S2-6
Small Current-High Voltage Steady-state Operation of MPD Arcjet
定常MPDアークジェットの小電流高電圧作動
Daisuke Ichihara, Hisashi Kataoka Tomoki Uno, John Jeong, Akira Iwakawa and Akihiro Sasoh
市原大輔, 片岡久志, 鵜生知輝, 鄭栽勲, 岩川輝, 佐宗章弘
Department of Aerospace Engineering, Nagoya University
Furo-cho, Chikusa-ku, Nagoya 464-0073, Japan
名古屋大学 工学研究科 航空宇宙工学専攻 〒464-8603 名古屋市千種区不老町
A steady state, applied field, rectangular magnetoplasmadynamics (MPD) thruster with permanent magnets
has been developed. The thrust and discharge voltage were measured by varying propellant mass flow rates,
discharge current and inter-electrode distance. Setting cathode at exit of the discharge channel, thrustincreased and 80-116% of the ideal Lorentz force obtained. Decreasing propellant mass flow rate, the ratioof back EMF to the discharge voltage increased and 16% of the thrust efficiency at 2876sec of specificimpulse was obtained.

1. General
A steady state, magnetoplasmadynamic (MPD)
thruster is considered one of the most promisinghigh-power thrusters for in space electric
propulsion [1]. It has a structural simplicity, highflexibility in propellant species, mass flow rate andspecific impulse. However, the most important
problems in MPD thruster are low thrust efficiency
and short operation time.
A conventional applied field MPD thruster has a
coaxial discharge channel, which consisted of acentral cylindrical cathode and a concentric
axisymmetric anode [2]. On the other hand,Institute of Space and Astronautical Science (ISAS)
investigated 2-dimension like applied field MPD
thruster have investigated and tested itsperformance. Introducing rectangular discharge
channel, the Lorentz force trends in the same
direction as exhaust velocity. As a results, itfacilitates the analysis of thruster performance andit has the potential to get higher thrust performance[3]. Generally, the laboratory model high power
MPD thruster is operated in quasi steady state,
which inputting high current, more than 1 kA, forless than 1 msec . However, the applied field MPD
thruster doesn’t need such a high current. For
example, small current and high voltage canapproach high power operation.
Because of these factors, we developed a steadystate MPD thruster with a rectangular discharge
channel in which electrodes are placed on the upperand lower walls. For better ignition and longer
lifetime, a gas fed type hollow cathode is used. The
cathode is placed at near the discharge channel exit
to distribute discharge current downstream. The
purpose of this study is to investigate operational
characteristics of the rectangular MPD thruster
varying with propellant mass flow rate ṁ, discharge
current Jd and inter electrode distance H.
2.

Rectangular MPD thruster headFigure 1 shows the schematic of the developed
rectangular MPD thruster and its electrical circuit.
The rectangular discharge channel has a width of25mm, length in the acceleration direction is 60mm.
The inter-electrode distance H, is variable. The
upper wall is water cooled anode made of copper.
On the lower wall, a hollow cathode (DLHC-1000,Fig. 1 Schematic of the developed rectangular MPD thruster and electrical circuit

40Thrust , mN

Kaufman & Robinson Inc.) is flush mounted. The
keeper electrode is made of 10mm length tantalum
plate and cathode tip is set at 5mm upstream fromdischarge channel exit. An external magnetic field
applied in the discharge channel is provided by twoneodymium magnets (50mm height, 50mm length
in the acceleration direction and 25mm width)connected with an iron yoke. The magnetic fieldstrength was 0.20 T at discharge region.

Mass flow rate , mg/s
1.3
30
1.7
2.1
F=JdBH
20
1003.

Thrust efficiency , %20
Mass flow rate , mg/s
1.3
1.7
2.115151010
5.0 mN/kW

500500 1000 1500 2000 2500 3000

Specific impulse , sec
Fig. 2 Thrust efficiency as a function of specific impulse(Jd=5.0-15A, B=0.20T, H=10,15mm)

010203040

Ideal Lorentz force , mNFig. 3 Thrust vs ideal Lorentz force (Jd=5.0-15A,B=0.20T, H=10,15mm)
20Back EMF ratio , %Experimental results
The thruster performance is summarized in Fig. 2.
Here, the specific impulse Isp and thrust efficiency ηare given byF(1)
I sp m gF2Fu BH
(2)
η


2m J d V d J d BH Vd
Generally speaking, thrust efficiency is anincreasing function of specific impulse. The lower
propellant mass flow rate, the higher specificimpulse range with high thrust efficiency becomes.
The maximum thrust efficiency was 16% at specificimpulse was 2876 sec.
Figure. 3 shows the relation between the ideal
Lorentz force JdBH, and measured thrust F. In casethat cathode was placed at center of the discharge
channel, thrust was lower than the ideal Lorentz
force [4]. However, setting cathode at near thedischarge channel exit, thrust increased and80-116% of the ideal Lorentz force obtained. Thisis because the discharge current distributed outsideof the discharge channel and thrust will increase byincreasing the magnetic field strength at outside ofthe discharge channel.
Subsequently, the ratio of back EMF to discharge
voltage ūBH/Vd is shown in Fig. 4. As decreasing
propellant mass flow rate, ūBH/Vd increased. This isbecause that effective exhaust velocity increased as
decreasing propellant mass flow rate.

15Inter-electrode distance , mm
1015105
0
0.0

0.5

1.0

1.5

2.0

2.5

Mass flow rate , mg/s
Fig. 4 Back EMF ratio vs mass flow rate (Jd=10A,B=0.20T)

4.

Conclusion
By setting cathode at near the discharge channel
exit, thrust increased and 80-116% of the ideal
Lorentz force was obtained. As decreasing
propellant mass flow rate, the ratio of back EMF todischarge voltage increased. Decreasing mass flowrate also increased thrust efficiency and maximum
thrust efficiency 16% was obtained at specificimpulse was 2876 sec.
References[1]. K. Sankaran, L. Cassady, A. D. Kodys and E. Y.
Choueiri: Annals of the New York Academy ofSciences. 1017 (2004). 450.
[2]. G. Krulle, M. A. Kurtz and A. Sasoh: J.P.P. 14(1998). 754.
[3]. A. Iwakawa, D. Nakata, and H. Kuninaka: Proc.
31th International Electric Propulsion Conference,Michigan, 2009, IEPC-2009-226.
[4]. S. Yokota, D. Ichihara, H. Kataoka, S. Harada andA. Sasoh: Proc. 33th International Electric
Propulsion Conference, Washington, 2013,
IEPC-2013-246.