Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYON

Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30, pp. Pb_13-Pb_22, 2016

1) 2) 2) 2) 2) By Hiroyuki KOIZUMI, Hiroki KAWAHARA, Kazuya YAGINUMA, Jun ASAKAWA, Yuichi NAKAGAWA, 2) 2) 2) 2) 3) Yusuke NAKAMURA, Shunichi KOJIMA, Toshihiro MATSUGUMA, Ryu FUNASE, Junichi NAKATSUKA 2) and Kimiya KOMURASAKI

1)Department of Advanced Energy, The University of Tokyo, Kashiwa, Japan 2)Department of Aeronautics and Astronautics, The University of Tokyo, Tokyo, Japan 3)Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, Sagamihara, Japan

(Received July 31st, 2015)

Initial flight operations of the miniature propulsion system I-COUPS (Ion Thruster and COld-gas Thruster Unified Propulsion System) are presented with problems found in space and its countermeasures for them. The I-COUPS was developed by the University of Tokyo and installed on a 70 kg space probe, PROCYON as main propulsion system to verify propulsive capability of the first micropropulsion in deep space. The PROCYON was successfully launched on December 3rd, 2014 and inserted into an orbit around the Sun. The PROYON project team started flight operation on the interplanetary orbit. Up to today, the cold-gas thrusters have successfully conducted unloading maneuvers since the launch. The ion thruster overcame several problems and achieved 223 hours operation with the averaged thrust of 346 μN. The I-COUPS will become the first electric propulsion and reaction control system operated on a small space probe (<100 kg) on an interplanetary orbit.

Key Words: Microthruster, Ion Thruster, Small Satellite, Deep Space

1. Introduction PROCYON were verification and demonstration of several BUS technologies of a 50 kg-class space probe for deep space Development and utilization of small satellites are exploration. Moreover, flyby observation of near-Earth continuing to grow up in the world. However, most of space asteroids was defined as advanced mission in addition to the missions using small satellites are currently limited to remote nominal missions (The PROCYON is an acronym of sensing in LEO. Exploring further possibilities by small PRoximate Object Close flYby with Optical Navigation). The satellites needs wide varieties of technology growth. Among detail of the spacecraft will be described in the reference 3). those, capability of a propulsion system has greatly limited the Developments of I-COUPS and PROCYON were started in activity of the satellites. Most of current small satellites have September 2013 and they were successfully launched on not been equipped with a propulsion system and they could December 3rd, 2014 as a small secondary payload by an not control their orbits. Developing a small propulsion system H-IIA launch vehicle along with the main payload suitable for small satellites can drastically enlarge the HAYABUSA-2 4) and inserted into the intended orbit around available missions 1). the sun. The PROCYON would the first small space probe to The University of Tokyo proposed and developed a new explore deep space independently in the class of less than 100 micropropulsion system named as I-COUPS (Ion Thruster and kg. After initial checkout, usage of the micropropulsion COld-gas Thruster Unified Propulsion System) for the first system, I-COUPS, started. On December 6th, reaction control step to pioneer missions of small spacecraft. The propulsion system using cold-gas thrusters was successfully operated and system unifies an ion thruster 2) and cold-gas thrusters by on December 28th, ion thruster accomplished the first sharing the same gas system. Combination of electric and acceleration and thrust generation was confirmed by Doppler chemical propulsion provides spacecraft with both high ΔV shift of the communication wave. This would be the first maneuver and high thrust, short time maneuver. Simple verification of micropropulsion system on the small spacecraft structure of cold-gas thrusters and sharing the same gas of less than 100 kg in deep space. In this paper, in-flight system with ion thruster realize very light weight and compact operation of the micropropulsion system I-COUPS from 2014 reaction control system suitable for small spacecraft. December to 2015 April will be described. The I-COUPS was installed on the small-space-probe PROCYON developed by the University of Tokyo and Japan Aerospace Exploration Agency. Nominal missions of the

Pb_13 Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30 (2016)

