Hall Thruster Development for Japanese Space Propulsion Programs*

Yushi HAMADA,1) Junhwi BAK,1) Rei KAWASHIMA,1) Hiroyuki KOIZUMI,1) Kimiya KOMURASAKI,1)

Naoji YAMAMOTO,2)† Yusuke EGAWA,2) Ikkoh FUNAKI,3) Shigeyasu IIHARA,4) Shinatora CHO,5)

Kenichi KUBOTA,5) Hiroki WATANABE,6) Kenji FUCHIGAMI,7) Yosuke TASHIRO,4)

Yuya TAKAHATA,8) Tetsuo KAKUMA,8) Yusuke FURUKUBO,8) and Hirokazu TAHARA8)

1)Department of Aeronautics and Astronautics, The University of Tokyo, Tokyo 113–8656, Japan2)Department of Advanced Energy Engineering Science, Kyushu University, Kasuga, Fukuoka 816–8580, Japan

3)Institute of Space and Astronautical Science, JAXA, Sagamihara, Kanagawa 252–5210, Japan4)IHI Aerospace Co., Ltd., Tomioka, Gunma 370–2398, Japan

5)Research and Development Directorate, JAXA, Chofu, Tokyo 182–8522, Japan6)Department of Aerospace Engineering, Tokyo Metropolitan University, Tokyo 191–0065, Japan

7)IHI Corporation, Yokohama, Kanagawa 235–8501, Japan8)Department of Aeronautics and Astronautics, Osaka Institute of Technology, Osaka 535–8585, Japan

Three different types of high power Hall thrusters—anode layer type, magnetic layer type with high specific impulse,and magnetic layer type with dual mode operation (high thrust mode and high specific impulse mode)—have been devel-oped, and the thrust performance of each thruster has been evaluated. The thrust of the anode layer type thruster is in therange of 19–219mN, with power in the range of 325–4500W. The thrust of the high specific impulse magnetic layer typethruster was 102mN, with specific impulse of 3300 s. The thrust of the bimodal operation magnetic layer thruster was385mN with specific impulse of 1200 s, and 300mN with specific impulse of 2330 s. The performance of these thrustersdemonstrates that the Japanese electric propulsion community has the capability to develop a thruster for commercial use.

Key Words: Hall Thruster, In-space Propulsion

Nomenclature

F: thrustg: gravitational accelerationId: discharge currentIsp: specific impulse_m: mass flow rate

Pcoil: power consumption at coilsPcathode: power consumption at cathode

T/P: thrust to power ratioVd: discharge voltage�t: thrust efficiency

Subscriptsa: anodec: cathode

1. Introduction

High power electric propulsion (EP) is increasingly in de-mand as a main propulsion system for planetary explorationmissions1) and all-electric satellites.2) It would also be appro-priate for space cargo missions for the Manned Mars Mission

and construction of the Space Solar Power System. US,European, and Japanese research organizations and compa-nies are competing in the development of high power electricpropulsion.3–5) The most mature EP system, the ion enginesystem, has shown good performance at low power levels,from several hundredW to kW; the successes of the “DeepSpace 1”6) and the asteroid explorer “Hayabusa”7) haveshown the superiority of the ion engine system. Hall thrustersare also promising, and they have shown competitive per-formance against ion thruster systems8,9) and superior per-formance at higher power levels.10,11) Traditionally, thrust-to-power ratio, specific impulse, and hours of operation forHall thrusters are around 50–60 mN/kW, 1500–2000 s and6000–10000 h, respectively. The characteristics of the Hallthruster are its high thrust density and wide power range op-eration, with acceptable efficiency. The use of Hall thrustersin satellites yields lower trip times and lower system grossmass than the use of ion thrusters.12)

There are two major types of Hall thrusters; the magneticlayer type and the anode layer type.13,14) Many satellites withthe magnetic layer type Hall thrusters have been launchedand operated. In 2003, the SMART-1 mission presented byESA adopted a PPS-1350, which was operated successfully.Total impulse was 1.1 millionN0s and delta-V was 3.9km/s.15) A Hall thruster named Hall Effect Rocket with Mag-netic Shielding (HERMeS) is currently under developmentfor the US Solar Electric Propulsion demonstration mis-sion.16) This is a 12.5 kW, 3000 s specific impulse thrusterwith “magnetic shielding”17) technology. Magnetic shieldingis expected to extend the life expectancy of the Hall thruster.

