IKAROS and Extended Solar Power Sail Missions for Outer Planetary Exploration

Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28, pp. Po_4_13-Po_4_20, 2012 Original Paper

1) 1) 1) 1) 1) 1) By Osamu MORI, Yuichi TSUDA, Hirotaka SAWADA, Ryu FUNASE, Takanao SAIKI, Takayuki YAMAMOTO, 1) 1) 1) 1) 1) Katsuhide YONEKURA, Hirokazu HOSHINO, Hiroyuki MINAMINO, Tatsuya ENDO, Junichiro KAWAGUCHI and IKAROS Demonstration Team

1)Japan Aerospace Exploration Agency, Sagamihara, Japan

(Received June 27th, 2011)

The Japan Aerospace Exploration Agency (JAXA) makes the world’s first solar power sail demonstration of photon propulsion and thin film solar power generation during its interplanetary cruise by IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun). It deployed and spans a membrane of 20 meters in diameter taking the advantage of the spin centrifugal force. It accelerates and controls the orbit using solar radiation pressure successfully. This is the first actual solar sail flying an interplanetary voyage. This paper presents the summary of development and operation of IKAROS and introduces the outline of the extended solar power sail mission toward Jupiter and Trojan asteroids.

Key Words: Solar Sail, Solar Power Sail, Membrane, Deployment, IKAROS

1. Introduction film power generation. It is also the world’s first actual solar sail flying an interplanetary voyage as shown in Fig. 2. A Solar Sail 1,2) is a space yacht that gathers energy for IKAROS is not hybrid propulsion system, because ion propulsion from sunlight pressure by means of a membrane. A propulsion engine is not mounted. solar sail can move forward without consuming propellant as long as it can generate enough energy from sunlight. This idea of a solar sail was born about 100 years ago and we often find it in science fiction novels. The solar sail missions are studied in the world. A Solar Power Sail is a Japanese original concept that gets electricity from thin film solar cells on the membrane in addition to acceleration by solar radiation. A solar power sail Fig. 1. Extended solar power sail mission toward Jupiter and Trojan craft can save the fuel using a solar sail and it can also gain asteroids. the necessary electric power using a vast area of thin film Table 1. Spacecraft systems for Jupiter system mission. solar cells on the membrane even when it is away from the Mission Mass Isp Electric power sun. It can be a hybrid propulsion system with a solar sail by around Jupiter activating the ultra-high specific impulse ion engines with the JUNO 3625 kg 300 s 486 W (Hydrazine and (Solar panel) power generated by thin film solar cells. Thus solar power nitrogen sails are suitable for outer planetary exploration. tetroxide) The authors have studied an Extended Solar Power Sail Extended solar 2150 kg 6000~10000 s 5 kW 3,4) power sail craft (Ion engine) (Thin film solar mission toward Jupiter and Trojan asteroids via hybrid >10000 s cells) electric photon propulsion as shown in Fig. 1. The extended (Solar sail) solar power sail craft has larger delta-V and can generate larger electric power compared with the conventional

spacecraft as shown in Table 1. The mission proposal passed Membrane the Mission Definition Review and now is eligible to go into Thin film the pre-project phase (Phase-A). solar cells However solar sail and solar power sail are not realized. It is considered as high risk mission. We applied first for the 20m small technology demonstrator mission, IKAROS 5) as a Front-Loading of new key technical issues of extended solar power sail craft. The proposal was endorsed in fall, 2007. Using IKAROS (Interplanetary Kite-craft Accelerated by Radiation centrifugal force Of the Sun) demonstrates the membrane deployment and thin Fig. 2. IKAROS mission.

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Reference 5) shows the mission design of IKAROS and the relationship between IKAROS and extended solar power sail. On the other hand, this paper presents the summary of the development and operation of IKAROS and introduces the current mission design of the extended solar power sail based on the results of IKAROS mission.

