Table of Contents
1. Introduction
1
1.1 Microsatellite ...……………………………………………......................1
1.2 Types of propulsions ………...………………….......................................3
1.3 Required performance…………………………………………………….4
1.4 Purpose …………….…………………………..........................................6
2. Miniature Microwave Discharge Plasma Thruster
7
2.1 Mechanism of ion acceleration………………………………………......7
2.2 Selection of the Propellant……………………………………………….8
2.3 Expectation of the Performance………………………………………...10
2.4 Microwave Discharge……………………………………………….….11

3. Experimental Facility and Setup
12
3.1 Experimental facility ...………………………………………………..12
3.1.1 Vacuum Chamber………..………………………………………….12
3.1.2 Devices being used…………….…………………………………....13
3.2 Miniature Microwave Discharge Thruster...…………………………….16
3.3 Ion beam measurement ………………………………………………….18
3.4 Propellant Supplying system………………………… …………………21
4. Experimental Result and Discussion
23
4.1 Orifice dependency……………………………………………………23
4.2 Magnetic field dependency……………………………………………27
4.3 Shape of Antenna dependency.….……………………………………31
4.4 Thrust Performance Estimation………………………………………35
5. Summary

36

References

37

Acknowledgement

39Chapter 1 Introduction
1.1

Microsatellite

Nowadays, space usage is an important theme for our life. There are many technologies
being used using in spaces such as, communication, GPS, weather observation, military
defense etc. The key technology is satellites. Satellites are launched from the earth byrockets and used in the space. The missions have wide range such as from orbiting aroundthe earth to planetary probe. Among the importance of the satellites, micro satellites arebeing attracted.
Recently, launches of microsatellites has been increasing. The launches of less than 100
1)

kg thrusters are over 100 in 2013 . Around 13 years have passed since first cubesat was
launched, number of recent launches is shown in the fig1-12). The increase in launches of
micro satellites are being caused by low cost and that the constructing time is short. Also
the difficulties in developing and designing are becoming less. By these, not only the spacecompanies and space agencies can develop, but educational industries and venture businesscompanies can also take part of it. The parts of the microsatellites kitting are advancing
recently. That is why number of launches of microsatellites is increasing and by theadvancing in technology even more, it would increase in the future. As challenging
microsatellites have a thruster like the “Hodoyoshi-3&4” and “PRoximate Object Close
flYby with Optical Navigation (PROCYON)” in these days3,4).
The attractive microsatellites also have disadvantages like size restriction and electric
power generating restrictions. Micro satellites don’t have much space inside to take things
with it. So most of the satellites don’t have thrusters with it. Also the solar panels are smallcausing restriction in the amount of electricity able to use. On the other hand, thrusters canmake the satellites to be used in wide range of missions. Thrusters are used to control theattitude and the orbital movement. As missions which thrusters are required, there is Super
low Altitude Satellites “SLATS” project5). It is a project requiring thrusters to cancel out
the drag force, since it is made to move at low altitude. They move around the height ofaround 200 km to detect the surface of the earth with high resolution by a detector.

1

Fig. 1-1 The number of launches of the 50 kg or less class satellites per year2

1.2 Types of propulsions
There are several type of thrusters used in space. The most famous one is the chemicalrocket. There are also electrical rockets which we call electric propulsion. Chemical
propulsion systems and electric propulsion systems have opposite features in terms ofthrust and specific impulse. Chemical rockets get thrust by exhausting high energy andpressure gas by the reaction between fuel and oxygen. It has a feature of high thrust andlow specific impulse. Specific impulse is a word used to express the efficiency of fuel. On
the other hand, electric propulsions have features of low thrust and high specific impulse.
The graph comparing the thrust and specific impulse for some thrusters is shown in fig
1-26). These two features give us habitat segregation in missions. Chemical rockets aremainly used when you want high energy within specific time. Electrical propulsion
systems are used in missions lasting for very long time period requiring low energy.
Electrical propulsion system is being used in more than 200 satellites until now. Electrical
propulsion systems gain thrusts mainly by exhausting high energy ions.

