Table of Contents
1. Introduction
1.1 Microsatellite……………………………………………….......................
1.2 Required performance……………………………………………………...
1.3 The Thruster for Microsatellites ...…………………..................................
1.4 Purpose …………….…………………………............................................
2. Miniature Microwave Discharge Plasma Thruster
2.1 Mechanism of ion acceleration………………………………………........
2.2 Selection of the Propellant………………………………………………...

1
1
3
5
7
8
8
9

2.3 Expectation of the Performance…………………………………………... 10
2.4 Microwave Discharge……………………………………………………... 11
2.5 Magnetic Tube…………………………………………………………….. 12
3. Experimental Facility and Setup
3.1 Vacuum Chamber ...……………………………………………………...
3.1.1 Small Vacuum Chamber……………………………………………..
3.1.2 Large Vacuum Chamber……………………………………………..
3.1.3 Space Science Chamber……………………………………………..
3.2 Miniature Microwave Discharge Thruster...……………………………..

14
14
14
15
15
17

3.3 Probe Measurement………………...…………………………………….
3.3.1 Langmuir probe……………………………………………………...
3.3.2 Emissive probe………………………………………………………
3.4 Retarding Potential Analyzer…………………………………………….
3.5 Ion beam measurement…………………………………………………..
3.6 Thrust measurement……………………………………………………...
3.6.1 Hanging Pendulum Thrust Stand…………………………………….
3.6.2 Aluminum Target Thrust Stand………………………………………

20
20
22
24
26
27
27
27

4. Experimental Result and Discussion
4.1 Ion Energy Measurement……..…………………………………………..

29
29

4.2 Plasma Diagnostics by Langmuir Probe………………………………….
4.3 Space Potential Measurement by Emissive Probe………………………..
4.4 Thrust Performance Estimation…………………………………………..
4.5 Thrust Measurement……………………………………………………...
4.5.1 Measurement by Hanging Pendulum Thrust Stand………………….
4.5.2 Measurement by Aluminum Target Thrust Stand……………………

30
33
36
39
39
41

5. Summary

45

List of References
Acknowledgement

45
48

Chapter 1
Introduction
1.1 Microsatellite
Nowadays, space utilization is entwined with human life in many aspects;
communication, weather forecast and car navigation system. On space utilization,
microsatellite is attractive. Microsatellites account for an increasing share of satellite
launches; 90% of the satellites launched in 2013 were in the 50 kg or less class. The Fig.
1-1 shows the number of launch of the 50 kg or less class satellites in 2004  2013.1)
Thanks to their relatively low cost and short development time, and the possibility of thelaunching piggyback on other payloads, 2,3) microsatellites enable space access moreachievable for academic teams, small companies, and developing countries.4,5)
Though the capabilities of microsatellites are restricted compared with large satellites,
microsatellites can be competitive with the large satellites if they adopt formation flightand constellation operation.6,7) In these configurations, the damage of a trouble can be
reduced and the addition of extra function is relatively easy.
Most microsatellites, however, have no thruster, since there is little room for the thrustsystems and a small solar array will not allow the large power consumption of a thruster.
Some challenging microsatellites, however, have a thruster like the “Hodoyoshi-3&4” and“PRoximate Object Close flYby with Optical Navigation (PROCYON)” in these days,8,9)
because thrusters enable advanced functionality, including the possibility of orbit transfer
or orbit keeping, or even super low altitude keeping. Super low altitude satellites flight at
about 180~250 km enable high resolution observation and reduce the emission power
requirements for active sensors.10,11) In addition, the observation device can beminiaturized, if super low altitude satellite realizes the resolution same to the one of the big
satellite in high altitude. Hence, super low altitude satellite can be miniaturized. However,micro super low altitude satellite has not been developed. An obstruction of theminiaturization is propulsion system. Satellites performing these missions must, however,compensate for drags by using thrusters. Therefore, the goal of this study is developmentof the thruster adoptable to the micro super low altitude satellite.
1

The number of launch

100
80
60
40
20

04

05

06

07

08

09

10

11

12

13

20

20

20

20

20

20

20

20

20

20

0

Year

Fig. 1-1 The number of launch of the 50 kg or less class satellites in these days2

1.2 Required performanceFirstly, we assumed the satellite and the mission as shown in Fig. 1-2 and Table. 1-1,
respectively. Generated power is estimated by the assumption that the area of the solar
panel is 0.4 m×0.5 m and the conversion efficiency of the solar panel is 40%. Secondly, we
estimate the magnitude of the drag. The drag is calculated by following equation.
1
F = 𝜌𝑣 2 𝐶𝑑 𝑆
2

(1-1)

In this equation, 𝜌, 𝑣, 𝐶𝑑 and 𝑆 represent the density of the air, the velocity of thesatellite, drag coefficient and the area of the projected area against the direction of
movement. These parameters are assumed that 𝜌 is 3.5×10-9 kg/m3 and 𝑣 is the velocityof the satellite at the altitude of 200 km (7.8 km/s). 𝐶𝑑 is changeable by the parameter of
the atmosphere and shape of the satellites. In the case of Super Low Altitude Test Satellite(SLATS), 𝐶𝑑 is 0.21 when the attack angle is 030 deg.12) If this parameter is adopted,
the drag is estimated to be 0.21 mN. Therefore, we set the objective thrust as 1 mN. The
power used for thrust system should be 2030% for widely application, so the objective
power consumption of the thruster system is under 20 W. Specific impulse is also
important parameters for thruster system. Specific impulse is defined in the following
equation and it represent the duration of the unit thrust by using unit propellant mass.15)
𝐼𝑠𝑝 =

