The magnetic field in the acceleration channel is applied by one inner coil and three outer coils. To stabilize the operation, the magnetic flux density near the anode was shielded by the magnetic screen. The normalized magnetic field profiles are shown in Fig.2. In this chapter, we used the thruster “without extended channel”.
TE D
Figure 3 shows a schematic diagram of the experimental setup. The Hall thruster was operated in a vacuum chamber whose diameter and length were both 1.0 m. The vacuum chamber was evacuated by a turbo molecular pump (3,000 L/s) backed by a rotary pump
EP
(1,500 L/min). The background pressure varied depending on the xenon mixture ratio and the backpressure increased gradually with the xenon fraction. The backpressure reached a
AC C
maximum of 3.8×10-2 Pa in pure xenon, and a minimum of 2.6 ×10-2 Pa in pure argon. The thrust was measured by a pendulum-type thrust stand, also developed at the
University of Tsukuba. Only knife-edge joints were used between the components. To measure the discharge current Id, the voltage between both ends of a 0.5 Ω metal-film resistor (inserted between the anode and discharge power supply) was measured with an oscilloscope using a differential probe. The ion beam current Ib was measured using the ion
4
Figure 1 Figure 2 Table 1
ACCEPTED MANUSCRIPT
collector, a 250×300 mm copper plate biased at -20 V with respect to the cathode. The collector was placed 30 cm downstream from the thruster. Another 0.5 Ω metal-film resistor
RI PT
was used for the current measurement. The collected ion current must be corrected to eliminate charge-exchange (CEX) collisions between the thruster and the ion collector. Owing to a lack of Ar-Xe+ and Xe-Ar+ CEX collision cross-section data, Ib was corrected
SC
using only the Ar-Ar+ collision cross section data or only the Xe-Xe+ collision
M AN U
． cross-section data. From Ib, the mass utilization efficiency (ηu=miIb/em) was deduced. Here, ． mi, e, and m are the ion mass, elementary charge, and mass flow rate, respectively. The results for ηu, corrected using the Ar-Ar+ data and Xe-Xe+ data separately, are plotted in Fig. 4.
The operation conditions are tabulated in Table 1. The total mass flow rate was 27.8
TE D
sccm. The xenon fraction α was varied from 0 to 100%. The discharge voltage was fixed at 200 V. The magnetic flux density was set to maximize the thrust efficiency ηt for each xenon fraction. For example, a peak value of 60 mT was applied in the case of pure xenon,
EP
and 10 mT was applied in the case of argon. Here, the thrust efficiency of an electrically
AC C
． powered thruster is defined as ηt = F2/mP, where, F and P are the thrust and input electrical ． power, respectively. [15] Also, we deduced the specific impulse Isp = F/mg using the measured thrust in order to discuss the “propellant consumption efficiency”. 2.2 Result and Discussion The thrust efficiency and the mass utilization efficiency are shown in Figs. 4 and 5, respectively, as functions of the xenon fraction. The error bars represent the standard deviations in the data. This chapter discusses only the normal case (denoted by circles); the
5
ACCEPTED MANUSCRIPT
extended discharge channel (square symbols) is explained in the next chapter. The thrust efficiency increases with the xenon fraction (Fig. 4). The thrust efficiency
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was less than 5% for pure argon but 48 % for the pure xenon case, consistent with earlier studies [12][14][16]. The dominant factor behind the ηt increase is the increase in ηu with increasing xenon fraction (Fig. 5). Thus, a low thrust efficiency when using argon
SC
propellant can be improved by enhancing ionization via a mixture with a low-ionization-energy substance like xenon. For simplicity, we here consider only singly
M AN U
charged ions. According to reference [17], the fraction of multiply charged ions is relatively small: the Xe2+ and Xe3+ fractions are 3% and 1%, respectively, for a discharge voltage of 300 V. Therefore, the contribution of multiply charged ions to ionization enhancement by xenon is small.
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The specific impulse Isp is shown in Fig. 6 as a function of the xenon fraction, displaying an increase with the xenon fraction and attaining a maximum when the xenon ． ． fraction is 70%. As mentioned above, Isp is defined as F/mg. Here, obviously m is
EP
monotonically increasing functions of α. On the other hand, measured F was also monotonically increasing functions of a (see Fig. 7). As a consequence, Isp is a upward-convex function of α.
AC C
Figure 3 Figure 4
We note that Isp for pure xenon reaches 1500 s, approximately 100 s more than with
other anode-layer-type Hall thrusters [14], so that Isp in the xenon-rich case may be overestimated because of the effect of the background pressure. On the other hand, argon is assumed not to contribute significantly to an increase in background pressure because of its higher pumping speed. Therefore, the estimate of Isp in the argon-rich case is likely to be
6
Figure 5
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3. Ionization enhancement by an extended discharge channel 3.1 Hall thruster with an extended discharge channel
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more correct than under xenon-rich conditions.