Fig 1-2 Graph of thrust against specific impulse31.3 Required performance
In missions which satellites stay at the height of about 200 km, satellites might movethrough air drag, in this case thrusters are needed to cancel out the resistance and also incases such that the satellites need to change its own position as expressed in section 1.1. To
solve the tasks given, a miniature microwave discharge plasma thruster for micro satellites
is under development.
Now we will assume a mission held at super low altitude to get high resonance. First we
assume the power generated by the solar panels. Generated power is estimated by theassumption that the area of the solar panel is 0.4 m×0.5 m and the conversion efficiency ofthe solar panel is 40%. Secondly, we estimate the drag force we must cancel out. The dragis 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0.5 when the attack angle is 015 deg8). We calculate Cd assumingthat attack angle is 0~15. By adopting the parameter, the drag is estimated to be 0.20.5
mN. This time, we set the objective thrust as 0.5 mN. The power used for thrust systemshould be 2030% for widely application, so the objective power consumption of thethruster system is under 20 W9). Specific impulse is also important parameters for thruster
system. Specific impulse is defined in the following equation and it represents the durationof the unit thrust by using unit propellant mass2
𝐼𝑠𝑝 =

𝐹
𝑣𝑖
≅
𝑚̇𝑔 𝑔

(1-2)

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.
From these parameters, the optimum specific impulse is estimated as 1000 sec. From theabove, the required performance is are 0.5 mN and 1000 sec, respectively, with a total
power consumption of 20 W in the thruster system. In addition, the simplicity of the thrust4system is also important. Simple structure provides high robustness, compactness, lightweight and low cost. These are very important point for microsatellite. Therefore, theobjective performance should be achieved with the simple thrust system.

Fig 1-3 Image of the satellite

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

1 years5

1.4 Purpose
The purpose of our research is to develop a thruster for the use in super low altitude microsatellite. In order to adapt to the micro satellites, the whole system including the tanks forthe propellant needs to be compact in size. In this study we tried to adopt water propellant
to the thruster. Propellant being in liquid state is an advantage since size of the tank is an
important factor. We will talk about choosing propellants and the size of the tank more indetail afterwards. The concept of the developing thruster will be also shown in chapter 2.
Developing thruster uses microwave discharge to generate plasma and it doesn’t have gridto accelerate the ions as ion thrusters. Ion thrusters use grids to accelerate ions by theirpotential difference. Removing the grids form the thruster will give us simple structurewith lower conservation energy.

6Chapter 2
Miniature Microwave Discharge Plasma
Thruster
2.1 Mechanism of ion acceleration
The schematic of ion acceleration is shown in figure 2-1. 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 thisconfiguration is used in some microwave discharge neutralizer and ion thruster10).
Electrons gain energy from the microwaves, they collide with propellant particles, and thepropellant atoms are ionized. Once plasma is generated, ions are accelerated by thepotential difference between the plasma inside the thruster and the space by keeping thepotential of the discharge chamber and space potential same. High energetic electrons canalso go out to downstream with ions against the potential hill. These electrons and ions areneutralized outside the thruster; it does not need a neutralizer. Therefore, this thruster isvery simple and small, with low power requirements.

Fig 2-1 Schematic of the thruster7

2.2 Selection of the Propellant
The most common propellant used for electric propulsion is Xenon. Xenon requires low
ionization energy to be ionized. However in this study, we mainly used water and argon.
Water is a liquid propellant at standard condition, so the volume of the tank can be kept
smaller than gas propellant.
In order to show that the size of tank changes, we calculated the size needed. We assume
the tank is sphere shape.
The equation shown below is used to calculate the thickness of the tank needed.