𝐹
𝑣𝑖
≅
𝑚̇𝑔 𝑔

(1-1)

In this equation, F, 𝑚̇ and g represent the thrust and mass flow rate and gravitational
acceleration. The nearly equal is valid only when 100% of the propellant gas is utilized.
The unit of specific impulse (Isp) is s. Optimum specific impulse is estimated fromfollowing equation.13)(𝐼𝑠𝑝 )𝑜𝑝𝑡 =

√𝛼𝜏
𝑔

(1-2)

In this equation,  is the ratio of the thrust power and thruster system weight, and the unit
is W/kg. The weight of the thruster system is assumed to be 5 kg in reference to theminiature ion engine system for microsatellite.14) From these assumption,  and  is
calculated as following.
=

1[𝑚𝑁] × 7.8[𝑘𝑚/𝑠]= 1.56[𝑊/𝑘𝑔]
5[𝑘𝑔]

(1-3)

 is the operational time of the thruster and the unit is s.  is calculated as 6.31×107 sec
3from mission period. From these parameters, the optimum specific impulse is estimated as1000 sec. From the above, the required performance is are 1 mN and 1000 sec, respectively,
with a total power consumption of 20 W in the thruster system. In addition, the simplicityof the thrust system is also important. Simple structure provide the high robustness,compactness, light weight and low cost. These are very important point for microsatellite.
Therefore, the objective performance should be achieved with the simple thrust system.

Fig. 1-2 Image of the assumed satellites

Table. 1-1 The Properties of the assumed satellite
Altitude

200 kmSize

0.1 m×0.1 m×0.5 m

Mass

50 kg

Power generation

100 W

Mission period

2 years4

1.3 The Thruster for Microsatellites
There is the two types in the thruster for space craft; chemical propulsion and electric
propulsion. Chemical thruster gain a thrust by exhaustion of the high temperature and highpressure gas made by combustion of fuel and oxidizer. The Feature of the chemical
propulsion is high thrust and low specific impulse.
Electric Propulsion use the electric energy to expel the ionized propellant. Manysatellites operate with electric propulsion; there are 236 operational satellites as of 2013.16)
The feature of the electric propulsion is high specific impulse, so propellant mass can besaved by adoption the electric propulsion. Electric propulsion systems are suitable for the
drag compensation because the thrusters should operate while keeping the orbits and themounted propellant mass is restricted. “Gravity Field and Steady-State Ocean Circulation
Explorer (GOCE)” with a mass of 1100 kg, used two 20 mN class ion thrusters.17) “SLATS”
will also use an ion thruster. Therefore, we also adopted electric propulsion. However, it isdifficult to fulfill this thrust performance using existing electric propulsions. Some thruster
for micro/nano satellites is under developing, but there is no thruster which achieved
objective performance with simple structure.18),19) A power consumption will be over 10 W,
if 1 mN thrust is realized by a miniature ion thruster.20) For example, Table. 2 shows thebreakdown of the power consumption is a miniature ion thruster developed in Kyushu
University.21) In this power consumption, the thrust performance is that thrust and specificimpulse is 0.57 mN and 2972 sec, respectively. As shown in this table, the total power
consumption is 30 W. In addition, it is also hard that the specific impulse of 1000 sec isrealized by the existing thruster. Figure 1-3 shows the relationship between the thrust andspecific impulse of the thruster. As shown in this Figure, there is no thruster which has the
specific impulse of 1000 sec. Therefore, it is necessary that the development of a novel
thruster.

5

Fig. 1-3 The relationship of the thrust and the specific impulse

Table. 1-2 Breakdown of the power consumption in miniature ion thruster
Ionization in the ion thruster

8W

Ion acceleration

12 W

Ionization in the neutralizer

2W

Others (Conversion efficiency etc…) 8 W
Total power consumption

6

30 W

1.4 Purpose
Development of the thruster for the micro super low altitude satellite is our definitive
purpose. Hence, we have been developing a miniature microwave discharge thruster andsome parameters were optimized.22) Figure 1-4 shows the photo of the thruster. We
adopted a thruster concept as shown in Chapter 2 using a microwave discharge, which
doesn’t have the device to accelerate ions actively. This is because the main power
consumption of the ion thruster is used for ion acceleration as shown in Table. 1-2. If theion acceleration power is saved, the power consumption of the thruster head will 10 W
and total power consumption of the thrust system can be under 20 W. In addition, thepower supply for the grid is also saved. However, the mechanism of ion acceleration is
not shown yet and the thrust performance is not clear. Therefore, the purpose of this study
is to show the ion acceleration mechanism of the thruster and evaluate the thrust
performance exactly.

Fig. 1-4 Photo of the miniature microwave discharge thruster

7

Chapter 2
Miniature Microwave Discharge PlanaThruster
2.1 Mechanism of ion acceleration
Figure 2-1 shows a schematic of the ion acceleration mechanism of this thruster. The
plasma is generated by microwave discharge. Electrons are trapped by a magnetic field, in
the area between magnetic mirrors. This magnetic configuration is called the magnetic tube
and this configuration is used in some microwave discharge neutralizer and ion thruster
head.23) Electrons gain energy from the microwaves, they collide with propellant particles,
and the propellant atoms are ionized. The ions are accelerated by the potential difference
between the plasma inside the thruster and the space, if the potential of the dischargechamber was kept at the space potential. High energetic electrons can also go out todownstream with ions against the potential hill. These electrons and ions are neutralized
outside the thruster; it does not need a neutralizer. Therefore, this thruster is very simple
and small, with low power requirements.