2. I-COUPS Table 1. Specifications of the I-COUPS

I-COUPS is the micropropulsion system installed on the Total (including propellant) 9.96 kg small space probe, PROCYON. It has three propulsive Mass Propellant (xenon) 2.57 kg functions: unloading wheel momentum as RCS, providing ΔV required for Earth gravity assist, and short-time, high thrust No thruster operation*1 7.3 W maneuver for TCM (RCS: reaction control system; TCM: Power Cold-gas thrusters operation*1 11.5 W trajectory correction maneuver). Its nominal missions were set consumption (two thrusters operation) Ion thruster operation*2 to 1) providing reaction control system using the cold-gas 40.0 W (beam current = 5.69 mA) thrusters, 2) verifying the operation of the miniature ion *1: at spacecraft bus voltage: 29.6 V thruster in deep space, and 3) achieving meaningful velocity *2: at spacecraft bus voltage: 30.4 V increment by the ion thruster. Additionally, its advanced missions were set to 1) achieving the orbit change required for which was developed for the 60 kg, LEO-satellite 8-10) earth gravity assist and 2) accomplishing the trajectory HODOYOSHI-4 . Difference from the MIPS is the correction maneuver by the cold-gas thrusters. addition of CTU whose gas is extracted from the xenon GMU. The total mass and power consumption of the system are The details of the gas system are described in the next section. summarized in Table 1. The mass was measured before the The controller ICU implemented an additional board to handle launch and the power consumption were values measured in newly added 12 valves, where eight were thruster valves and the space operation (those are not averaged value, but typical four made its redundant system. The ITU and PPU have no ones). The total mass was 9.96 kg including 2.57 kg xenon. difference from the MIPS installed on HODOYOSHI-4. The power consumption was 7.3 W at its standby condition 2.1. GMU: Gas Management Unit where telemetries are recorded and no thruster is operated. The GMU has a two-staged pressure regulation to provide Two cold-gas thrusters operation increased the power to 11.5 two different mass flow rates for an ion thruster and cold-gas W and an ion thruster operation increased it to 40.0 W. The thrusters. The system diagram of the GMU is shown in the Fig. power consumption depends on the spacecraft bus voltage and 1. Pressurized xenon is stored in a CFRP/GFRP tank, HPGT 3 ion thruster current. The telemetries of bus voltage and current (high pressure gas tank). The capacity of the tank is 2000 cm , have uncertainties of 0.196 V and 0.049 A respectively, its dry mass is 1.00 kg, and the service pressure is 19.6 MPa. associated with the resolution of A/D-conversion. Hence the Charging pressure of the I-COUPS was set to 7.75 MPa at power consumption calculated from those telemetries has 30 °C. The high pressure gas is regulated down to 0.27 MPa uncertainty of maximum 1.75 W. by a pre-fixed-mechanical pressure-regulator. The I-COUPS consists of five units: an ion-thruster unit The gas line is divided to the cold gas thruster unit and ion (ITU), a cold gas thruster unit (CTU), a power-processing unit thruster unit at the downstream of the mechanical regulator. (PPU), a gas-management unit (GMU), and a I-COUPS The cold-gas thrusters directly use this pressure 0.27 MPa for control unit (ICU). The components and structure of I-COUPS their operation. In the ion-thruster gas line, 0.27 MPa gas is are based on the miniature ion propulsion system: MIPS 5-7), further regulated down to about 30 kPa using a low-pressure

High pressure gas tank

7 – 8 MPa

Pressure sensor Drain valve

Mechanical pressure regulator

0.30 MPa High-pressure cutoff valve (latching valve) Pressure sensor Regulation valve (solenoid valve)

Drain valve

Feedback control Cold-gas cutoff valves 30 kPa Low pressure gas tank (solenoid valves) Pressure Cold-gas sensor Flow restrictor thruster valves (solenoid valves) Ion thruster valve (solenoid valve) 〜1 kPa Cold-gas thrusters

Ion thruster CT1 CT2 CT6 CT7 CT3 CT4 CT5 CT8 Fig. 1. Diagram of the gas management system.

Pb_14 H. KOIZUMI et al.: Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYON gas tank (LPGT), a solenoid valve, and pressure sensor. Table 2. Summary of the performance measurement of a BBM So-called “bang-bang control” is used for this second-stage cold-gas thruster. pressure regulation. The low pressure tank has the capacity of Upstream pressure: 0.212 MPa 80 cm3 to extend the control period. The exit of the low Upstream gas temperature:݌ୡ୮ୱ 294.4 K pressure tank is connected to the ion thruster through a flow ୡ୮ୱ restrictor. The pressure is controlled depending on the Thrust: ܶ 21.3 mN required mass flow rate. Mass flowܶ஼ rate 88.4 mg/s The high-pressure-gas system was assembled by COTS Specific impulse 24.5 s (commercial-off-the-shelf products). The high-pressure 5.85 nN Pa-1K-1/2 CFRP/GFRP tank was provided by Teijin Engineering ܣൌ ܶ஼Ȁሺ݌ୡ୮ୱඥܶୡ୲ୱሻ Limited (ALT764J). The solenoid valves were by Lee Z axis. In this case, translational force is not applied to the Products Ltd. The mechanical regulator and tube fittings were spacecraft. As for TCM, the thrusters always need a paired from Swagelok Company. Only the space qualified product operation. A pair of 1 and 2, 3 and 4, 5 and 6, or 7 and 8 was the HC (high pressure cutoff valve), which was added to provides minus Z, plus Z, minus X, or plus X direction increase the reliability of the gas system. respectively. 2.2. ICU: I-COUPS Control Unit The cold-gas thrusters are assembled from thruster heads, The I-COUPS control unit is an electrical interface with the tube connectors, tubes, and valves. The thruster head is made spacecraft on-board computer (OBC). All of the analog of aluminum and has a conical nozzle. The operation of the telemetries and statuses of the propulsion components are thrusters is controlled by open/close of cold-gas thruster acquired by the ICU and sent to the OBC after the electrical valves. The eight thruster valves are divided into two groups. isolations and analog-to-digital conversions. The ICU receives Parallel-connected, two valves are connected at the further commands from the OBC, to open/close all of the valves and upstream of each group as shown in Fig. 1. Those upstream to switch on/off of the HVPS and MPS. This control includes valves, called as cut-off valves, are used as redundant systems the open/close of the solenoid valve for the second-stage for opened/closed failures of the cold-gas valves. All of the pressure regulation at the low pressure tank. valves, twelve in total, are installed on a single plate. The Powers to drive valves, switches, and a heater were directly lengths of stainless tubes ranged from 200 – 1100 mm. provided from ICU, not through PPU. The valves used at CTU The thrust of the cold-gas thrusters: is described by the and ITU need 0.5 W power for its single operation, and the equation ܶେ operation results in 0.7 – 0.8 W power consumption at the ( 1 )