Trans. Japan Soc. Aero. Space Sci.Vol. 60, No. 5, pp. 320–326, 2017

© 2017 The Japan Society for Aeronautical and Space Sciences+Presented at the 8th Asian Joint Conference on Propulsion and Power,March 16–19, 2016, Takamatsu, Kagawa, Japan, and the 60th SpaceSciences and Technology Conference, September 6–9, 2016, Hakodate,Hokkaido.Received 26 August 2016; final revision received 24 December 2016;accepted for publication 17 March 2017.†Corresponding author, yamamoto@aees.kyushu-u.ac.jp

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The other type of Hall thruster, the anode layer type, was de-veloped in Russia as a Thruster with Anode Layer (TAL).The D-55 is the most famous Russian anode layer type Hallthruster, but it has less operational time than the magneticlayer type Hall thrusters.18)

In Japan, there have been many studies on Hall thrusterssince the late 80’s,19) as reported by Kaufman.8) These stud-ies have focused on the improvement of thrust performancein both single stage and two stage Hall thrusters (Periplasma-tron ion source,20) microwave discharge ion source21)),understanding of ion production/acceleration/loss processes,and understanding of the plasma-wall interaction mechan-ism. In addition to experimental studies, since the 1990’s,numerical simulation codes have been developed to studythe physics behind the Hall thruster.22–26) These studies havedemonstrated that numerical simulation is a powerful tool forclarification of the physics in the system as well as for the de-velopment of Hall thrusters.

Based on this experience, several Hall thrusters haverecently been developed at the University of Tokyo,27,28)

Osaka University/Osaka Institute of Technology(OIT),29,30) Tohoku University,31) Nagoya University,32)

Gifu University,33) Kyushu University,34) Tokyo Metropoli-tan University (TMU),35) and The Institute of Space and As-tronautical Science (ISAS) in the Japan Aerospace Explora-tion Agency (JAXA).36) This work has not been confined touniversities and JAXA, industry has also been involved, incollaboration with universities; the IHI Corporation andOsaka University37) developed a 1 kW class Hall thruster,and Mitsubishi Electric Corporation38) developed a 250mNclass Hall thruster.

The understanding developed through this work, and thesuccessful demonstration of an all-electric propulsion satel-lite by Boeing,2) have inspired the development of highpower electric propulsion in Japan; three 5 kW class Hallthrusters are currently under development. One is a2–6 kW class dual mode operation Hall thruster developedby JAXA, IHI, IHI Aerospace Engineering Corporation(IA), and TMU. (2–6 kW is considered to be the most usefuloperational range for a variety of missions.) The second is a5 kW class anode layer type Hall thruster developed as partof the Robust Anode-layer Intelligent Thruster for the Japa-nese IN-space propulsion system (RAIJIN) project, com-prised of nine universities and JAXA.39) The third is a mag-netic layer type Hall thruster with high specific impulsedeveloped at OIT. In the present paper, we report on the cur-rent status of these thrusters.

2. Experimental Setup

Figure 1 shows the 5 kW class anode layer type Hallthruster, RAIJIN94, developed as part of the RAIJIN project.The inner and outer diameters of the acceleration channel are60mm and 94mm, respectively. An inner solenoid coil andfour outer solenoid coils create a predominantly radial mag-netic field in the acceleration channel. There is a trim coil,which can change the shape of the magnetic field configura-

tion. The thruster has a hollow annular anode, which consistsof two cylindrical rings, with a propellant gas fed throughthem. The gap between the tip of the anode and the exit ofthe acceleration channel is fixed at 3mm. A hollow cathode(Veeco, HECS) is used as an electron source. Tests were con-ducted in an ISAS/JAXA ion engine endurance test vacuumchamber,40) 2m diameter by 5m length, evacuated by fourcryogenic pumps (44,000 l/s for xenon), with the pressurekept below 6:6� 10�3 Pa (for xenon) during thruster opera-tion, with total mass flow rates of 13.5mg/s (140 sccm).Pressure was measured using an ionization gauge, whichwas positioned close to the top of the thruster behind ashroud. The chamber baseline pressure was below1� 10�5 Pa. High-purity (99.999%) xenon gas was used asthe propellant. Thrust measurement was performed using adual pendulum thrust stand.41) The uncertainty of the thrustwas estimated as 3% at the thrust level of 200mN, due tothe friction of the calibration system, thermal drift effect,and the magnetic interference between the thruster and thethrust stand/vacuum facility.

The THT-VI magnetic layer type Hall thruster for highspecific impulse, which was developed at Osaka Instituteof Technology, is shown in Fig. 2. The outer diameter ofthe discharge channel is 100mm and the inner diameter is56mm, that is, the channel is 22mm wide and 40mm long.These channel dimensions are the same as those of the Rus-sian SPT-100.42) The discharge channel wall is made of bor-

Inner coil

Trim coil

Outer coil

Anode

(a) (b)

Fig. 1. RAIJIN94 Hall thruster, (a) Photo, (b) Magnetic coil position.