2. IKAROS Mission

The success criteria of the IKAROS mission are summarized as follows; (a) Orbit (1) Deployment of Large Membrane Sail

-Deployment and Expansion of a Large Membrane in space Venus 6) Flying by Venus using similar mechanical device and procedures to those in Sun Full success Extended Solar Power Sail craft. (in six months) -Obtaining a number of data indicating the expansion status of Earth Minimum success the membrane. (in several weeks) 5) Demonstration of guidance, navigation (2) Generating Electricity by Thin Film Solar Cells and control by solar sail 4) Demonstration of photon -Demonstrating Solar Power from Thin Film Solar Cells propulsion by solar sail 1) Launch -Evaluating performance of thin film solar cells on the Spin separation 2) Radio telemetry ON 3) Membrane deployment Initial operation check Power generation by thin film solar cells membrane in space. Tip mass separation Separation camera experiments These two items belong to the Minimum Success Criteria. (3) Demonstrating Photon Propulsion (b) Mission sequence -Verification and Comparison of Reflectance with Specular & Fig. 4. Orbit and mission sequence. Diffuse Property. -Measurement of overall Reflectance with a rigorous relation 3. Development of IKAROS Spacecraft examination of the temperature and surface status. (4) Demonstrating Guidance, Navigation Control Skills for 3.1. IKAROS system Solar Sail Propulsion The major characteristic of IKAROS is shown in Table 2 -Navigation / Orbit Determination under continuous and small and Fig. 5. The spacecraft spans a square membrane sail acceleration. whose tip-to-tip length is 20 m long, and weights about 310 kg. -Acceleration Direction Control by Steering via appropriate The sail, spanned using the its spin centrifugal force, includes attitude control means. thin film solar cells which occupy approximately 5 % of the These latter two items belong to the Full Success Criteria. total area. The main body is a spinner and its shape is It was supposed that IKAROS was launched together with cylindrical. It carries a drum around which a membrane is the Venus Climate Orbiter, AKATSUKI in May 2010 as wound to be unfolded via a special mechanics aboard. shown in Fig. 3. The orbit and mission sequence was designed Where sail membrane is wound is the mission portion that as shown in the Fig. 4. After sun-pointing and spin separation is independent of the bus hub portion as shown in Fig. 5. The from the rocket, the initial operation check including the link spacecraft design started from fall of 2007, just 2.5 years prior establishment was performed. IKAROS span up to 25 rpm and to the launch as shown in Fig. 6 and the design process deployed the large membrane using the centrifugal force. The intentionally adopted No-EM (Engineering Model) power generation by thin film solar cells was confirmed after development strategy taking the advantage of relatively the deployment. It took several weeks to achieve the minimum sufficient mass margin indicated from the launch vehicle. success. The photon propulsion was verified and the guidance, Besides, the existing hardware surplus as well as reproduction navigation and control by solar sail were demonstrated for full of existing design was fully exploited to shorten the success mission in the next six months. IKAROS flew by development period and to cut cost down as shown in Fig. 7. Venus at the beginning of December 2010. This paper introduces the major development components: sail, deployment system, gas-liquid equilibrium propulsion system and optional equipments. Table 2. Major characteristics. Diameter 1.6 m * height 0.8 m Body (cylinder shape) Configuration Square of side 14m / cross section 20 m Membrane (after deployment) Wet mass 310 kg (including membrane mass) Dry mass 290 kg (including membrane mass) Mass Membrane 15 kg (including 2 kg of 4 tip masses) IKAROS H-IIA launch vehicle Launch mass Attitude control system Spin Fig. 3. Launch of IKAROS by H-IIA launch vehicle.

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8) 1.6m SAP/Upper deck Reflection control device : Orbit control requires changing the angle (facing direction) of the membrane to the sun as Mission module 0.8m shown in Fig. 12. Liquid crystal films are loaded near the tips Thrust tube Middle deck of the membrane. They can be switch specular and diffuse Thruster deck Lower deck reflection by power ON and OFF according to the spin rate to

Fig. 5. IKAROS appearance. control the sun angle as shown in Fig. 13. ALADDIN: PVDF (PolyVinylidene Di-Fluoride) film is