𝑡=

𝑃𝑟𝛽
2𝜎

(2-1)

As we assumed in section 1-3, we calculate with the amount of propellant lasting for 1 year
kept in the tank made of plastic. The thickness of the tank becomes 24.7 μm. So the weightof the tank becomes 536 mg in case of using polyethylene as a tank. In the case of Xenon,
thickness is 670 μm and it is usually kept in a titanium tank. So the mass of 229.5 g which
is much higher than water tank. Xenon is kept in supercritical state so a pretty thick tank isrequired to keep it safe. In both cases we used 2 as safe value β.
Also we had a choice of other liquid such as alcohol or hydrogen peroxide. We chosewater since it is cheap without causing contaminations as alcohols and less oxidization ashydrogen peroxide.11) Argon is used as propellant on account of high specific impulse andlow cost. Specific impulse is in inverse proportion to the mas of the propellant, so specificimpulse become longer by adopting Argon. The price of Argon is about 1/1000 of the oneof Xenon, because 1% of atmosphere is Argon. The price of water is 1/5 of Argon. Costand size occupation is the important for widely application of thrust system in the future.
The table comparing the propellants is shown in table 2-1. The important part of the tableis mainly the big difference in the cost of each propellant.

8

Table 2-1 The comparison of the propellants
Unit Price (JPY/Kg)

Ionization energy (KJ/mol) Volume

atcondition (dm3)
Xenon

500,000

1170.4

24

Argon

500

1520.6

24

Water

100

1214.60.18

9standard

2.3 Expectation of the Performance12)
It is very important that the thruster is available to achieve the required performance.
Required performances will be calculated by the following equations.
The potential difference between the plasma inside the thruster and the space is estimatedas 20 V, because the sheath potential is 20 V by the previous research done using thedeveloping thruster in Kyushu University. In that case, the velocity of the extracted ion 𝑣𝑖
is calculated by following equation from the ion’s equation of motion. This time weassume, we obtain only water ions H2O+.
𝑣𝑖 = √

2𝜑𝑒
𝑚𝑖

(2-1)

In this equation, 𝜑, 𝑒 and 𝑚𝑖 represents the potential difference, elemental charge
and ion mass respectively. From this equation, the ion velocity is 14.6 km/s.
𝐼𝑠𝑝 =

𝐹
𝑣𝑖
≅
𝑚̇𝑔 𝑔

(2-2)

By the equation above the specific impulse becomes around 1489 sec and estimating thepropellant utilization as 70 %, we obtain 1000 sec of specific impulse. 𝐼𝑏 means the ion
beam current. In this case, we need to obtain 0.5 mN and we calculate the ion beam current
required to get enough thrust.
𝐹 = 𝑚̇𝑣 =

𝑚𝑖
𝐼 𝑣
𝑒 𝑏 𝑖

(2-3)

Therefore, this thruster concept can meet the objective performance. And the ion beam
current needed is 183 mA.

10

2.4 Microwave Discharge
We used the method of microwave discharge to generate plasma. The utility of
microwave discharge is shown in the mission of HAYABUSA13). Microwave discharge
have the merits for thruster shown in the below.

1) Long lifetime and simplification by no need of the electrodes2) Needlessness of the beforehand heating
3) Low power consumption by using existing electron

In this thruster, Electron Cyclotron Resonance (ECR) is used to give energy to theelectrons to accelerate. The accelerated electros hit the neutral particles. By hitting highenergy electrons to neutral particles, plasma will be generated. In this thruster, themagnetic tube is formed between the central yoke and front yoke as. Electrons make around trip between the central yoke and front yoke along the magnetic lines of force andreflected 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 byapproaching the strong electric field area near the antenna many times. Plasma can begenerated 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 themagnetic tube.

11

Chapter 3
Experimental Facility and Setup
3.1Experimental facility
3.1.1Vacuum chamber
There is a vacuum chambers in the laboratory. The small vacuum chamber shown in fig
3-1 and the large vacuum chamber. The inner diameter and the length of the small vacuum
chamber are 0.6 m and 1.0 m, respectively. The pumping system consists of a rotary 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

123.1.2 Devices being used
In this study we use several devices to do experiments. WE 7000 in Fig 3-2 is used tocollect the data being used in the experiment such as Forward and Reflection power
gaining data from the amplifier. And ion beam current obtained form ion collector.
Microwave amplifier in fig 3-5 is used to amplify incident microwave power from themicrowave generator n fig 3-4. This microwave is used generate plasma for the thruster.
Bipolar DC supply is used to apply wide range of voltage to the ion collector.