ICU. The heater warming up the CTU valves results in 3.0 W where and େ are theୡ୮ୱ gas pressureୡ୲ୱ and temperature at power consumption at the ICU. ܶ ൌ ܣ݌ ඥܶ the upstreamୡ୮ୱ of theୡ୲ୱ cutoff valves and is an coefficient 2.3. CTU: Cold-gas Thruster Unit including݌ the effectܶ of nozzle expansion and pressure drops by The cold-gas thruster unit, CTU, has 8 cold-gas thrusters valves and tubes. ܣ that are allocated at the center of each side of the spacecraft as The coefficient A was experimentally measured using a shown in Fig. 2. Nozzle directions of all the thrusters were thrust balance developed for an electric propulsion system. canted from the panel surface to provide both of rotational The detail of the thrust stand was described at the reference 11). force for RCS and translational force for TCM. As for RCS, This experiment was performed using bread-board-model the cold-gas thruster 1 and 2 provides the torque around plus thrusters and gas system. A light weight pressure tank was X and minus X respectively and thruster 3 and 4 provides the installed at the upstream of the pressure regulator and located torque around minus Y and plus Y respectively. Basically outside the vacuum chamber. The mass of the tank before and those thrusters are singly used, accompanying translation after the 30 s-operation of the BBM thruster was measured by force. On the contrary, the thrusters 5 to 8 base a coupled a precise balance. The mass difference divided by the operation. A couple of the thrusters 5 and 7 provides the operation time was used as averaged mass flow rate. The torque around plus Z axis and the thrusters 6 and 8 for minus measurements were conducted with different tube length from

CT-2

CT-6 CT-5 CT-2 CT-1 CT-1,2 PZ CT-1 CT-6 MY PX X Z Z MX MX PZ MY Y Y X CT-5 MY Ion thruster Z X CT-7 CT-8 Z CT-8 MZ CT-3,4 CT-3 CT-4 Y X Y CT-7 CT-3 CT-4 (b-1) (b-2) (b-3) (a) (b) Fig. 2. Configuration of the thrusters (a) and details of cold-gas thruster arrangement (b); (b-1)operations of CT-6 and CT-8 produces torque around plus-Z axis, (b-2)operation of CT-2 produces torque around minus-X axis with unwanted translational motion, and (b-3)operation of CT-4 produces torque around minus-Y axis with translational motion; abbreviations: CT: Cold-gas Thruster, PX: Plus-X panel, MX: Minus-X panel, PY: Plus-Y panel, MY: Minus-Y panel, PZ: Plus-Z panel, MZ: Minus-Z panel.

Pb_15 Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30 (2016) N

I number of ion propulsion systems in space are operated = NPS: Neutralizer S

I Neutralizer power supply without this power supply and electrons are emitted by a slight that

potential drop of the satellite against the space potential. so NPS current: IN Electrons However, this potential drop cannot be measured unless APS: Accelerator power supply special equipment measuring spacecraft potential is installed on the spacecraft. The reasons we employ the NPS is to avoid SPS: the potential drop and also to monitor the status of the

control Voltage Screen power supply Ion beam neutralizer (health-check). 2.5. PPU: Power Processing Unit SPS current: IS A power processing unit consists of a high voltage power Ion beam source supply (HVPS) accelerating ions and generating thrust and a Satellite ground microwave power source (MPS) generating plasma. The high voltage power supplies: HVPS includes the screen Fig. 3. Electrical connections of ITU and high-voltage power supplies. power supply (SPS), the accelerator power supply (APS), and the neutralizer power supply (NPS). All of the power supplies 270 to 1000 mm and with different number of the are operated by unstable 28 V (typically 24-32 V). Output tube-bending from 2 to 6. The measured thrust and mass flow voltage of the SPS is 1.5 kV and its maximum current is at 7.0 rate did not depend on those parameters. The result was mA. Output voltage of the APS is -200 V and its typical summarized in Table 1 by averaging the data of the all current is less than 50 μA. Output of the NPS can range from conditions. In this experiment, the upstream pressure was set 0 to -100 V depending on the required current and the plasma to 0.21 – 0.23 MPa. status. The typical voltage is from -10 to -40 V and this 2.4. ITU: Ion Thruster Unit voltage is referred as to contact voltage. The total energy An ion thruster unit, ITU, consists of an ion beam source, a conversion efficiency was about 50% depending on the output neutralizer, a thruster valve, a gas distributer, a gas isolator, current, where the efficiency was defined as summation of the and DC-blocks. The ion beam source and the neutralizer have output powers of three supplies divided by the input unstable ECR-plasma sources driven by microwave injection 28 V power. respectively and both sources have almost identical design. The microwave power supply (MPS) has two outputs of the Both plasmas need the same amount of microwave power, microwave power of 1.4 W by each output. The 1.4 W about 1.0 W each other. The detail of the plasma sources was microwave is transmitted to the ion source or neutralizer 12-14) described in the reference . The ion beam source installs a respectively. The semi-rigid cables and DC-blocks installed two-grids system to accelerate and exhaust ions. The between the transmission lines have the total 7% transmission neutralizer installs a four-holed orifice for electron emission. loss, and 1.3 W microwaves are transmitted to the ion source Contrasting to the microwave power, lower gas conductance and neutralize respectively. The MPS includes an oscillator of of the neutralizer orifice causes the less than half gas flow rate 4.25 GHz microwave, GaN amplifiers, dividers, and isolators. of the neutralizer compared with the ion beam source. Microwave isolators were implemented after the divider on Applying three types of DC voltages to the ITU generates the both lines to prevent unexpected coupling by reflected thrust by ion beam exhaust and its neutralization by electrons. waves. The three types of voltages are provided by a screen power The final-stage of the amplifier is a GaN-FET to achieve supply (SPS), an accelerator power supply (APS), and high energy conversion efficiency. The conversion efficiency neutralizer power supply (NPS). Electrical connections of of the MPS-BBM showed the maximum efficiency of 48% those supplies are shown in the Fig. 3. Thrust generated by the from three DC powers to the output microwave power (single ITU: was calculated using measured beam current by output). Three DC voltages are 24 V to the FET drain, -5 V to � , ( 2 ) the FET gate, and 5.0 V for the oscillator. Flight operation of � where is the thrust coefficient, which was unknown before the MPS needs to consider more two efficiencies. DC/DC �� ⁄� the space operation and had been assumed to 0.90 on the conversion from unstable 28 V, which is supplied from the �� ground test, is the ion beam current, difference between satellite bus, to the regulated those three has typical efficiency the SPS and APS currents, is the mass of ion particle, around 70%. Splitting and isolating the output microwave �� is the beam voltage, equal to the SPS voltage, and is the power has efficiency around 80%. As a result, the MPS-FM � �� elementary charge. The NPS applies negative voltage to the had about 20% of the total efficiency. � neutralizer against the spacecraft common potential to enhance the electron emission. The voltage of the NPS is 3. Initial Checkout of Propulsion System controlled such that the NPS current equals to the SPS current. The electrons are emitted from the neutralizer by the potential 3.1. Initial checkout of GMU difference between the neutralizer and space. This means that The gas management unit, GMU, has three major functions: potential drop of the whole spacecraft against the space causes safety storage of high pressure gas, first-stage pressure electron emissions. Hence, the NPS is not necessarily needed regulation for CTU, and second-stage pressure regulation for for the ion propulsion system if potential drop of the satellite ITU. Burst or leakage of the high pressure components was against the space is accepted to emit the electrons. Actually, a the most fatal failure for the spacecraft, and most of efforts