FY 2007 2008 2009 2010FY22 attached on the back side as dust counter. It is the optional MonthMonth equipment which is not for the mission of solar power sail Development Critical Design Flight Model Integrated Test phase demonstration. ▲ Rocket I/F FIX ▲CDR for ▲CDR for Flight Launch ▲ Tip mass: A 0.5 kg weight is connected to each tip of the Configuration FIX bus portion mission portion Operation Fig. 6. Development schedule. membrane. It supports the deployment of the membrane. Tether and harness: Tether connects the membrane with the main body mechanically and harness connects them electrically. Reuse In addition, electric charge-measuring patches and Mission system Reproduction 必要に応じてADRV数を拡張するミッション系レートジャイロRate gyro （MPU I/Fには影響ない） VLBI-LGAVLBI-LGA ×3軸 CAM#1 RAM1 RAM1 CAM#1 thermometers are equipped on the membrane. ADRV CAM#1 RAM#1 RAM#1 CAM#1 Development ADRVADRV ADRV CAM-H#1CAM_H #1 DCAM#1DCAM_ ActuatorActuator1 1 XSSPA ActuatorActuator2 2 CAM#2CAM#2 RAM#2RAM2 RAM#2RAM2 CAM#2CAM#2 Actuator3 SSAS-ESSAS-S SSAS-S Actuator 3 CAM-H#2CAM_H #2 DCAM#2DCAM_ NDN D ActuatorActuator4 4 XVLBIXVLBI ActuatorActuator5 5 CAM#3 RAM3 ActuatorActuator6 6 CAM#3 RAM#3 PVDF ActuatorActuator7 7 CAM-H#3CAM_H #3 PVDF Actuator8 Actuator 8 FSA ActuatorActuator9 9 CAM#4CAM#4 RAM#4RAM4 ActuatorActuator10 10 ＳiSi DC CAM_H #4 DCAM-SEPDCAM-SEP TLMTLM CAM-H#4 RCD ADRV Motor RCD Env.MonEnv.Mon RAMRAM PWR ADRVADRV IF MDRV GAP TMPx12TMP ＤＲＵ CAM-C SAIL-IFSAIL IF DRU CMDCMD サンプレゼンス MPUMPU DR

Heater推進系ヒータ制御 control PCDUPCDU SAPSAP SSRSSR 各機器分離 Switchスイッチ Tanktank RCS

Pressure Sensor Fill/DrainFill/Drain ValveValve P Pressure Sensor DHU 中利得アンテナ BAT MGA 低利得アンテナ1LGA1 BAT DHU Filter (送信/受信) Filter (送信/受信) Latching LatchingValve Valve TLMTLM Solenoid XDIPXDIP SSPASSPA XTRPXTRP Valve XSW-A CMD XSW-B XSW-A CMD Solenoid Valve PSU ThrusterThruster 低利得アンテナ2 PSU A系（主系） B系（冗長系） (送信/受信)LGA2

Fig. 7. System block-diagram. Fig. 8. Sail shape and layout.

3.2. Sail A solar sail requires a thin film mirror that is extremely 20m light and strong enough to hold a vast area. The sail of IKAROS is made of ultra-thin polyimide resin whose thickness is 7.5 μm (about 1/10 of a human hair.) It has a high tolerance for space environment requirement. Fig. 8 shows the sail shape and equipment layout. 6) Membrane : The shape of the membrane is a square whose Polyimide-2 diagonal distance is 20 m. The direction of folded lines is perpendicular to the direction of the centrifugal force. It is Polyimide-1 made of two kinds of polyimide resin as shown in Fig. 9. Polyimide-1 is an existing film and it needs to be connected to each other by adhesive. On the other hand, polyimide-2 is Fig. 9. Membrane size and polyimide resin. thermoplastic polyimide which is developed and manufactured by JAXA newly for the extended solar power Front side sail mission. It can be fused with each other by heater control (Facing to the sun) without adhesive. If the membrane is made of thermoplastic polyimide, the membrane becomes lighter and stronger and can decrease outgassing. Aluminium is vapor deposited on front side of the membrane as shown in Fig. 10, in order to reflect sunlight more efficiently. In addition, the film is Back side reinforced in such a way as to prevent it from splitting all the (Not facing to the sun) way if it's ripped. If it is ripped, the solar sail performance declines slightly, but it can still continue its space travels. Thin film solar cells 7): a-Si (amorphous silicon) solar cells as shown in Fig. 11 are attached to certain areas of the Fig. 10. Quarter of the square membrane. membrane. They generate almost 300 W. The area ratio is 5 %.