Fig 3-2 WE7000

Fig 3-3 Bipolar power supply
13

Fig 3-4 Microwave generator

Fig 3-5 Microwave amplifier

14

Table 3-1 Models of the apparatus used in experiments
ObjectCompany

Name of the apparatus

Model

Save the data

Yokogawa

Measuring station

WE7000 series

Oscillate thenF

Bipolar DC Supply

BP 4610

Anritsu

Digital Modulation Signal

MG3660A

power
Generate
microwave
Amplify
microwave

Generator
R&K

High frequency power
amplifier15

CA801M162-4747R

3.2 Miniature Microwave Discharge Thruster
In this study, we use a microwave discharge thruster as shown in figure 3-2. The size is
50 mm×50 mm×27 mm. Fig. 3-3 shows the parts assembly of the thruster with 12 mm anddischarge chamber length. Sm-Co magnets surround the discharge chamber and are locatedbetween yokes made of soft iron. The magnet size is 4 mm×4 mm×12 mm. The innerdiameter and the height of the discharge chamber are 21 mm and 12 mm respectively.
There are a spacer (diameter 21 mm) and an orifice plate(12 mm for the one shown in thephoto) in front of the front yoke. The thickness of the spacer and orifice are 4 mm and 2mm respectively. The antenna is made of molybdenum. The antenna, which is star shaped,
has a thickness of 1 mm and a circular diameter of 9 mm. This star shaped antenna showed
best performance in the previous study.14) We studied about several types of antenna which
are star, circle and cross. For argon star shape showed the best performance.

Fig 3-2 Microwave discharge plasma thrusters of 12 mm diameter thruster

16

Fig 3-3 Parts assembly of the thruster17

3.3 Ion Bean Measurement
Ion beam current was measured to estimate the thrust performance. Fig. 3-4 shows theschematics of the ion beam current measurement. Ion collector is used to collect the ionsand measure ion beam. We only need ion to measure but our thruster also emits electrons
so we have to repel them. In order to repel the electrons, minus voltage is applied to the ioncollector. Fig. 3-5 shows the ion collector used in this study. It has cylindrical shape to beable to measure wide range of plasma from the thruster. At first we tried using platecollector but there are some scattered ions caused by charge exchange15). Fig 3-6 shows thescattered ion beam not being collected by a plate collector.
Cylindrical collector has an aperture ratio of 48% to prevent the confinement of neutral
particles. We applied voltage with 1000 Hz of frequency with the range of 10V ~ -100 V tothe collector. This range is to give saturation to the current. At first when we tried tomeasure ion beam current by simple negative voltage, we couldn’t see saturation in ion
beam current, meaning we can’t measure collect ion beam current value. So we did theexperiment with other ranges but between the range of 10 V ~ -100 V. We need to swing
the voltage to positive voltage to catch the electrons. The photo of the circuit used in thisexperiment is shown in figure 3-7. It uses instrumentation and isolation amplifier. And we
use Bipolar power supply with the circuit explained. Since the ion beam current didn’t
saturate enough so we used fitting with the measured point using the equation 3-1.

𝐼 = 𝐼𝑖 + 𝐼𝑒 𝑒𝑥𝑝 {−(𝜑𝑝 − 𝑉)
}
𝑇𝑒

(3-1)

In this study, optimization of the thruster is a main theme. In order to find the
optimization, we changed the size of orifice, magnetic field distribution and the shape of anantenna using 1.6 GHz of microwave frequency for 10 W ~ 30 W of power. We choose tochange them because it made a better result in previous reaserch16). We use argon andwater propellants with the mass flow rate of 2.8 sccm and 7.52 sccm.