Pb_16 H. KOIZUMI et al.: Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYON

Table 3. Comparison of the last ground data and the initial flight data Pressure regulation using RV without any CTU operation was as to the gas system. confirmed in good health and ITU operation had been On the ground On the orbit conducted in this mode for a while. th th (Nov. 5 , 2014) (Dec. 5 , 2014) Another problem, leakage of the ion thruster valve (IV), was Pressure at the HPGT 6954 kPa 5402 kPa found during this activity. There was no sign of leakage on the Pressure at the CTU 107.6 kPa 113.8 kPa ground test and at the initial checkout of the GMU, but the IV Pressure at the LPGT 155.5 kPa 161.3 kPa began to leak after second close in the space. One of the Temperature at the HPGT 25.49 ℃ 10.16 ℃ possible causes of the valve leakage is particle contamination of the sealing face. Although we conducted trials to flush out Temperature at the LPGT 26.27 ℃ 15.36 ℃ the particles by xenon flow, the leakage rate was not recovered. Fortunately leakage of ion thruster valve was not developing GMU were spent to increase that reliability. The fatal for the propulsion system, because that main effect is to first operation on ICU and checkout of the gas system was waste gas stored in the low pressure tank after finishing the performed on Dec 5th, 2014, two days after the Dec. 3rd ion thruster operation. The gas wasting was typically 30 mg launch. for an ion thruster operation. All of the four pressure sensors showed nominal values 3.2. Initial checkout of CTU measured on the ground before launch. The acquired initial Initial checkout of CTU was conducted on December 6th flight data are shown in Table 3 in comparison with the last just after confirmed the functions of high pressure components ground data. Only the high pressure sensor showed quite and first-stage pressure regulation. The main function of the different value from the ground, and this was caused by the CTU is unloading the reaction wheels and any incident of this temperature difference. Highly compressed xenon shows unit immediately affects the spacecraft function. The healthy strong nonlinearity, that is, the behavior is far from the ideal operation of the CTU is inevitable to achieve the nominal gas. Based on the database of the real gas, the pressure mission, verifying BUS technology, and has higher priority decrease can be explained by the decrease of the temperature. than ITU which is necessary for advanced mission, asteroids The next activity was opening the high pressure cut-off flyby observation. valve (HC) which was closed during the launch. Before the All of the possible combinations of CC and CV were tested HC opened at the first time, several downstream valves, cutoff and cold gas was ejected during a few seconds. Operation of valves (CC) and cold-gas-thruster valves (CV) were opened in the valves was confirmed from the pressure sensor located at preparation for unexpected high pressure between the between the mechanical regulator and cold-gas cutoff valves. mechanical regulator (MR) and the HC. The HC was opened, A mechanical regulator makes secondary pressure depending CC and CV valves were closed after 3.0 s, and HC was closed. on the gas flow rate and that pressure, CPS, was used to check After this operation CPS (CTU pressure sensor) showed the nominal value and reaction wheels changed the spin rate corresponding to the ejected gas. This operation showed ITU plate -T/℃ HVPS-T/℃ mechanical regulator, first-stage pressure regulation, was in Screeen current/mA Microwave power/W good health. 50.00 The third function, the second-stage pressure regulation, was tested on December 15th for the plasma ignition and a malfunction was found at the opening control of the valve. 40.00 The test started from opening ion thruster valve for 1-2 hours to release the gas in the low-pressure tank. That tank had been filled with 0.15 MPa xenon since the final ground test to avoid 30.00 contamination from atmosphere. When the pressure at the low-pressure tank achieved down to 30 kPa, the pressure regulation started. However, the regulation valve (RV) opened 20.00 in much longer period than ordered by the controller and the pressure excessively increased. Ordered opening period of the RV was 48 ms, and the actual opening period was estimated 10.00 as 1 – 2 s. from that pressure increase. After several trials and analysis, the project team founds this unexpected opening was caused by software bug of the controller associated with the 0.00 CTU (cold-gas thruster unit) operation. Once one of CTU 10:00 11:00 12:00 13:00 14:00 15:00 valves was operated, this anomaly was active and clearing this Time (hh:mm) error needed the restart of the ICU. On the ground tests, operation of CTU and ITU was separately carried out on th different days due to the difference of required background Fig. 4. Time history of temperature on the ITU plate (Jan. 4 , pressures for each. The initial checkout of the GMU on the 2015). The temperature was affected by the microwave power and not by high voltage powers. orbit became the first successive operation of CTU and ITU.