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stage deployment and second stage deployment. (1) Tip mass separation 220 mm In the folded configuration, each quarter of the membrane is line-shaped and rolled up around the main body. After the 330 mm separation from the launch vehicle with the slow spin at 5 rpm,

Fig. 11. Thin film solar cells. IKAROS spins down to 2 rpm using RCS in order to decrease the centrifugal force of four tip masses. They are separated at the same time by activating the tip mass separation Changed orbit mechanism. Sunlight force received (2) First stage deployment Original orbit by the solar sail IKAROS spins up to 25 rpm using RCS to increase the centrifugal force. In the first stage deployment, each quarter of Solar sail the membrane is extracted like a Yo-Yo despinner and the membrane forms a cross shape. If the deployment is Sunlight performed dynamically, each quarter is twisted around the To go away from the sun To go closer to the sun main body again just after the deployment. Therefore the first stage deployment is performed quasi statically by activating Fig. 12. Theory of orbit control using solar sail. the rotation guides that hold the membrane through the relative rotation mechanism (motor drive). The membrane is deployed slowly and gradually through centrifugal force and the spin rate dwindles as the relative rotation mechanism (Sun light) activates. After the completion of the first stage deployment, Incident ray Reflected light Incident ray Reflected light the cross shape is maintained by rotation guides.

Power ON Power OFF (Specular reflection) (Diffuse reflection) (Torque) Tip mass separation F F2 1 z 5rpm 2rpm Solar pressure F2 < F1 Separation from Spin down launch vehicle (RCS)

x y Tip mass First stage deployment (quasi static) 25rpm 5rpm OFF Spin up (RCS)

ON Rotation guide Second stage deployment (dynamic) 2.5rpm 5rpm Fig. 13. Attitude control by reflection control device.

3.3. Deployment system The most difficult mission of IKAROS is the deployment of membrane sail. There are two types of solar sail. First is the mast type 9,10) which uses some rigid support structure to Fig. 14. Deployment method. deploy and maintain the sail. The other is the spin type which uses spinning centrifugal force. The deployment motion and Sun sensor φ60 mm attitude control of mast type are simpler than those of spin Separation camera type. On the other hand, spin type can be accomplished with lighter-weight mechanisms than mast type because it does not require rigid structural elements. Thus spin type should be 2 separation cameras 60 mm selected in case of a large membrane. IKAROS demonstrated 0.8 m Monitor camera a spin type solar power sail whose area was 200 m2 for 2 extended solar power sail whose area is 2000 m . The RCS feasibility was verified by a lot of experiments and numerical 1.6 m 4 monitor cameras simulations 11,12). The proposed deployment method 13) of the sail is shown in Fig. 15. Sensors checking motion of main body and membrane. Fig. 14. It consists of three processes: tip mass separation, first