18

Fig 3-4 Schematic of ion beam current measurement

Fig 3-5 Cylindrical shape ion collector

19Fig 3-6 Ion beam measurement using plate collector

Fig 3-7 photo of Circuit used to measure ion beam current

203.4 Propellant supplying system17)
Figure 3-8 and 3-9 shows the propellant supplying system used in this study. The mass
flow rate of the propellant used in the experiment which is water in this case is controlled
by needle valve. First, water is storage in the vaporizing tank. Water will be vaporized at
the tank and be supplied to the temporary vapor storage tank. The pressure of thetemporary vapor storage tank is kept between 1050 and 950. This pressure is controlled bysolenoid valve and needle valve set between two tanks. Finally mass flow rate is controlled
by needle valve and solenoid valve between temporary vapor storage tank and the thruster.
The big tank in fig 3-9 is vaporizing tank and the narrow size one is the temporary vapor
storage tank. The two green valves are solenoid valves which are controlled by labview.
Calibration was done by using pressure drop method. First, temporary vapor storage tank
was with initial pressure P0. Valve will be open and time taken for the change in pressure is
the parameter due to propellant consumption.

Fig 3-8 Diagram of the propellant supplying system

21Fig 3-9 Photo of the supplying system

Fig 3-10 calibration graph for the supplying system

22

Chapter 4
Experimental Result and discussion
4.1 Orifice Dependency
In this study we change the size of the orifice to find the best condition in terms ofigniting and the thrust. The orifice used were 5 mm, 6 mm, 8 mm, 10 mm, 12 mm and 14
mm with 7 number of magnets. When the size of the orifice is small, the density of neutral
particles will increase so the ignition will occur easily but the amount of exhausting ionswill decrease. On the other hand, large size orifices can exhaust more ions but the densityof the neutral particles will decrease. In order to see the orifice dependency we use an ion
collector. Figure 4-1 and Figure 4-2 show the photo of the ignited microwave discharge
thruster using water and argon as a propellant respectively. The frequency of the incident
microwave power is 1.6 GHz and power was from 10 W to 30 W by 2 W. We compare
water and argon propellant with 2.8 sccm and 7.52 sccm for each propellant.

Since therequired performance of the thruster is calculated with 20 W. We compared the ion beam
current when 20 W of incident microwave power is applied. Figure 4-3, figure 4-4 andfigure 4-5 are graphs which show the orifice dependency for the microwave discharge
thruster.
When 2.8 sccm of propellant is supplied to the thruster, the highest ion beam current
value was 4.90 mA and 15.54 mA when water and argon was supplied respectively. For
water the best orifice size was 12 and argon was 14. For the mass flow rate of 7.52 sccm
they both got the highest ion beam current when 14 mm of orifice was used. When 14 mmof orifice was used with 2.8 sccm of mass flow rate, water plasma could not be generated
but when 7.52 sccm was used it did ignite.
The reason why when water propellant’s mass flow rate was 2.8 sccm didn’t show goodresult is because the neutral particle density wasn’t kept in the optimized condition for thegeneration in plasma. In order to generate plasma using microwave discharge, mean freepath is an important factor.

23

Fig 4-1 Photo of water propellant plasma

Fig 4-2 Photo of Argon propellant plasma24Fig 4-3 Graph of orifice dependency (2.8 sccm)

Fig 4-4 Graph of orifice dependency (7.52 sccm)