Pb_17 Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30 (2016) the open/close of the valves. Additionally, gas-ejection applies hours after the plasma ignition at 11:00. First, short time torque to the spacecraft and causes the change of the spin rate acceleration of 10 s was examined and we confirmed the of the wheels. The measured pressure and torques were close healthy ion acceleration. Then we extended the operation to the expected values. There was no trouble for this checkout. period step by step from 10 s, 300 s, 300 s and finally 1800 s. 3.3. Initial checkout of ITU and PPU Although breakdowns (over-currents) were observed during Checkout of PPU and bake out of the ion thrusters were the operation, there was no effect to every component of the conducted in advance of the ion thruster acceleration. Initial spacecraft. After 1800 s operation, we confirmed thrust checkouts of accelerator power supply, neutralizer power generation from the two-way Doppler shift data of the supply, and microwave power supply were conducted on communication waves and estimated thrust was 366 μN. This December 16th. Each power supply was turned on without gas first ion beam acceleration was depicted in Fig. 5. and plasma and we confirmed expected outputs. Only the Two unknown phenomena were found which were different screen power supply to generate 1500 V was tested on from the ground tests. Firstly, the measured ion beam current December 25th after the bake out finished. There was no was higher by 7 – 13 %. Secondly, the neutralizer voltage was trouble for the PPU checkout. higher and it had increasing trend. The neutralizer voltage was The main purpose of the baking was to decrease out-gas typically 25 V on the ground tests, and the initial flight from discharge chamber and grid system. The ion thruster of operation on December 28th it increased from 36 V to 38 V in I-COUPS does not have a heater heating the thruster and 1800 s. plasma ignition was used for this purpose. Plasmas of ion The first phenomenon settled in successive operations. In beam source and neutralizer were first ignited on December the operations of December 29th and 30th, the beam current 22nd and bake out of 5.5 hours in total was conducted. The was decreased to the ground test values. This decrease was not plasma was produced by 1.4 W microwave power and a total continuous but stepwise change. Although the reason of this 2.8 W would be dissipated on the ion thruster plate. The actual change was not clarified, the ion beam current kept at this temperature increase was around 10 – 15 ℃ at the sensor on level after this change. The second phenomenon, high the ion thruster plate. neutralize voltage, was not solved in the succeeding two The initial firing of ion thruster was successfully conducted weeks operations. Characteristics of the neutralizer operation on December 28th. Plasmas were ignited two hours before the were summarized as follows: first ion acceleration, and ion beam source and neutralizer ・ Neutralizer voltage when the acceleration started was reached an equilibrium temperature when the acceleration different operation by operation. started. Fig. 4 shows the example of the time history of the ・ Increasing rate of the voltage after the acceleration started IPU plate, where the temperature became almost flat two was different operation by operation.

Two-way Doppler shift/mms-1 Jan 06 (2015) 14:50 Jan 05 (2015) 13:50

Fitted line Dec 30 (2014) 14:50 Dec 28 (2014) 14:22

Ion beam current/mA Jan 06 (2015) 12:20 Jan 07 (2015) 10:40

25.00 40.0

20.00 38.0

15.00 36.0

10.00 34.0 Neutralizervoltage/V

5.00 32.0

0.00 30.0 14:00 14:30 15:00 15:30 16:00 0 600 1200 1800 2400 3000

Time (hh:mm) Time/s

Fig. 5. The first ion beam acceleration of I-COUPS accompanying Fig. 6. Time histories of neutralizer voltages operated on different the Doppler shift (Dec. 28th, 2014). The thrust was estimated as 366 dates, where the acceleration starts at t = 0. Each operation shows μN from the slope of the fitted line. gradual increase and different start point.