(a) (b)

Fig. 2. Schematic of SPT-type Hall thruster THT-VI, (a) Overview, (b)Cross section.

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on nitride (BN). A hollow cathode (Veeco, HC-252) is usedas the electron source. Tests are conducted in the ISAS/JAXA ion engine endurance test vacuum chamber with xe-non as the propellant. The thruster mass flow rate is fixedat 3.0mg/s. The hollow cathode mass flow rate is 0.1mg/sexcept in high discharge voltage operation; it changes to0.2mg/s when the discharge voltage is 950–1000V for sta-ble operation. The inner coil current, outer coil current andback coil current are fixed at 0.3A, 0.3A, and 0.9A, respec-tively.

Dual mode operation magnetic layer type Hall thrusterbreadboard models were developed by JAXA, IA, IHI/IA,and TMU. For the requirements of the both high-thrust andhigh-Isp modes, we intend to put the ion production regionat the appropriate downstream position. The effective diam-eters of the channels are 100mm for BBM1a and 140mm forBBM1b, as shown in Fig. 3. A larger model, BBM2, 170-mm effective diameter was also fabricated and tested. Thedesign of the thruster head was based on pre-design numer-ical calculations; not only the geometric and magnetic fieldconfigurations, but also plasma production and loss mecha-nisms were quantitatively estimated prior to fabrication.43)

The use of numerical simulation in the preliminary designphase allowed optimization of the configuration and drasti-cally shortened the development time.

Performance tests were conducted in the IHI vacuumchamber (2m diameter and 3m length) and the vacuumchamber at the Georgia Institute of Technology (5m diame-ter and 9m length).44) In the experiment, the pressure insidethe vacuum chamber was kept below 4� 10�3 Pa, with a fewexceptions. In the experiment, xenon mass flow rates werechanged from 5 to 30mg/s for the anode, and the cathodemass fraction was kept to 10% of that of the anode. Dischargevoltages were set at 150 to 800V.

3. Results and Discussion

For evaluation of the Hall thruster performance, specificimpulse, Isp, and thrust efficiency, �t, are defined as

Isp ¼F

ð _ma þ _mcÞgð1Þ

�t ¼F2

2 _ma þ _mcð Þ VdId þP

Pcoil þ Pcathode� � ð2Þ

Figure 4 shows the thrust of the RAIJIN94 5 kW classanode layer thruster versus input power. The dependencyof the thrust on incident power shows some inconsistency,as the thrust was measured for various mass flow rates andvarious magnetic field configurations. The thrust is in therange of 19–219mN for power in the range of 325–4500W. The lower limit is imposed by low efficiency andthe upper limit is imposed by a combination of the cathodecapacity, instability, and overheating. As in conventionalHall thrusters, the thrust and power consumption are almostproportional to mass flow rate. The thrust is almost propor-tional to the square root of the discharge voltage, and powerconsumption is almost proportional to the discharge voltage;thrust is proportional to the square root of power consump-tion, if the mass flow rate is fixed.

Figure 5 shows the thrust versus coil current at a dischargevoltage of 300V, mass flow rate of 4.9mg/s, and trim coilcurrent of 0A. The thrust depends strongly on the magneticfield configuration, since the ion production and accelerationdepend strongly on magnetic field configuration.20) Thethrust at the inner coil current of 0.5A and the outer coil cur-rent of 0.84A goes to a minimum value, 71mN, in thesemeasurements. The thrust reaches maximum, 91mN, at innercoil current of 0.3A and outer coil current of 0.5A. There is a20mN difference, or about 25% of the average thrust, be-tween the two different magnetic field configurations, evenwhen the mass flow rate and discharge voltage are fixed.

Figure 6 shows the relation between trim coil current andthrust at the inner coil current of 0.4A, the outer coil currentof 0.4A, discharge voltage of 300V, and mass flow rate of4.9mg/s. The increase in the trim coil pushes the lines offorce toward the downstream region, as shown in Fig. 7.This will also push the plasma generation and accelerationregion toward the downstream end of the field. This would

(1) (2)

Fig. 3. Photo of breadboard models, (1) BBM1a, (2) BBM1b.

0 1000 2000 3000 4000 50000

50

100

150

200

250

2.9 mg/s 4.9 mg/s 6.8 mg/s 9.8 mg/s

Thru

st, m

N

Power consumption, W

Fig. 4. Thrust vs. power consumption for various mass flow rates for theRAIJIN94 thruster.