25

Fig 4-5 Graph of the optimized orifice using water as a propellant

26

4.2 Magnetic field dependency

The magnetic field distribution inside the discharge chamber can be changed by the
number of magnets. Figure 4-6 ~ 4-7 shows the magnetic field distribution for magnets.
For microwave discharge, ECR layer is a very important factor. The surface of an antenna
needs to touch to accelerate the ions efficiently. 9 magnets was the best performance forthe previous research using argon as a propellant. This value was focusing on ion beam
current and not in igniting conditions. In this study we will measure the ion beam current
for the magnets of 6, 7, 8, 9 and 10. The frequency of the incident microwave power is 1.6
GHz and power was from 10 W to 30 W by 2 W as we used in orifice dependency
experiment. Plasma couldn’t be generated with 6 magnets either for argon nor water. The
best ion beam current obtained for water was, when 7 magnets were used and it went down
as we increase the number of magnets. For argon propellant 9 magnets gave us the highest
ion beam current value as shown in figure 4-7. But for 7.52 sccm, the best result isobtained from 9 magnets in both of the propellant.
This suggests that electrons didn’t accelerate enough so energy was low. When we look at
the magnetic field distribution, the ECR layer is inside the antenna so the electrons
couldn’t get enough energy. For more number of magnets, the tendency was different
depending on the mass flow rate for water. For 2.8 sccm water propellant, we can say thatplasma is trapped by magnetic field.
So, if we try using 2.8 sccm, 7 magnets give us the best performance but for 7.52 sccm, 9is the best. Optimization exists for both mass flow rates and type of propellant and they are
different. If we look at figure 4-11, when mass flow rate is 7.52 ion beam current value
increases but not for 2.8 sccm. When mass flow rate is larger, there are cases which plasma
is developed also in front of an antenna is shown in the previous research. So The
importance is plasma being trapped inside the discharge chamber rather than antenna
touching ECR layer

27

Fig 4-6 magnetic field distribution when 6 magnets (left) and 7 magnets (right)

Fig 4-7 magnetic field distribution when 8 magnets (left) and 9 magnets (right)

28

Fig 4-8 magnetic field distribution when 10 magnets (left) and strength (right)

Fig 4-9 Graph of the magnetic field dependency (2.8 sccm)
29Fig 4-10 graph of magnetic field dependency (7.52 sccm)

Fig 4-11 graph when optimized thruster for water propellant (7.52 sccm)

30

4.3 Antenna dependency
Fig 4-9, fig 4-10, fig 4-11 and fig 4-12 show the result of the experiment of the antenna’s
shape dependency. The experiment was done under the condition of 1.6 GHz of incident
microwave power, 12 mm of orifice and 7 magnets. The experiment was done again using
10 W ~ 30 W of incident microwave power. The experiment using circle antenna showed
good result for water propellant but not for argon. For argon it didn’t even ignite plasma.
But when we tried using low frequency we could ignite so ignite condition of 7 magnets
wasn’t suitable for argon. Focusing to the water propellant, at high incident microwave
power the value of ion beam current was higher than other types of antenna. For 20 W ofincident microwave power, there wasn’t much difference in the ion beam current when 2.8
sccm of water was used. When we apply higher incident microwave power, we obtainlarger difference in the antennas than other power. In previous study it is known that low
density propellants, circle shape antenna gave best result but when amount of plasma ishigh, then the ion beam current becomes less18).

Fig 4-12 Star shape antenna31

Fig 4-13 Circle shape antenna

Fig 4-14 Cross shape antenna32

Fig 4-15 Antenna dependency Ar (2.8 sccm)

Fig 4-16 Antenna dependency water (2.8 sccm)

33

Fig 4-17 Antenna dependency Ar (7.52 sccm)

Fig 4-18 Antenna dependency water (7.52 sccm)

34

4.4 Thrust Performance Estimation
We have been changing the size of orifice, number of magnets and shape of the antenna to
optimize liquid propellant microwave discharge plasma thruster. By the above section, we
succeeded to ignite plasma using water propellant to microwave discharge thruster. Up tothis section, we were measuring the ion beam current by using cylindrical collector. Since
the aperture ratio is 48% in case of cylinder collector, we need to recalculate the value ofion beam current.
Table 4-1 shows the thrust and specific impulse calculate using the formulas expressed in
the section above and the recalculated ion beam current. The mass flow rate which isalmost equal to 2.8 sccm of argon is 7.52 sccm of water but the result obtain is not highenough. The objective value is 500 μN.