Pb_18 H. KOIZUMI et al.: Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYON

・ The increased neutralizer voltage was recovered after rest. 95%) even with the valve leakage instead of additional gas Typical behaviors of the neutralizer were shown in Fig. 6. loss. The project team selected reliable ignition method rather High neutralizer voltage could be explained by microwave than saving some propellant. Additionally the number of transmission loss and gas leakage in neutralizer microwave open/close cycle of the ion thruster valve was reduced to the and gas lines. However, it was difficult to explain the slow minimum through all of the operation to avoid further leakage. increase (a few volts per hour), because the acceleration 4.2. Error control of the pressure regulation valve started after the ion beam source reached to the thermal The error control at the second-stage pressure regulation equilibrium in plasma heating. Large fraction of the could be solved by not using the cold-gas thrusters. However, DC-power consumption was used for ion beam acceleration or this method did not allow the continuous operation of ion dissipated at power supplies as conversion loss. thrusters longer than 40 hours. The I-COUPS does not have a Another malfunction, unexpected freeze of the I-COUPS gimbal system controlling thrust vector of the ion thruster, and controller unit, was found in two weeks of trouble shooting as the ion thruster operation was always associated with the to the neutralizer voltage. The frozen controller did not disturbance torque. Safety management of the total angular receive any command and send any telemetry data and its momentum of the spacecraft needed unloading by the cold-gas occurrence rate was once or twice 10 hours operation. After thrusters per 30 – 40 hours under the ion thruster operation. several weeks of trials and analysis, we found the occasional Another solution of this problem was sending the pressure freeze was also caused by software bug of the controller. That regulation command at intervals. The pressure regulation bug exists in the command receiving processes and the starts after the controller received a dedicated command (RV controller stopped when total bytes count of commands command). The controller checks the pressure at the matched to the factor of 256 bytes. ICU commands which are low-pressure tank and if the pressure was less than preset sent from spacecraft OBC to ICU ranged from 25 to 32 bytes. value, it opened the RV in a designated period. The judgement For example, when a set of commands: 25, 25, 25, 25, 25, 25, to open or not to open the valve repeats every 5 s. It was found 25, 25, 28, and 28 bytes was sequentially sent to the ICU, the that the software bug became active at only second and controller failed to process the last command and could not subsequent judgements of the pressure. Hence repetition of process any command after that. sending and cancelling the RV commands enabled to use only the first judgement, preventing the occurrence of that bug. The 4. Countermeasures against the Problems project team assembled this gimmick using the OBC command function. In the steady operation of the ion thruster, Following problems were found after the initial checkout of the RV command followed by its cancelling command was the propulsion system. issued every 200 s. The period 200 s was needed to send other #01: Leakage at the ion thruster valve commands between repeated RV command issues. This #02: Error control of the pressure regulation valve command handling enabled simultaneous operations of ITU #03: Occasional freeze of I-COUPS controller and CTU. #04: High and gradual increasing neutralizer voltage 4.3. Freeze of the I-COUPS controller The problem #01 was not critical for the system although The OBC of PROCYON stops the ion thruster operation wasting some propellant, #02 and #03 interrupted a long-time, when it finds freeze of the I-COUPS controller to avoid continuous operation of the ion thruster, and #04 had a thruster operation without any monitoring. Then occasional potential decreasing the neutralizer lifetime and might disable and unexpected freeze prevented continuous operation of the a few thousands operation. The PROCYON project team spent ion thruster. That controller freeze was solved by managing two months (Jan – Feb in 2015) to settle those problems to the total bytes of all the sent commands. However, the achieve the goal defined as the advanced mission which above-mentioned malfunction with RV command and its requires 2000 – 3000 hours of ion thruster operation. countermeasure complicated the situation, which requires 4.1. Leakage in the ion thruster valve sequential issues of RV commands. Particularly, the size of An important role of that valve is to provide instantaneously RV command was 31 bytes, prime number. The total byte high flow rate to surely ignite plasmas. The ion thruster counts reached to the factor of 256 bytes before sequential 256 plasmas are ignited by the microwave input and gas injection RV-commands were sent even in any condition. and high microwave power and high gas-flow-rate increase A set of commands whose total size matches to 256 bytes that ignition probability. The nominal ignition process opened was used to solve this problem. The project team sent one the ion thruster valve after the microwave injection. The gas RV-command associated with eight “meaningless” commands stored between the valve and flow restrictor causes every 2000 s. The eight commands were selected such that the instantaneously high gas flow at the moment of valve opening. total size of the nine commands matches to 256 bytes. If we The leakage from the closed valve lowers this effect and has continue to send this command set, the remainder (total bytes possibility to decrease that probability. divided by 256) circulates a certain pattern. Then if initial After the leakage was found, the gas flow rate was increased condition was appropriately selected, the set of commands up to 160 μg/s only when plasmas were ignited from the was sent any number of times without freezing the controller. nominal rate of 30 μg/s (actual change was the regulation 4.4. Neutralizer voltage pressure at the low pressure tank from 30 kPa to 75 kPa). This The nominal value of the neutralizer voltage of the new ignition process enabled high ignition probability (> I-COUPS ion thruster was less than 30 V based on the