As discussed above, a greater Isp is achieved using argon gas rather than pure xenon.
SC
However, a higher ηt under argon-rich conditions is needed to reduce the propellant cost. Therefore, the discharge channel was extended to enhance argon ionization. The
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ionization-collision probability is thus expected to be higher because the propellant channel transit time becomes longer. Figure 8 shows the Hall thruster with the extended discharge channel. The distance between the anode tip and thruster exit plane is 9 mm. The magnetic flux density distribution is the “with extended channel” case in Fig.3. The experimental
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procedure is described above.
3.2 Result and discussion
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The open squares in Figs. 4 and 5 display ηt and ηu, respectively, as functions of
α when using the extended discharge channel. As shown in Fig. 4, ηt is improved using the
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extended discharge channel under argon-rich conditions (α < 50%) thanks to the enhancement of ionization resulting from the extended discharge channel, as intended. There is also an improvement in ηu when using the extended discharge channel (Fig. 5). On the other hand, as a result of the discharge channel extension, ηt decreases under xenon-rich conditions (α > 50%) despite ηu increasing in all the mixtures. This result indicates that the extended discharge channel is too long for the xenon propellant; a 7
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proportion of ionized xenon collides with the long-channel wall before exhaust release, so that the extra energy is used for the re-ionization of the recombined xenon atoms.[6]
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Figure 6 plots Isp as a function of α. Especially under argon-rich conditions, Isp increases to 1500 s over the entire α range when the discharge channel is extended. Because argon is lighter than xenon (39.95 amu, compared to 131.29 amu for xenon), the
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greater the increase in α, the higher the atomic mass and the lower the exhaust speed. Therefore, Isp increases much more than the efficiency in the case of an extended discharge
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channel because Isp is also defined as v/g, where g is the acceleration due to gravity and v is the exhaust speed. Moreover, considering the back-pressure effect mentioned above, it is expected that Isp under argon-rich conditions is greater than under xenon-rich conditions. This means that an argon or an argon-rich mixture is more advantageous than xenon in
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terms of cost for space transportation.
Whereas the extended discharge channel is effective for achieving ionization enhancement, it is likely to shorten the thruster lifetime owing to wall erosion. Optimizing
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future work.
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the channel length for low-ionization-energy gas propellants thus remains an objective for
5. Conclusion
(1) The thrust performance of the Hall thruster using an argon-xenon gas mixture was investigated. The thrust efficiency and the mass utilization efficiency increased with increasing xenon fraction in the gas mixture. The specific impulse reached a maximum for a
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xenon fraction of approximately 70 % as a result of the trade-off between ion mass and thrust efficiency.
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(2) We tested the thrust performance using an extended discharge channel, which is expected to enhance ionization. The thrust efficiency increased by up to 300 % under argon-rich conditions, but decreased under xenon-rich conditions. This result indicates that
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the longer discharge channel increases not only the ionization rate but also the rate of ion loss to the wall. As the result of the improvement in the thrust efficiency, the specific
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impulse reached approximately 1500 s regardless of the xenon fraction.
Acknowledgements
References [1]
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This work was supported by JSPS KAKENHI Grand Number 15H0558.
K. Kinefuchi, N. Nagao, Y. Saito, K. Okita, K. Kuninaka. Test Facility Concept for
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High Power Electric Propulsion, STEP-2013-024, Space Transportation Symposium FY2013, 2014. (in Japanese) Y. Hamada et al. Hall thruster development for Japanese space propulsion programs,
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[2]
Transactions of JSASS, 60 (2017) 320-326.
[3]
K. Komurasaki, Y. Arakawa, Definition of Hall Current Ion-Thruster Performance, Journal of Propulsion and Power 8 (1992) 1212-1216.
[4]
J. A Linnell, A. D. Gallimore, Efficiency Analysis of a Hall Thruster Operating with Krypton and Xenon. J. Propul. Power 22 (2006) 1402–1412.
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[5]
Lorand, O. Duchemin, S. Zurbach, D. LE. Mehaute, N. Cornu, Alternate propellants for PPS Hall-Effect Plasma Thruster. IEPC-2013-074, 33rd International Electric
[6]
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Propulsion Conference, 2013.
Eunsun Cha1, David B. Scharfe, Michelle K. Scharfe, and Mark A. Cappelli, Hybrid Simulations of Hall Thrusters Operating on Various Propellants.
[7]
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IEPC-2009-075, 29th International Electric Propulsion Conference, 2009.