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tend to extend the lifetime of the thruster, though the thrustand discharge current might be decreased due to degradationof propellant utilization. This result shows that with increasein trim coil current, the thrust and discharge current decrease,as mentioned above, and the thrust achieves a maximum val-ue of 82 mN at the trim coil of ¹1A. The thrust reaches amaximum at trim coil of ¹1A, because with a decrease intrim coil current, the lines of force are attracted into the hol-low anode, and the ion production region moves upstream.This improves propellant utilization and thrust increases. Itshould be noted that excessive trim current does not affectthe plasma generation region due to the anode geometry.

These results show that magnetic field configurationgreatly affects both thrust and power consumption; optimiza-tion of the magnetic field is crucial, not only to the lifetime ofthe thruster17) but also to thruster performance.

Figure 8 shows the relation between power consumptionand thrust of the THT-VI at a mass flow rate of 3.0mg/s.

The magnetic coil ratio and magnitude are optimized at thedischarge voltage of 750V and mass flow of 3.0mg/s. Themagnetic coil ratio and magnitude are fixed to demonstratethe dependency of current and thrust on discharge voltagein a fixed magnetic field configuration. Thrust is almost pro-portional to the square root of the input power, as in conven-tional Hall thrusters, and thrust of 100mN was achieved atdischarge voltage of 850V. THT-VI is a high specific im-pulse aimed Hall thruster and, indeed, the specific impulsewas 3:3� 103 s at discharge voltage of 850V, with goodthrust efficiency of 0.60. This is a sufficiently high specificimpulse to perform planetary exploration missions. TheTHT-VI shows stable operation below a discharge voltageof 900V, and it shows unstable operation above 900V, ow-ing to overheating of the thruster body; the outer acceleration

0.3 0.4 0.5 0.6 0.7 0.80.3

0.4

0.5

0.6

0.7

0.8

Inne

r coi

l cur

rent

, A

Outer coil current, A

70

75

80

85

90

95

Thrust, mN

Fig. 5. Thrust versus coil configurations.

0

1

2

3

4

5

-3 -2 -1 0 1 2 3 4 5 660

70

80

90

Dis

char

ge c

urre

nt, A

Thru

st, m

N

Trim coil current, A

Fig. 6. Thrust and discharge current for various trim coil configurations.

(a) (b) (c)

Fig. 7. Calculated magnetic field configuration for three values of trim coilcurrent (inner coil current of 0.4A and outer coil current of 0.4A), (a)Trim coil current of ¹1A, (b) Trim coil current of 1A, (c) Trim coil cur-rent of 4A (calculated using Magnum4.0, Field Precision LLC).

0 500 1000 1500 2000 2500 3000 3500 40000

20

40

60

80

100

120

Thru

st, m

N

Power consumption, W(a)

200 300 400 500 600 700 800 900 1000 11000.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Thru

st e

ffic

ienc

y

Discharge voltage, V

(b)

Fig. 8. THT-VI thruster thrust performance at mass flow rate of 3mg/s,inner coil current of 0.3A, outer coil current of 0.3A, and trim coil currentof 0.9A, (a) Thrust vs. power consumption, (b) Thrust efficiency vs. dis-charge voltage.

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channel wall was observed to be red hot. Overheating couldbe overcome by the adoption of a large radiation shield, in-crease in the thruster size, and/or the adoption of a pyrolyticcarbon anode. Maximum efficiency is 0.60 at discharge volt-age of 750V; thrust performance would be improved if themagnetic field configuration (inner coil current, outer coilcurrent and trim coil current) were optimized for the dis-charge voltage.

The thrust of the breadboard models developed at IHI/JAXA/IA/TMU, the BBM1a, BBM1b and BBM2 thrusters,is plotted against input power in Fig. 9. The magnetic fieldconfiguration was optimized for each condition (each massflow and each discharge voltage). These thrusters are de-signed for dual mode operation, that is, high thrust modefor quick orbital transfer from GTO to GEO, and high specif-ic impulse mode for north-south station keeping (NSSK).The target thrust of the high thrust mode is 320–430mN,with specific impulse of 1180 s or higher. The second targetthrust is 140mN at power consumption of 3 kW with specificimpulse of 2500 s or higher. BBM1b demonstrates good bi-modal operation; thrust of 354mN with specific impulse of1100 s at power consumption of 5140W and mass flow rateof 30mg/s, and 130mN with specific impulse of 2350 s atmass flow rate of 5mg/s and power consumption of3500W. BBM2 also demonstrated dual mode operation;thrust of 380mNwith specific impulse of 1200 s at mass flowrate of 30mg/s, and thrust of 140mN with specific impulseof 2050 s at mass flow rate of 6mg/s and 300mN with spe-cific impulse of 2330 s at mass flow rate of 12mg/s. Theupper limit of input power is defined by thermal issues andfacility restrictions, the thrust range is limited to 136–473mN for BBM2 in the power range of 1500W to6970W. The thrust efficiency is in the range of 0.4 to 0.6,which is quite good as compared to the other Hall thrust-ers.14–17)