Pb_19 Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30 (2016) ground-test results, and neutralizer voltage higher than 30 V February show clearly different dependency, and it was was regarded as off-nominal operation. Additionally changing from February to March. The project team increased according to the past researches on the microwave neutralizer, the gas flow rate up to the twice as high as the nominal flow the project team had limited the operation to less than 40 V to rate. As a result, the neutralizer voltage was recovered down ensure the life time of the neutralizer. Hence the high to 40 V and in addition the increasing trend was stopped. This neutralizer voltage and its gradual increase with the extremely high mass flow rate solved the abnormal behavior acceleration in the space was clearly off-nominal state and of the neutralizer. However, the mass flow rate to the ion prevented the project team from continuous operation of the source was also increased in conjunction with the neutralizer ion thruster over one hour. flow rate. To clarify the causes of the off-nominal behavior and to keep the voltage at healthy value, the project team examined 5. Steady Operation of the Propulsion System several trials: high and low gas flow rate, hot or cold temperatures, interval operations, and beam acceleration 5.1. Operations of the CTU without neutralizer power supply operation (electrons are Unloading operations by the cold-gas thrusters were emitted through a bypass diode). High neutralizer voltage and successful from the first operation on December 6th, 2014 to in its increasing trend had changed operation by operation and May 2015. One of the twelve valves (CV-5) had a tendency to any methods did not solve the problem. On the contrary the occasionally fail the open in a low-temperature environment. status seemed to become worse gradually during those trials. That was found on the ground tests and the CTU was Eventually, on February 15th, the neutralizer trend equipped with a heater dedicated to warm up the valves. In the drastically changed and the voltage rapidly increased up to flight operation, that valve showed the same opening failures 100 V as shown in Fig. 7. However, voltage dependence on and the heater was always turned on prior to opening that the gas flow rate in this rapid increase of the voltage showed a valve. There was no opening failure with usage of that heater. new trend different from the previous one. Before that Accumulation of the wheel momentum was mainly caused operation, the neutralizer voltage showed saturation against by ion thruster operation and miss-alignment of the thrust the gas flow rate, that is, increase of the flow rate decreased vector. The miss-alignment was estimated to less than 7.0 mm the neutralizer voltage in a certain range but further increase from the torque accumulation and it was in the range predicted had no effect for the voltage. However, the operation of by the ground measurement of the ion beam direction. February 15th showed strong dependence of the neutralizer Another expected disturbance torque was solar radiation voltage on the gas flow rate while the voltage was going up to pressure. That magnitude was also within the expected order 100 V. Dependency of neutralizer voltage on mass flow rate and currently its accurate measurement was continued. On the were summarized in Fig. 8. Operations in January and contrary, raw gas ejection from the ion thruster produced a

Jan 04 - Jan 09, 2015 Jan 10 - Jan 13, 2015

Screen current/mA Neutralizer voltage/V Feb 15 - Feb 24, 2015 Feb 25 - Mar 02, 2015

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Time (hh:mm) Mass flow rate/μg s-1

Fig. 7. Neutralizer voltage on February 15th, 2015. The neutralizer Fig. 8. Mass flow rate vs. neutralizer voltage. The dependency was voltage reached to 100 V, and high mass flow rate decreased the different in February 2015. status.

Pb_20 H. KOIZUMI et al.: Initial Flight Operations of the Miniature Propulsion System Installed on Small Space Probe: PROCYON

Screen current/mA Neutralizer current/mA Accelerator current/mA Microwave power/W Screeen voltage/200 V Neutralizer voltage/10 V Accelerator volrate/50 V Mass flow rate/10 μg s-1 8.00

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Fig. 9. Time histories of the ion thruster operations of 223 hours. All the telemetry data were averaged over 5 minutes and the data where screen current was higher than 1.0 mA were extracted and depicted in this graph. torque higher than expected. The torque associated with that 223 hours after carrying out the countermeasures for problems gas ejection was around minus X direction. The project team #01-05. Time histories during that ion thruster operation are concluded that it was due to the reflection of the gas by the shown in Fig. 9, where all of the telemetry data were averaged backside of the solar array panel. The momentum of raw gas over 5 minutes and the data set whose screen current was particle was little but the torque arm was much longer than the higher than 1.0 mA were extracted and depicted. The accelerated ion particles. accumulated operation time in this graph was calculated based Simultaneous operation of the cold-gas thrusters and the ion on those averaged data set and it was different from the actual thruster was successfully achieved. Gas ejection from the total operation time (summation of non-averaged data gave us cold-gas thrusters had potential to increase the gas pressure the total operation time of 223 hours). First 12 hours of the surrounding the spacecraft and to disturb the operation of the operation included a number of trials to settle the ion thruster. Particularly, the increased background of the ion above-mentioned problems. The steady operation of the ion thruster can affect the accelerator current. The project team thruster started after that. The averaged ion beam current was stopped the ion acceleration during the unloading maneuver at 5.62 mA corresponding to the thrust of 346 μN. The thrust the initial phase of the operation. However, this procedure coefficient was updated to 0.964 according to the thrust complicated spacecraft operation and had a risk of operational estimated from୘ the Doppler data. The beam current showed a miss. Then the team examined the simultaneous operation of trend of gradualߛ decrease by 1.7% over 200 hours. The mass the both thrusters at the later phase of the steady operation and flow rate (total flow rate to both of the ion source and the showed that operation had no interference. neutralizer) ranged from 57 – 67 μg/s. This variation was 5.2. Continuous operation of the ITU caused by the off-nominal method to solve the error control of The ion thruster accumulated the total operation time up to the pressure regulation system. Judgement to open or not open

Nominal control #1 (Opening: 48 ms, Limit: 45.0 kPa) from Feb 15 10:00

Nominal control #2 (Opening: 144 ms, Limit: 45.0 kPa) from Feb 22 07:30

Off-nominal control (Opening: 144 ms, Limit: 47.5 kPa, RV-CMD loop:200 s) from Feb 25 11:10

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50.5

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44.5 Pressure at the LPGT/kPa the at Pressure 42.5 0 1000 2000 3000 4000 5000 6000 7000 8000

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Fig. 10. Comparison of pressure regulations.