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Conference, 2017. [8]
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and Plum Experiments, IEPC-2017-25, 35th International Electric Propulsion
M. A. Hopkins and L. B. King, Performance Comparison Between a Magnesiumand Xenon-Fueled 2 Kilowatt Hall Thruster, J. Propul. Power 32 (2016) 1015-1021. J. Szabo, M. Robin, J. Duggan, Light Metal Propellant Hall Thrusters,
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[9]
IEPC-2009-138, 31st International Electric Propulsion Conference, 2009. [10]
Adam Shabshelowitz, Alec D. Gallimore, Peter Y. Peterson, Alternative propellant
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Performance of a Helicon Hall Thruster Operating with Xenon, Argon, and Nitrogen, Journal of Propulsion and Power, 2014, Vol.30: 664-671. James Szabo, Bruce Pote, Surjeet Paintal, Mike Robin, Adam Hillier, Richard D.
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[11]
Branam, Richard E. Huffmann, Performance Evaluation of an Iodine-Vapor Hall Thruster, Journal of Propulsion and Power, 2012, Vol.28: 848-857
[12]
A.Lorand, O.Duchemin, S. Zurbach, D.Mehaute , N. Cornu, Alternate propellants
for PPS® Hall-Effect Plasma Thruster, IEPC-2013-074, 2013 [13]
J. Yamasaki, S. Yokota, K. Shimamura, Operation Characteristics of Anode Layer
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Hall Thruster Using Alternative Propellant, ISTS-2017-b-13, 31st International Symposium on Space Technology and Science. 2017. D. Manzella et al. Evaluation of Low Power Hall Thruster Propulsion,
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[14]
AIAA-96-2736, 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 1996. [15]
D. M. Goebel, I. Katz, Fundamentals of Electric Propulsion: Ion and Hall Thrusters,
[16]
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first ed., Wiley, 2008.
T. Shoenherr, R. Kawashima, H. Koizumi, K. Komurasaki, D. Fujita, Y. Ito: Design
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and Performance Evaluation of Thruster with Anode Layer UT-58 for High-Power Application, IE`PC-2013-242, 33rd International Electric Propulsion Conference, 2013. [17]
R.R. Hofer, A. D. Gallimore, Efficiency Analysis of a High-Specific Impulse Hall
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Thruster, AIAA-2004-3602, 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 2004.
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Figure Captions:
Fig. 1 Cross section of the anode-layer-type Hall thruster at the University of Tsukuba.
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Fig. 2 Normalized magnetic flux density Fig. 3 Schematic of the measurement system. Fig. 4 Thrust efficiency as a function of the xenon fraction. Fig. 5 Mass utilization efficiency as a function of the xenon fraction. Fig. 6 Specific impulse as a function of the xenon fraction. Fig.7 Thrust as a function of the xenon fraction (without extended channel).
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Fig. 8 Cross section of the anode-layer-type Hall thruster with an extended discharge
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channel, at the University of Tsukuba.
Table captions:
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Table 1 Operation conditions.
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Fig. 1 Cross-section of the anode-layer-type Hall thruster at the University of Tsukuba.
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Thruster exit plane (with extended channel)
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1 without Extended channel with Extended channel
0.8
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Normalized magnetic flux density
Thruster exit plane (without extended channel) Anode tip 1.2
0.6 0.4 0.2
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0 -15
-10
-5
0
5
10
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Axial potition, mm
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Fig. 2 Normalized magnetic flux density
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Fig. 3 Schematic of the measurement system.
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60
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40 30 20 10
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Thrust efficiency, %
50
without Extended channel with Extended channel
0
20
40 60 Xenon fraction, %
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0
80
100
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Fig. 4 Thrust efficiency as a function of the xenon fraction.
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80
60
40
without Extended channel (Xe correction) without Extended channel (Ar correction) with Extended channel (Xe correction) with Extended channel (Ar correction)
20
0
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Mass Utilization Efficiency, %
100
0
20
40
60
80
100
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Xenon fraction, %
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Fig. 5 Mass utilization efficiency as a function of the xenon fraction.
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2500
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1500
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1000
without Extended channel with Extended channel
500 0
0
20
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Specific Impulse, s
2000
40 60 80 Xenon fraction, %
100
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Fig. 6 Specific impulse as a function of the xenon fraction.
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40
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35
25
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20 15 10 5 0
0
20
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Thrust, mN
30
40
60
80
100
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Xenon fraction, %
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Fig. 7 Thrust as a function of the xenon fraction (without extended channel).
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9 mm
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Fig. 8 Cross-section of the anode-layer-type Hall thruster with an extended discharge channel,
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at the University of Tsukuba.
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Values
Propellant
argon-xenon gas mixture
Total flow rate
27.9 sccm
Xenon fraction
0 ~ 100 %
Discharge voltage
200 V
Magnetic flux density
10 ~ 50 mT
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Parameters
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Table 1 Experimental Conditions
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Thrust efficiency of argon propellant Hall thruster can be improved by ionization enhancement by mixing of xenon.
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Longer discharge channel is effective to promote argon ionization.