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(3) Second stage deployment dust counter using PVDF (PolyVinylidene DiFluoride). The second stage deployment is performed dynamically by Engineering mission: activating the rotation guides and releasing the hold of the VLBI (Very Long Baseline Interferometry) 19): High accuracy membrane. The membrane expands quickly to form a square orbit determination test by the DDOR (Delta Differential shape. The spin rate is decreased in the first and second stages One-way Range) technology. of the deployment, because the inertial momentum of the sail is increased. The motion of the main body and the membrane is checked by a sun sensor, a three-axis gyro, range rate, four monitor cameras, two separation cameras 14) and so on as shown in Fig. GAP detector ALADDIN sensor VLBI-Tx and LGA 15. 3.4. Gas-liquid equilibrium propulsion system Fig. 17. Optional equipments. IKAROS has a reaction control system (RCS), shown in Fig. 16, to spin up itself and control its attitude through all the 4. IKAROS Operation mission life. Because IKAROS is a sub-payload of AKATSUKI, IKAROS has some difficulty in use of IKAROS was launched together with AKATSUKI aboard hazardous propellant such as hydrazine due to the safety H-IIA F17 at 6:58:22 am on May 21, 2010 (Japan Standard 15) requirement . In addition, IKAROS’s required maneuver is Time) from the Tanegashima Space Center. The initial not so little that the size of tank will be too large to be operation was performed to achieve the minimum success mounted on IKAROS if IKAROS uses cold gas jet such as until June 29. The normal operation was conducted for full GN2. success. Therefore IKAROS RCS adopted Gas–Liquid Equilibrium 4.1. Initial operation 16) Propulsion System with chlorofluorocarbon alternative Summary of the initial operation is as follows. HFC-134a. This system stores HFC-134a as liquid phase in May 21: Launch, Spin separation (=> 5 rpm) the tank, extracts the vapor of HFC-134a from the tank, and May 23~25: Spin down (=> 2 rpm) ejects the vapor from the thruster nozzle. This system is good May 26: Tip mass separation at cost and schedule because cold gas through the nozzle May 27~31: Spin up (=> 25 rpm) realizes simplified thruster system. Also this system requires June 2~8: First stage deployment (=> 5 rpm) not so large space because the propellant is stored in liquid June 9: Second stage deployment (=> 2.5 rpm) phase and pressurized by its own vapor through the mission June 10: Solar power generation through thin film solar cells (no pressure degradation in accordance with propellant June 14: 1st separation camera experiment consumption). Besides, the small satellites are prohibited to June 16~18: Spin down (=> 1 rpm) use hazardous propellant such as hydrazine due to the strict June 19: 2nd separation camera experiment safety requirements. Therefore this gas-liquid equilibrium June 21~25: Start up of optional equipments propulsion system is very suitable for small satellites. June 29: Transition to normal operation From May 26 to June 9, the membrane was deployed. Fig. Main body 18 shows the data of the tip mass separation. Due to conservation of angular momentum, the spin rate was decreased, when these tip masses were separated from the main body. The angular velocities of x, y-axes were zero Tank THR 2 THR 1 before and after this event. If one of tip mass does not be separated, the nutation motion of the main body was occurred. THR 3 Thus four tip masses were separated at the same time, despite THR 4 a lot of locks for flight security. The pictures by four monitor Fig. 16. Thruster layout. cameras also confirm that four tip masses were separated successfully. 3.5. Optional equipments Fig. 19. shows the data of the first stage deployment. The In addition to solar power sail mission, IKAROS performs rotation guides hold the membrane. By rotating them the science observation and engineering experiment using relatively, the membrane was deployed gradually and optional equipments as shown in Fig. 17. They are not for the similarly. The pictures by four monitor cameras confirm that mission of solar power sail demonstration and this paper the membrane was deployed smoothly and formed a cross introduces them compactly. shape after the first stage deployment. Science missions: Fig. 20 shows the data of the second stage deployment. Just GAP (GAmma-ray burst Polarimeter) 17): Gamma-ray burst after deployment start, the vibrations of angular velocities observation experiment by a polarized light detector. were occurred. However they were dumped gradually. The ALADDIN (Arrayed Large-Area Dust Detector for spin rate of z-axis was decreased due to conservation of Interplanetary Space) 18): Dust distribution is observed by a angular momentum. Thus the membrane was deployed dynamically and symmetrically. The pictures by four monitor

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cameras confirm that the membrane, tethers and harnesses 6 were not gotten hung up on the released rotation guides. 二次展開実行Deployment start On June 10, the membrane deployment and solar power 0.2 generation by thin film solar cells were confirmed. 5 On June 14 and 19, two separation cameras successfully 0 took impressive images as shown in Fig. 21. The membrane 4 was expanded to form a square shape without damages. It was x-axis y-axis -0.2 shining in the dark space. The second separation camera made 3 sure that the reflection control device was working properly. Z Spin Rate , Spin [rpm] Z