Pb_21 Trans. JSASS Aerospace Tech. Japan Vol. 14, No. ists30 (2016) the valve was once 200 s and variation was increased by 40 for Scientific Research (B), No. 25289304. times of the designed value. Different setting and control methods of pressure regulations are compared in Fig. 10. The Reference neutralizer voltage was decreased from 42 V to 29 V in 200 hours operation. Particularly, there were stepwise reductions 1) Micci, M. M. and Ketsdever, A. D.: Micropropulsion for Small from 45 to 60 hours. Spacecraft, American Institute of Aeronautics and Astronautics, Although that ion thruster operation and thrust generation Washington, D.C., 2000. 2) Goebel, D.M. and Katz, I.: Fundamentals of Electric Propulsion, were stable over 200 hours, that operation was far different Wiley, New Jersey, 2008. from the nominal condition. The major difference was the 3) Funase, R., Koizumi, H., Nakasuka, S., Kawakatsu, Y., mass flow rate of 61.4 μg/s, which was twice as high as the Fukushima, Y., Tomiki, A., Kobayashi, Y., Nakatsuka, J., Mita, nominal flow rate. Although this flow rate reduced the M. and Kobayashi, D.: 50kg—Class Deep Space Exploration specific impulse, the spacecraft had 2.50 kg xenon propellant Technology Demonstration Micro—spacecraft PROCYON, 28th Annual AIAA/USU Conference on Small Satellites, Utah State and it was enough for ion thruster. Another influence of the University, 2-7 August 2014. high flow rate was increase of the accelerator current up to 4) Kuninaka, H. and Hayabusa-2 Project: Deep Space Exploration more than twice. Although this flow rate was not examined of Hayabusa-2 Spacecraft, the 30th International Symposium on enough on the ground tests and life time was unknown, Space Technology and Science, 2015, 2015-k-61. keeping low neutralizer voltage required this off-nominal 5) Koizumi, H. Komurasaki, K. and Arakawa, Y.: Development of the Miniature Ion Propulsion System for 50 kg Small Spacecraft, operation. 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, , Atlanta, Georgia, 30 July - 01 August 2012, AIAA 6. Conclusion 2012-3949. 6) Koizumi, H. Komurasaki, K., Aoyama, J. and Yamaguchi, K.: Engineering Model of the Miniature Ion Propulsion System for The PROCYON project has developed the micropropulsion the Nano-satellite, Trans. JSASS Space Tech. Japan, 12, ists29 system I-COUPS since September 2013 and continued the (2014), pp.Tb_19-Tb_24. operations in deep space for more than six months. 7) Koizumi, H., Inagaki, T., Naoi, T., Kasagi, Y., Komurasaki, K., ・ Initial operations of the micropropulsion system, I-COUPS, Aoyama, J. and Yamaguchi, K.: Development of the Miniature of the ion thruster and cold-gas thrusters were completed Ion Propulsion System: MIPS for the 50-kg-Nano-satellite HODOYOSHI-4, Space Propulsion 2014, Cologne, Germany, on the space probe PROCYON. It will be the world’s first 19-22 May 2014. operation in deep space on a spacecraft smaller than 100 8) Nakasuka, S.: From Education to Practical Use of Nano-satellites kg. –Japanese University Challenge towards Low Cost Space ・ The high pressure gas system of 7.75 MPa xenon was Utilization –, 8th IAA Symposium on Small Satellites for Earth assembled by COTS products and has been in good health Observation, Berlin, April, 2011. 9) Nakasuka, S.: Low Cost Practical Micro/Nano-Satellites with through the launch and space operation more than six Reasonably Reliable Systems "HODOYOSHI" Concept, The 3rd months. Aerospace Innovation Workshop, University of Tokyo, Japan, ・ The reaction control system using xenon cold-gas thrusters 2011. has been successfully operated for six months to unload 10) Shiraska, S. and Nakasuka, S.: Realization of the Concept of the wheel momentum accumulated by the ion thruster Reasonably Reliable Systems Engineering in the Design of Nano-Satellites, Trans. Jpn. Soc. Aeronaut. Space Sci., operation and solar pressure. Aerospace Tech. Jph., 10 (2012), pp.7-12. ・ The miniature ion thruster achieved 223-hours operation in 11) Nagao, N., Yokota, S., Komurasaki, K. and Arakawa, Y.: total in interplanetary orbit without any interference with Development of a Two-dimensional Dual Pendulum Thrust the spacecraft system. Its thrust was evaluated to 345 μN Stand for Hall Thrusters, Review of Scientific Instruments, 7 from the Doppler shift data. (2007), 115108. 12) Koizumi, H. and Kuninaka, H.: Antenna Design Method for Low Power Miniaturized Microwave Discharge Ion thrusters, J. Acknowledgments JSASS, 57 (2009). 13) Koizumi, H. and Kuninaka, H.: Development of a Miniature The authors would like to express their deepest gratitude to microwave discharge ion thruster driven by 1 W microwave all of the members of the PROCYON project team in the power, Journal of Propulsion and Power, 26 (2010), pp.601-604. 14) Koizumi, H. and Kuninaka, H.: Performance Evaluation of a University of Tokyo and ISAS/JAXA for a lot of efforts to Miniature Ion Thruster μ1 with a Unipolar and Bipolar Operation, operate the word-first small space probe, PROCYON. This the32nd International Electric Propulsion Conference, work was supported by JSPS KAKENHI grant: Grant-in-Aid IEPC-2011-297.

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