Fig. 2-1 Mechanism of ion acceleration.
8

2.2 Selection of the Propellant
Xenon is widely used as propellant in electric propulsion, because the ionization cost
become low owing to a low ionization energy relatively in the rare gas. However, in thisstudy, Argon is used as propellant on account of high specific impulse and low cost.
Specific impulse is in inverse proportion to the mas of the propellant, so specific impulse
become longer by adopting Argon. The price of Argon is about 1/1000 of the one of Xenon,
because 1% of atmosphere is Argon. Table. 2-1 shows the cost and production of the Argon
and Xenon.24) Cost is the important for widely application of thrust system in the future.

Table. 2-1 Cost and production of Xenon and Argon
Production[t /year]

Potential production[t /year]

Unit price[JPY/kg]

Xe

32

160

500,000

Ar

3×106

5×106

500

9

2.3 Expectation of the Performance
It is needed to validate that the thruster concept can meet the required performance. The
potential difference between the plasma inside the thruster and the space is estimated as 20
V, because the sheath potential is 20 V on the simulation of a miniature microwave
discharge neutralizer in Kyushu University.25) In that case, the velocity of the extracted ion

𝑣𝑖 is calculated by following equation from the ion’s equation of motion.
𝑣𝑖 = √

2𝜑𝑒
𝑚𝑖

(2-1)

In this equation, 𝜑, 𝑒 and 𝑚𝑖 is represent the potential difference, elemental charge
and ion mass. From this equation, the ion velocity is 9800 m/s, and the specific impulse is1000 sec in the case of the propellant utilization of 1. Propellant utilization represent thepercent of the propellant extracted as ion beam current to the mass flow rate and defined in
the following equation.
𝜂𝑢 =

𝐼𝑏 𝑚𝑖
𝑒𝑚̇

(2-2)

𝐼𝑏 means the ion beam current. In this case, thrust is estimated as 1 mN from following
equation.
𝐹 = 𝑚̇𝑣 =

𝑚𝑖
𝐼 𝑣
𝑒 𝑏 𝑖

Therefore, this thruster concept can meet the objective performance.

10

(2-3)

2.4 Microwave Discharge
We adopted the microwave discharge for plasma generation. The utility of microwave
discharge is shown in the mission of HAYABUSA.26) Microwave discharge have the merits
for thruster shown in the below.

1) Long lifetime and simplification owing to the lack of the electrode2) Needlessness of the beforehand heating
3) Low power consumption by using existing electron

In this thruster, Electron Cyclotron Resonance (ECR) is used to give the energy to theelectron. In this thruster, however, it seemed that the electron cyclotron wave doesn’t
propagate because the size of the discharge chamber is much smaller than the wave lengthof the microwave. Electric field to accelerate electron seemed to be generated by thepotential difference of the discharge chamber and antenna. ECR is occurred by theoscillation of the antenna potential at microwave frequency. In this thruster, plasma is also
generated in front of the antenna.27) This shows that the electron can be accelerated enoughwithout ECR.

11

2.5 Magnetic Tube
In this thruster, the magnetic tube is formed between the central yoke and front yoke asshown in Fig. 2-2. Electrons make a round trip between the central yoke and front yokealong the magnetic lines of force and reflected at the both ends by the sheath potential.
Magnetic mirrors in the both ends also contribute the reflection of the electron. Electrons
gain energy from microwave by approaching the strong electric field area near the antenna
many times. Plasma can be generated effectively on account of this effect. In the magnetic
tube, guiding center also do B drift because the magnetic field strength is large and it
generate the B along the magnetic tube. Therefore, the total movement of electrons is asshown in Fig. 2-3.

Fig. 2-2 Confinement by magnetic tube

12

Fig. 2-3 B drift in the magnetic tube

13

Chapter 3
Experimental Facility and Setup
3.1 Vacuum Chamber
3.1.1 Small Vacuum Chamber
Experiments was conducted in the three types of the vacuum chamber. Fig. 3-1 shows
the small vacuum chamber in Kyushu University. The inner diameter and the length of thevacuum chamber are 0.6 m and 1.0 m, respectively. The pumping system consists of arotary pump (pumping speed 1.5×10-2 m3/sec) and turbo molecular pumps (pumping speed
5.2×10-1 m3/sec). The chamber baseline pressure is below 5.7×10-4 Pa. The back pressure is
2.7×10-2 Pa at the argon mass flow rate of 82 g/sec. This vacuum chamber is connected to
GND through whole experiments and this is the baseline voltage.

Fig. 3-1 Small vacuum chamber

143.1.2 Large Vacuum Chamber
Fig. 3-2 shows the large vacuum chamber in Kyushu University. The inner diameter andthe length of the vacuum chamber are 1.0 m and 1.2 m, respectively. The pumping system
consists of a rotary pump (pumping speed 2.7×10-2 m3/sec), mechanical booster pump
(1.0×10-1 m3/sec) and turbo molecular pumps (pumping speed 2.2 m3/sec). The chamberbaseline pressure is below 1.7×10-3 Pa. The back pressure is 3.6×10-3 Pa at the argon mass
flow rate of 82 g/sec. This vacuum chamber is connected to GND through wholeexperiments and this is the baseline voltage.