Table 4.1 Thrust and specific impulse
Ar

Ar

Water

Water

(2.8 sccm)(7.52 sccm)(2.8 sccm)(7.52 sccm)

Thrust190 μN490 μN28.2 μN

135 μN

Specific impulse

242 s

233 s

57 s

145 s

35

SummaryIn this study we investigated about the optimization of the miniature microwave plasma
thruster by changing the size of orifice, number of magnets and shape of an antenna. The
final thrust obtained was 135 μN when using 7.52 sccm and 28.2 μN when 2.8 sccm ofwater propellant. The result of our research is shown below.
1. The best size of orifice was 12. Larger than this orifice we couldn’t ignite plasma for
2.8 sccm water. Because the neutral particle density wasn’t kept in enough condition
for the generation in plasma. In order to generate plasma using microwave discharge,
mean free path is an important factor.
2. The number of magnets condition changes by the mass flow rate of water. When highmass flow rate is being used plasma is generated in front of an antenna so the distancefrom the ECR layer is not an important factor. Plasma will be trapped stronger inside
discharge chamber. So the density of plasma will increase and more ions are expected
to exhaust.
3. The shape of antenna also affects the plasma generation we suppose. The largest thrust
was obtained from circle shape when 30 W but our object is 20 W and at that stage theshape of the didn’t affect much.
4. Thrust obtained was lower than the object value. But thinking of the mass of the tank,
we could try using much higher mass flow rate.

In this study, we succeeded to show the thrust from the thruster. In the future we need tosee the mechanism of the plasma inside the microwave discharge thruster to be able to dochanges more specific. The investigation of the plasma generation mechanism of plasma is
important for the improvement of the thrust performance. In future, numerical simulation
can help us. Also that, the thrust should be measured directly without using ion beam
current or other methods by more reliable method for accurate evaluation of the thrust
performance such as using thrust balance with high resolution.

36

Reference1.) Kaneoka, M., : International and national studies of miniature size artificial satellite,
JSASS-2014- 4035 (in Japanese)2.) Kaneoka, M. : A summary and market of small SAR Satellite , JSASS-2016-4459 (inJapanese)3.) Tsuruda, Y., Nakasuka S., Matsui, M., Mase, I.: Reasonably Reliable System Design
and Operation for Twin Satellite: Hodoyoshi-3 & 4, Proceedings of 58th Space
Sciences and Technology Conference, Nov.2014 (in Japanese).

4.) Koizumi, H., Kasagi, Y., Inagaki, T., Kawahara, H, Yaginuma, K., Asakawa, J.,
Komurasaki, K.: Development of small propulsion system for nano-satellites and its
challenges for the future, Proceedings of 58th Space Sciences and Technology
Conference, Nov. 2014 (in Japanese).

5.) Konoue, K, Imamura, S, Satoh Y, Kawasaki, H, Kohata, H.: Development Status ofSuper Low Altimde Test Satellite “SLATS”, The Institute of Electronics,
Information and Communication Engineers, 111 2011, pp.1-5 (in Japanese).

6.) Hall Thruster research group Tokyo University
http://www.al.t.u-tokyo.ac.jp/hall/jp/projects.html (in Japanese)7.) Hoskins, W. A., Cassady, R. J., Morgan, O., Myers, M. R., Wilson, F., King, Q. D.,
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Acknowledgements
I would like to express the deepest appreciation to my supervisors; Emeritus Professor
Hideki Nakashima, Professor Naoji Yamamoto, and Assistant Professor Taichi Morita.
I would like to thank my labmates; Lee Hou-Yi, Naoya Saito, Koichi Ushio, Kensuke
Iijima, Tomohiro Ichimaru, Yusuke Egawa, Masafumi Edamoto, Parsa Yousefian, Satoshi
Miura, Yutarou Itadani, Masataka Iwamoto, Yutarou Kawahara, Hirotaka Fuchigami,
Tomoki Koyama, Yu Fukumoto, Ippei Takesue, Tomihiko Kojima and graduates in mylaboratory.
Finally, I would like to express the deepest gratitude to my family for their kind supportand help.

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