How ON and OFF reflection control devices looks like in z-axis -0.4 Rate , Spin [rpm] X,Y 2 X軸スピンレート orbit was confirmed. Y軸スピンレート Spin rate (z-axis) [deg/s] (z-axis)rate Spin Z軸スピンレート 1 -0.6 5800 5900 6000 6100 6200 Tip masses were separated from the main body [deg/s] (x,y-axes) velocities Angular 12 Time , [sec] Z axis X axis 0.4 (a) Spin rate and angular velocity of main body (June 9, 2010) 11.9 Y axis z-axis 11.8 0.2

11.7 x-axis

0 Z Spin Rate , [deg/s] , SpinZRate 11.6 [deg/s] , SpinX,YRate Harness y-axis Tether Spinrate (z-axis) [deg/s] Rotation 11.5 -0.2 guide 3900 4000 4100

Time , [sec] Angular velocities (x,y-axes) [deg/s] (b) Pictures taken after second stage deployment by monitor camera 1 (a) Spin rate and angular velocity of main body (May 26, 2010) (June 10, 2010) Fig. 20. Second stage deployment.

Tip mass

Membrane

(b) Pictures taken after tip mass separation by monitor camera 1 (May 29, 2010) Fig. 18. Tip mass separation.

Monitor camera 1

(a) Pictures shot by first separation camera (June 14, 2010)

(a) Pictures taken in the middle of first stage deployment by monitor camera 1 (June 4, 2010)

ON OFF (Specular reflection) (Diffuse reflection) (b) Picture shot by second separation camera (June 19, 2010) Fig. 21. Separation camera experiments.

4.2. Normal operation After the initial operation, the normal operation was (b) Pictures taken after first stage deployment by monitor camera 1 conducted for full success. The photon acceleration was (June 8, 2010) Fig. 19. First stage deployment. confirmed in the course of determining its precise orbit after its sail deployment. Fig. 22 shows the difference between