Fig. 3-2 Large vacuum chamber

3.1.3 Space Science Chamber
Fig. 3-3 shows the Space Science Chamber at the Institute of Space and Astronautical
Science (ISAS) of Japan Aerospace Exploration Agency. The diameter and the length ofthe vacuum chamber are 2.5 m and 5.0 m, respectively. The pumping system consists of aturbo molecular pump (pumping speed 3.4×103 l/s at N2) and two cryo pumps (pumping
speed 2.8×104 l/s at N2). The chamber baseline pressure is below 2.28×10-4 Pa. The backpressure is 3.28×10-4 Pa at an argon mass flow rate of 84 g/sec. The back pressure is 1.69
×10-2 Pa with the back ground plasma. The back ground argon plasma is generated bybackward diffusion type plasma generator and the density is 7×1011 - 1×1012 m-3. The
15

discharge current is 100-200 mA and the pressure is 0.02 Pa. Fig. 3-4 shows the photo of
the plasma generator.

Fig. 3-3 Space science chamber at ISAS

Fig. 3-4 Photo of the plasma generator

16

3.2 Miniature Microwave Discharge Thruster
In this study, the two types of the thruster is developed. Figures 3-5 (a) and (c) show thetwo microwave discharge thrusters; one is 50 mm×50 mm×27 mm and the other is 50
mm×50 mm×53.5 mm. Fig. 3-5(b) shows the parts assembly of the thruster with 12 mm
and discharge chamber length. Sm-Co magnets surround the discharge chamber and arelocated between yokes made of soft iron. The magnet size is 4 mm×4 mm×12 mm. The
inner diameter of the discharge chamber is 21 mm and its height is 12 mm. There are aspacer (diameter 36 mm) and an orifice plate (diameter 12 mm) in front of the front yoke.
The thickness of the spacer and orifice are 4 mm and 2 mm respectively. The antenna is
made of molybdenum. The antenna, which is star shaped, has a thickness of 1 mm and acircular diameter of 9 mm. This star shaped antenna showed best performance in theprevious study.29)
Figure 3-5 (d) shows the parts assembly of the thruster with 33 mm discharge chamber
length. The discharge chamber is extended to 33 mm. The magnets and the central yoke areextended to fit the discharge chamber. The back yoke is enlarged from 35 mm×35 mm×3
mm to 50 mm×50 mm×12 mm. The other yokes are the same dimensions as of the 12 mm
discharge chamber thruster.
The magnetic field strength inside the discharge chamber can be changed by changing
the number of magnets. Figure 3-6 shows the magnetic field distribution for 9 magnets (12mm discharge chamber) and 14 magnets (33 mm discharge chamber). The number ofmagnets is optimized for the best performance, the detail was shown in Ref. 30.

(a)

17

(b)

(c)

(d)
Fig. 3-5.

Microwave discharge thrusters (a) photo of 12 mm discharge chamber type, (b)parts assembly of 12 mm discharge chamber type, (c) photo of 33 mm discharge chamber
type, (d) parts assembly of 33 mm discharge chamber type.

18

(a)

(b)
Fig. 3-6. Magnetic field distribution for each discharge chamber (a) 12 mm
discharge chamber type and (b) 33 mm discharge chamber type.

19

3.3 Probe Measurement31)
3.3.1 Langmuir probe
Ion acceleration mechanism can be investigated by probe measurement. In this study,
two types of the probe is used; Langmuir probe and emissive probe. Figure 3-7 shows thephoto of the Langmuir probe used in this study. The Langmuir probe is installed in thelinear actuator and inserted to the thruster from the downstream on the center axis of thethruster as shown in Fig. 3-8.
In Langmuir probe measurement, plasma parameter is gained following procedure. The
I-V characteristic as shown in Fig. 3-9 is gained by the probe measurement. The horizontal
axis Vp shows the probe voltage and the horizontal axis Ip shows the current flowing into
the probe. The positive direction of the vertical axis corresponds to the electron current
flows into the probe. Ies, Iis, Vf and Vs, shows the electron saturation current, ion saturation
current, floating potential and plasma potential, respectively. This V-I characteristic divided
three regions. First region is the right region and this region is ion saturation current region.
In this region, the probe voltage is much lower than the plasma potential and electrons arereflected by potential difference and only ion flows through probe. Second region isretarding potential region and it is a middle region of the figure. In this region, probe
current increases with the increase of the probe voltage. This is because electrons begin toflow into the probe. Third region is an electron saturation region. In this region, all ion isreflected and only electrons flow into the probe. Plasma parameters can be founded asfollowing.

1) Electron temperature Te
In the retarding potential region, the following equation can be obtained.
𝐼 ( 𝑝)

=

𝑒

(3-1)

As shown in this equation, electron temperature can be obtained from the slope in the I-V
characteristic plotted in the semilog graph. Note that Ie is not equal to the Ip, so Ie isobtained by subtracting Ii form Ip. Ii in the retarding potential region can be obtained byextrapolation as shown (a) in Fig. 3-9.