Thrust-to-power ratio (T/P) is a good indicator of thrusterperformance--large T/P is required for orbit transfer from

GTO to GEO or drag compensation at low altitude orbit.Figure 10 shows T/P for three thrusters versus specific im-pulse. The T/P is in inverse proportion to the specific im-pulse, if the thrust efficiency is constant. The BBM2 has ahigh T/P with specific impulse of 1000–2000 s and the max-imum T/P exceeds 70mN/kW. The BBM2 is a remarkablethruster, since it achieved a high thrust efficiency of 0.4 at alow specific impulse of 1050 s, and higher efficiency at high-er specific impulses.

Conventional Hall thrusters have a low T/P with specificimpulse of less than 1000 s, as a results of low thrust effi-ciency. this is due to low propellant utilization for dischargevoltages below 150V. The THT-VI also has a high T/P ratio,considering its specific impulse of 1800–3300 s. This is dueto high thrust efficiency (0.5–0.6). RAIJIN94 also showshigh T/P, with specific impulse of 1200 s to 2000 s.

These thrusters have a wide variety of T/P (from 30mNto 60mN) and specific impulse (from 1800 s to 3300 s),and they demonstrate capable performance for practical ap-plications, as shown in Table 1. Lifetime assessment of thethruster head and development of a neutralizer will be neces-sary next steps.

4. Conclusion

Several 5 kW class Hall thrusters, RAIJIN94, THT-VI andBBM1a, BBM1b and BBM2 have been developed at UT/KU, OIT, and JAXA/IA/IHI/TMU, and the thrust perform-ance evaluated. The specific impulse was found to be in therange of 1000 s to 3300 s with sufficient thrust efficiency(0.4–0.6) and thrust to power ratio (30–78mN/kW). The

0

100

200

300

400

500

0 2000 4000 6000 8000 1 104 41.2 10

BBM1a-6mg/sBBM1a-10mg/s

BBM1b-6mg/sBBM1b-12mg/sBBM1b-20mg/sBBM1b-30mg/sBBM2-6mg/sBBM2-12mg/sBBM2-20mg/sBBM2-30mg/s

Thru

st,

mN

Power Consumption, W

Fig. 9. Thrust characteristics of breadboard models (BBM1a, BBM1b andBBM2; mass flow rate is for anode mass flow rate).45)

0 500 1000 1500 2000 2500 3000 35000

10

20

30

40

50

60

70

80

90 η=0.6η=0.5

THT VI RAIJIN 94 IHI/JAXA/TMU_BBM2Th

rust

to p

ower

ratio

, mN

/kW

Specific impulse, s

η=0.4

Fig. 10. Thrust to power ratio versus specific impulse.

Table 1. Comparison of thrusters.46)

Thrusters Power, kW Efficiency Specific impulse, s

RAIJIN94 4.5 0.53 2200THT-VI 2.7 0.60 3300IHI/JAXA/IA/TMU 4 0.40 1050BBM2 7 0.54 2300SPT-140 4.5 0.55 1800BPT-4000 4.5 0.57 2000

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thrust to power ratio of BBM2 was 78mN/kW, with specificimpulse of 1050 s; this is a remarkable thrust to power ratioand it would be sufficient to shorten the transition from GTOto GEO. The specific impulse of THT-VI achieved 3300 swith a high thrust efficiency of 0.60; this performance rivalsthat of other high specific impulse Hall thrusters. RAIJIN94shows good performance, and thrust efficiency of 0.5–0.6with specific impulse of 1200–2200 s. Hall thruster develop-ment in Japan has been highly successful, and has led to thecurrent state-of-the-art. The stage is set for technological andapplication breakthroughs in the near future.

Acknowledgments

The authors gratefully acknowledge support from JAXA, IA andIHI. Research results were obtained using the ion engine endurancetest facility at ISAS, JAXA. The contributions of Prof. HitoshiKuninaka, Prof. Kazutaka Nishiyama, Prof. Satoshi Hosoda, Prof.Ryudo Tsukizaki, and their staff are greatly appreciated. For theexperiment at GIT, the contributions of Prof. Mitchell Walker andhis staff are much appreciated.

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Y. OhkawaAssociate Editor

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