202) Plasma potential
Plasma potential Vs is obtained as the voltage of the cross point of the tangent lines in
the retarding potential region and electron saturation current region on the semilog graph.

3) Electron density
Electron density is obtained by electron saturation current or ion saturation current. In
this study, electron saturation current is decreased by magnetic field, so electron density is
obtained from ion saturation current. Electron density is obtained from following equation.

𝐼𝑖𝑠 =

𝑒𝑆(

𝑘𝑏
) /2
𝑚𝑖

(

1
)
2

Fig. 3-7 Photo of the Langmuir Probe

Fig. 3-8 Schematics of the probe measurement
21

(3-2)

Fig. 3-9 I-V characteristic gained by Langmuir probe

3.3.2 Emissive probe
Emissive probe is used for potential measurement. Emissive probe can measure thespace potential more exactly than the Langmuir probe. Figure 3-10 shows the I-V
characteristic gained by emissive probe. Emissive probe is heated and emit the electrons.
Emitted electrons are accelerated by ion sheath and flow into the plasma. Ion current flow
into the probe increase and the slope in the retarding potential region increase as shown in
Fig. 3-10. Therefore, if the heating and electron emission is enough, the floating potential
Vf can be regard as the plasma potential Vs. Therefore, we measured the floating potentialof the emissive probe. Fig. 3-11 shows the missive probe used in this study. Emissive
probe is made of the 1% thoriated tungsten and the diameter of the filament is 0.125 mm.
The emissive probe is connected to the power supply via transformer and applied 8 kHzsine wave of 10-11 A. This sine wave prevents the potential difference in the filament.

22

Fig. 3-10 I-V characteristic gained by emissive probe

Fig. 3-11 Photo of the emissive probe23

3.4 Retarding Potential Analyzer
Retarding Potential Analyzer (RPA) is used to measure the extracted ions energy. Fig.
3-12 and Fig. 3-13 shows the photo of the RPA and the schematic of retarding potential
analyzer, respectively. The RPA consists of four grids and one collector, as shown in Fig.
3-13. We measured the collector current while scanning the voltage of the Ion Retarding
Grid (IRG) from 0 to 65 V. We apply -40 V to the Electron Retarding Grid (ERG) and
Secondary Electron Suppression Grid (SESG). The Floating Grid (FG), ERG and SESG
are 200 mesh/inch, the IRG is 400 mesh/inch. The FG is the three-ply 200 meshes. This is
to prevent the ionization of the neutral particles and space charge limitation by decreasing
the number of ions and neutral particles flowing into the RPA. The IRG has fine mesh toprevent generation of an area where the voltage is lower than that applied in the mesh. The
following equation was used to calculate Ion Energy Distribution Function (IEDF)
f(𝐸𝑖 ).32)
1
𝑒𝐼c0

f(𝐸𝑖 ) =In this equation,

RG

𝐼c
RG

(3-3)

and 𝐼c represent grid sweep voltage and collector current. 𝐼c0

represent the collector voltage when the sweep voltage is 0 V. Average of the ion energy
Ei,ave is gained from following equation.
+∞

𝐸𝑖,𝑎𝑣 = ∫ 𝐸𝑖 𝑓(𝐸𝑖 ) 𝐸𝑖
0

24

(3-4)

Fig. 3-12 Photo of the retarding potential analyzer

Fig. 3-13 Schematic of retarding potential analyzer

253.5 Ion beam measurement
Ion beam was measured to estimate the thrust performance. Fig. 3-14 shows theschematics of the ion beam measurement. In order to repel the electrons, -30 V is applied
to the ion collector and the ion current is saturated at -30 V. Fig. 3-15 shows the ioncollector used in this study. In previous study, it seemed that extracted ion is scattered bycharge exchange collision.33) The cylindrical collector can collect the scattered ions. This
collector is perforated, and has an aperture ratio of 48% to prevent the confinement of
neutral particles.

Fig.3-14 Schematics of ion beam measurement

Fig. 3-15 Photo of the cylindrical collector

26

3.6 Thrust measurement
3.6.1 Hanging Pendulum Thrust Stand34)
Thrust stand was used to measure the thrust directly. The thrust stand is divided to three
types; hanging pendulum type, inverted pendulum type and torsional pendulum type.
Hanging pendulum type can operate stably and easily. Inverted pendulum type can realize
the high resolution, but the operation is unstably. Torsional pendulum type can realize highresolution, but the size of the stand is large. In this study, hanging pendulum type isadopted for simplicity. Fig. 3-16 shows the thrust stand used in this study. In this thruststand, thruster is installed at the bottom of the pendulum, and the counter weight isinstalled in the top of the pendulum. Displacement is measured by the LED displacement
sensor (Z4D-F04A, made by OMRON) and the displacement is calibrated by using weight.
Eddy current damper is installed in the side of the bottom. Damping system is important tominimize a noise and reduce the time required to reach a stable position of thependulum.35)

Fig. 3-16 Schematics of the thrust stand3.6.2 Aluminum Target Thrust StandThrust was also measured by using aluminum target. Figure 3-17 shows the Schematicsof the thrust measurement by using aluminum target. Aluminum target is hung in front of
the thruster. The distance of the thruster and aluminum target is 3 mm. Aluminum target isdisplace by receiving the ion beam. The displacement is calibrated by using weight. The
27

number of weights pull the target is changed by taking up the thread using vacuum
rotational feedthrough.

Fig.3-17 Schematics of the thrust measurement by using aluminum target

28

Chapter 4
Experimental Result and Discussion
4.1 Ion Energy Measurement
Ion beam was measured by RPA to confirm the acceleration of ion. The length of thedischarge chamber is 33 mm and the number of magnets is 14. The diameter of theorifice is 12 mm. The mass flow rate is 82 g/sec. In this experiment, thruster isconnected to GND. This experiment is conducted in the large vacuum chamber in Kyusyu
University. Fig. 4-1 shows the measured IEDF. The average ion energy was 13.7 eV, 18.4
eV and 19.3 eV for incident power of 10 W, 20 W and 32 W, respectively. The average ion
energy increased with the increasing of the incident power. This is because the incident
power per unit mass increase and electron temperature increase. As a result, the plasma
potential of the discharge chamber increases and ion velocity increase. The acceleration
was confirmed. The IEDF is not symmetric and is biased to the low energy side. This iscaused by charge exchange collisions before the ion reaches the RPA from the thruster.
Therefore, the average ion energy is underestimated.

Fig. 4-1 Ion energy distribution function for several incident power29

4.2 Plasma Diagnostics by Langmuir Probe
The plasma diagnostics is conducted by Langmuir probe. The length of the dischargechamber is 33 mm and the number of magnets is 14. The diameter of the orifice is 12mm. The mass flow rate is 82 g/sec. In this experiment, thruster is connected to GND.
The Langmuir probe is installed in the linear actuator and inserted to the thruster from thedownstream on the center axis of the thruster as shown in Fig. 4-2. Plasma potential is thedischarge chamber at the deepest point in the discharge chamber was 19.0 V, 23.2 V and25.4 V for incident power of 10 W, 20 W and 32 W, respectively. The plasma potential
increase with the increasing of the incident power. This result is consistent with the IEDF
measured by RPA. The potential difference was gentle than expected. The electron densityand electron temperature in the discharge chamber was 7×1017～2×1018 m-3 and 35 eV.
There is some consideration in the Langmuir probe measurement. Firstly, electron density
may be overestimated by the sheath effect. The surface area of the probe become big
virtually and the electron current increase. Secondly, the electron saturation current
decrease, so the potential may be underestimated and the electron temperature may beoverestimated.

Fig. 4-2 Photo of the Langmuir probe measurement

30

(a)

(b)

31

(c)
Fig.4-3 Results of probe measurement, (a) Distribution of the potential, (b) Distribution of
the electron density,(c) Distribution of the electron temperature.

32

4.3 Space Potential Measurement by Emissive Probe
Up to this point, ion acceleration and generation of the potential difference wasconfirmed with the thruster connected to GND. However, in practical operation, thethruster is floated electricity. Therefore, we measured the space potential distribution ismeasured by using emissive probe with the thruster floated electricity. In addition, theeffect of the back ground plasma was investigated, because there is the weak plasma in the
super low altitude. The length of the discharge chamber is 12 mm and the number ofmagnets is 6. The diameter of the orifice is 10 mm. This experiment is conducted in Space
Science Chamber at the Institute of Space and Astronautical Science (ISAS) of Japan
Aerospace Exploration Agency. In this experiment, the aluminum plate was connected tothe thruster to keep the thruster potential equal to the space potential. The size of the plateis 27 cm × 42 cm. The plate is located in the side of the thruster as shown in Fig. 4-4.
Figure 4-5 shows the photo of the emissive probe measurement. Figure 4-6 shows thespace potential distribution measured by emissive probe with the back ground plasma. The
mass flow rate is 82 g/s. The probe current is 10.55 A. As shown in this figure, the spacepotential inside the discharge chamber is 7.65 V, 7.75 V and 8.59 V at incident power of 10
W, 20 W and 30 W, respectively. These potentials are lower than expected, but the true
potentials may be 1.5~2Te higher than the measured potentials by space charge effects.36)
Te is the electron temperature of the plasma around the emissive probe. The space potential
in the discharge chamber increases with the increasing of the incident power. This isbecause the incident power per unit mass increase and the electron temperature increase.
As a result, the space potential increases. This space potential inside the discharge chamber
is about 10 V lower than the potential measured by the Langmuir probe. This difference
seemed to be caused by the effect of the boundary condition of the vacuum chamber. In
fact, the potential distribution measured by Langmuir probe is almost flat outside thedischarge chamber. This means that IEDF measured by the RPA was overestimated about
10 eV which is correspond with the space potential at the point of the RPA installation.
Therefore, the true potential difference generated by this thruster seemed to be about 10 V
as shown in Fig. 4-6.

33

Fig. 4-4 Setup of the tether

Fig. 4-5 photo of the emissive probe measurement

34

Fig. 4-6 Space potential distribution for several incident power

35

4.4 Thrust Performance Estimation
Up to the above section, the ion acceleration mechanism was investigated and the ion
acceleration and the generation of the potential difference is confirmed. Next, we estimatedthe thrust performance. Thrust performance is estimated using ion beam current and ion
velocity as shown in Eq. 2-3. Ion velocity is calculated from the potential difference
measured by emissive prove. Ion beam current was measured by cylindrical collector tocollect the scattered ion by charge exchange collision. The length of the discharge chamberis 12 mm and the number of magnets is 9. The diameter of the orifice is 12 mm. Thruster isconnected to the GND. The mass flow rate is 82 g/sec. This experiment was conducted in
the small vacuum chamber in Kyushu University. Figure 4-7 shows the photo of the ion
beam measurement. Figure 4-8 shows the ion beam current measured by cylindricalcollector and plate collector. The ion beam current is 13.5 mA, 32.3 mA and 51.1 mA at theincident power of 10 W, 20W and 32 W. As shown in this Figure, the ion beam current
measured by cylindrical collector is about twice of the ion beam current measured by theplate collector. This shows that about 50% of the exhausted ions were scattered by charge
exchange collision. Considering the ions escape through the aperture of a perforated
texture of the cylindrical collector, the measured ion beam current should be compensated.
The coefficient of compensation is 1.91, which is the ratio between collector area andaperture area and it is determined by the assumption that the rate which ions escape isequal to the aperture ratio. This is because the Debye length is very small compared withthe diameter of apertures. The Debye length estimated to be 50 m outside of the thrusterfrom past probe measurement and a diameter of apertures on the collector is 3 mm. This
Debye length is estimated from the electron density (1.0×1016～4.0×1016 m-3) in thedownstream obtained by ion saturation current. This estimation is supported by theadditional experiment; the plate collector and perforated plate set in the sides of thethruster and ion beam current was measured. As a result, the ion beam current measured by
the plate collector is approximately the twice the one of the perforated plate. The base of
the cylindrical collector doesn’t have aperture, so the coefficient of compensation doesn’t
correspond with the result. Table 4-1 shows the compensated ion beam current andestimated performance. Estimated thrust is 61 N at the incident power of 10 W. This isunder 1/10 of the objective thrust. However, the performance may be overestimated due tothe effects of the confinement of neutral particles and possible uncertainties in the method36

of the compensation.
For confirmation of the validity of this performance estimation, ion beam current wasestimated by another method. Ion beam current was calculated by following equation.

𝑘 2
)
𝑚𝑖

𝑜 = (

𝐼𝑖𝑛 = 𝑒 𝑖 0 𝑆𝑒𝑥𝑝( 1⁄2)

(4-1)
(4-2)

Eq. (4-1) is the Bohm velocity. From this equation, the ion velocity at the sheath edge is
obtained. Using this ion velocity, ion current in the discharge chamber is calculated from
Eq. (4-2). In this equation, 𝑒,𝑖and 𝑆 are elemental charge, ion density and area. Ion

density is assumed to be equal to electron density. 𝑆 is the orifice diameter for simplicity.
𝑒𝑥𝑝( 1⁄2) is used to calculate the density at the sheath edge from the density indischarge chamber. Table. 4-2 shows the results of the estimation. Comparing thecompensated ion beam current shown in Table. 4-1 and estimated ion beam current shownin Table. 4-3, the difference is only 5% at incident power of 10 W. However, the differenceis 20% and 50% at incident power of 20 W and 32 W. Even after taking into consideration
the uncertainly of the Langmuir probe measurement, this difference is large. Therefore,
direct thrust measurement is needed.

37

Fig. 4-7 Photo of the ion beam measurement

Fig. 4-8 Comparison of the ion beam measurement for cylindrical collector and platecollector

Table. 4-1 Estimated performance
Incident power
Ion beam current
ThrustSpecific impulse
Propellant utilization

10 W
24.2 mA
61 μN
75 sec12%

20 W
66.0 mA
167 μN
206 sec
33%

32 W
105 mA
280 μN345 sec
53%

Table. 4-2 Ion beam current in the discharge chamber
Incident power
Potential current

10 W
20 W
32 W
23 mA 54 mA 69 mA

38

4.5 Thrust Measurement
4.5.1 Measurement by Hanging Pendulum Thrust StandThrust was measured by the pendulum thrust stand. This experiment was conducted inSpace Science Chamber at the Institute of Space and Astronautical Science (ISAS) of
Japan Aerospace Exploration Agency. At first, the displacement was calibrated by using
weight. Before the experiment, the effect of the damper material to the attenuation of theoscillation was investigated. This investigation was conducted in the air. Fig. 4-9 shows theattenuation of the oscillation for several damper materials. As shown in this figure, theattenuation time is minimum when the damper material is copper. This seemed to be due tothe low resistance of the copper. The M 2.6 washer was used as the calibration weight.
However, the displacement for the weight is not constant. Fig.4-10 show the calibrated
relation between the thrust and the voltage change of the displacement sensor. As shown inthis figure, the dispersion was large. The obtained calibration coefficient is 0.864 V/N±
0.321 V/N. The length of the discharge chamber is 12 mm and the number of magnets is
6. The diameter of the orifice is 10 mm. Thruster is floated electricity. The mass flow rateis 82 g/sec. Fig. 4-12 show the measured data when the microwave stopped. The incident
power was 20 W. From this displacement, the thrust is estimated to be 54 N±20 N.
However, this displacement might be not due to the thrust. Fig. 4-13 shows the measured
displacement without plasma. As shown in this figure, the displacement is as much as thedisplacement with plasma vanishment. This seemed to be due to the elasticity changing of
the microwave cable on account of the heating. Therefore, the reliability of the thrustestimated by this experiment is low.

39

Fig. 4-9 The attenuation of the oscillation for several damper materials

Fig. 4-10 Calibration result40

Fig. 4-11 Measured displacement

Fig. 4-12 Measured displacement without plasma

4.5.2 Measurement by Aluminum Target Thrust StandThrust is also measured by using the aluminum target thrust stand. This experiment was
conducted in the small vacuum chamber in Kyushu University. The diameter of the orifice
is 10 mm and the number of magnets is 6. Thruster is floated electricity. The mass flowrate is 82 g/sec. Fig. 4-13 shows the Photo of the thrust measurement by the aluminumtarget. The distance between the thruster and the target is 30 mm. The displacement wascalibrated by the weight. However, displacement was not measured when the plasma
41vanished with the calibration weight connected. This is because the friction force between
the thread and the pulley. Therefore, the thrust was measured with the thread cut. The
displacement was measured when the plasma was ignited and vanished. This means thethrust was generated by the thruster. The displacement increase with the increasing of theincident power and it is consistent with space potential and the IEDF showed in the above
sections. However, the calibration coefficient became invalid because the mobility of thetarget changed by cutting the thread. Therefore, the thrust was estimated by the ideal
relationship between the displacement and thrust. Fig. 4-14 shows the estimated thrust.
Thrust was measured with the thruster connected the GND or floated. Thrust was 3.4 N,
6.7 N and 8.7 N for incident power of 10 W, 20 W and 30 W, respectively. As shown inthis figure, the difference of the thrust for two electrical state of the thruster is very small at
incident power of 10 W and 20 W. Even when the incident power is 30 W, the difference isabout 10%. Therefore, it seemed that the electrical state of the thruster doesn’t affect to thethrust drastically. In this figure, the thrust doesn’t include the contribution of the raw gas.
This estimated thrust is much lower than expected. Therefore, we compared the thrustgenerated by the plasma and raw gas. The Figure. 4-15 shows the thrust measured when
the diameter of the orifice is 12 mm and the number of magnets is 9. Thruster wasconnected to GND. The thrust was 8.6 N, 17 N and 22 N at incident power of 10 W, 20
W and 30 W, respectively. Those thrust don’t include the thrust of the law gas. As shown inthis figure, the thrust generated by the plasma is almost same as the thrust of the law gas atincident power of 10 W. Even at incident power of 30 W, the thrust is only three times of
the thrust of the low gas. These thrust are only 1/10 of the objective thrust. However, thosethrust are 1/61/10 of the estimated thrust shown in section 4.4. There is some reason ofthis disagreement: error of the gas calibration, changing of the potential difference byaluminum target and overestimate of the ion beam current. In this experimental, calibration
was conducted by using gas. However, the reaction of the gas and ion is different. All gas
is reflected by target though some ion attach the target. Therefore, the thrust wasunderestimated by difference of the momentum transfer. The potential difference might beunder 10 V by the effect of the aluminum target. As shown in section 4.2 and 4.3, thepotential distribution was affected by boundary condition, especially the distance of thethruster and the aluminum target is close and it is 30 mm. Even there is no effect for the
potential difference, the potential doesn’t drop enough at the point of the aluminum target
as shown in Fig. 4.6. The ion beam was overestimated as shown in section 4.4. Taking into
42

these error, it seemed that the practical thrust is between the estimated thrust from ion
beam current and the measured thrust by aluminum target.

Fig. 4-13 Photo of the thrust measurement by the aluminum target

Fig. 4-14 Dependence of the thrust in the electrical state of the thruster

43

Fig. 4-15 Comparison of the thrust between the plasma and law gas.

44

Chapter 5
Summary
We investigated the ion acceleration mechanism of the miniature microwave discharge
thruster and estimated the thrust performance. The results is showed as following.

1. Ion acceleration was confirmed by using RPA.
2. The generation of the potential difference was confirmed even in the back groundplasma. The potential difference is about 10 V and it is the half of the expectation.
3. Thrust performance was estimated from the ion beam measured by using cylindricalcollector. The estimated performance was 61 N at incident power of 10 W.
4. It was tried that the thrust measurement by pendulum thrust stand. However, it failed by
the elasticity changing of the microwave cable.
5. The thrust was also measured by the aluminum target and the generation of the thrust
was confirmed. The thrust was 8.6 N.

In this study, the ion acceleration and the generation of the potential difference wasconfirmed. However, the mechanism of the potential difference generation remains unclear.
The investigation about these mechanism is important for improving of the thrust
performance. In future, the simulation code should be established for elucidation of themechanism. After that, the thrust should be measured directly by more reliable method for
accurate evaluation of the thrust performance.

45

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48

Acknowledgements
I would like to express the deepest appreciation to my supervisors; Emeritus Professor
Hideki Nakashima, Associate Professor Naoji Yamamoto, and Assistant Professor Taichi
Morita.
I would like to thank my labmates; Lee Hou-Yi, Naoya Saito, Koichi Ushio, Yu Okuda,
Kohei Takase, Kazuhiko Nakano, Atsushi Yamaguchi, Kensuke Iijima, Tomohiro Ichimaru,
Bunsuke Ueno, Yusuke Egawa, Masafumi Edamoto, Parsa Yousefian, Satoshi Miura,
Yutarou Itadani, Masataka Iwamoto, Yutarou Kawahara, Hirotaka Fuchigami and graduatesin my laboratory.
Some research results were obtained using Space Science Chamber is Space Plasma
Laboratory of ISAS, JAXA.
Finally, I would like to express the deepest gratitude to my family for their kind supportand help.

49