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calculated value (without photon acceleration) and observed Fig. 24. The distance at closest point to Venus was 80800km. value (result of range rate measurement) of IKAROS's IKAROS was arrived at Venus one day later than AKATSUKI, line-of-sight velocity on June 9. For about one hour around because IKAROS was decelerated by solar sail. 9:36, when the second stage deployment was in operation, In the normal operation, the missions of optional data was lost; however, you can clearly see that the graph line equipments (GAP, ALADDIN and VLBI) were also starts inclining after the second-stage deployment compared to performed steadily between the solar power sail mission. the flat line prior to the event. The calculated value does not incorporate photon acceleration, thus, the velocity difference 5. Extended Solar Power Sail Mission to Jupiter and becomes a flat line if there is no photon acceleration. It is not Trojan Asteroids equal to zero due to the error of estimated orbit motion. With photon acceleration, the line inclines. The thrust force is In the next step of IKAROS, the extended solar power sail calculated as 1.1 mN and it is nearly equal to the designed craft is designed as shown in Fig. 25. It consists of the main value. This is the world’s first data of solar sailing. spacecraft and Jovian orbiter. It uses the world’s first hybrid IKAROS continued to be accelerated by solar sail. The photon / ion propulsions taking the advantage of thin film delta-V of 100 m/s was achieved within half a year. photo-voltaic technology. The mission has very new The attitude control performance using the reflection multi-purposes. control device was successfully accomplished. Fig. 23 shows First of all, the mission aims at the exploration of the Trojan the evaluation result of the attitude control experiment by the asteroids for the first time in the world as shown in Fig. 26. reflection control device on July 13 based on sun angle. The The spacecraft is supposed to drive its ultra-high specific solar angle was gradually increasing without control, while it impulse ion engines aboard with the combination of the Earth was decreasing when the reflection control device was used to gravity assist. EDVEGA (Electric-DV Earth Gravity Assist) control the attitude toward the sun direction. Very smooth strategy is best to maximize science payload, compared with attitude maneuver was achieved. The reflection control double EDVEGA and direct departure strategies. By 5 years enabled IKAROS to change its attitude for orbit control. flight, the mother spacecraft jettisons small Jovian orbiter, which independently decelerates the orbital speed to stay around the Jupiter for scientific observation. The mother Filtered Doppler O-C Second stage 100 Period of data loss spacecraft leaves the Jupiter to rendezvous to Trojan Asteroids deployment due to deployment 90 (09:36 UTC) operation by additional 5 years flight. Here are summarized major 80 characteristics of the extended solar power sail mission; It is 70 60 the 50 1. World’s First Combined Jovian Orbiter / Flyby Mission, Velocity Difference Velocity value (=Observed – calculated value) [mm] 09:00 09:30 10:00 10:30 11:00 11:30 12:00 12:30 2. World’s Highest Performance Ion Engines, Time on June 6 [Universal Time] 3. World’s First Photon / Electric Hybrid Sail Propulsion, 4. World’s First Background Emission Mapping, Fig. 22. Confirmation of photon propulsion by range rate measurement. 5. World’s First Access to Trojan Asteroids, 6. World’s First Formation Flight in Jovian Magnetosphere. The proposed extended solar power sail craft is innovative as it carries the following new technology demonstrations. 1. A Large Membrane Space Structure including Deployment strategy, 2. A Hybrid Propulsion using both Photon and Ultra High Performance Ion Engines, 3. Thin Film Solar Cells, 4. Reaction Control at very Low Temperature, 5. An Integrated Propulsion / Power System using Fuel Cells. Fig. 23. Attitude control performance using reflection control device. 6. Formation Flight in Jovian system, 7. Electric Delta-VEGA (Delta-V Earth Gravity Assist) Venus technique for outer solar system, 8. Ultra Stable Oscillator for 1-way range or VLBI (Very Long Baseline Interferometry) orbit determination, 9. Ka-band Communication for Interplanetary Missions, 10. Radiation-Resistant Technology for Jovian Orbiter, 11. Membrane Phased Array Antenna It is full of new technologies requisite for future solar system voyages. A part of technologies 1-3 was demonstrated by IKAROS. Fig. 24. Flying by Venus (Dec. 8, 2010). Not only the technology demonstration, there are more new IKAROS flew by Venus on December 8, 2010 as shown in innovative scientific purposes carried by this spacecraft. This

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spacecraft performs the cruise infrared astronomy observation Jupiter System Exploration, JpGU International Symposium 2011, 2011, PPS001-04. together with the ecliptic dust detection, and also makes 5) Mori, O., Sawada, H., Hanaoka, F., Kawaguchi, J., Shirasawa, Y., magnetosphere observation of Jupiter and visits Torojan Sugita, M., Miyazaki, Y., Sakamoto, H. and Funase, R.: asteroids. The scientific objectives are shown in Fig. 26. This Development of Deployment System for Small Size Solar Sail demonstrator aims at the first-class planetary science and Mission, Trans. JSASS Space Tech. Japan, 7, ists26 (2009), pp.Pd_87-Pd_94. space physics in addition to the new technologies. 6) Yokota, R., Miyauchi, M., Suzuki, M., Andoh, A., Kazama, K., JAXA has been seriously investigating the spacecraft Iwata, M., Ishida, Y., Shimamura, H. and Ishizawa, J.: Heat development and the mission proposal passed the Mission Sealable, Novel Asymmetric Aromatic Polyimide Having Excellent Definition Review in 2005. The extended solar power sail Space Environmental Stability and Application for Solar Sail, IKAROS Membrane, 28th International Symposium on Space craft is supposed to be launched in 2020 on the basis of Technology and Science, 2011, 2011-o-4-02v. IKAROS achievement. 7) Tanaka, K., Soma, E., Yokota, R., Shimazaki, K., Tsuda, Y., Kawaguchi, J. and IKAROS Demonstration Team: Development of Thin Film Solar Array for Small Solar Power Demonstrator “IKAROS,” 61st International Astronautical Congress, 2010, IAC-10.C3.4.3. 8) Funase, R., Shirasawa, Y., Mimasu, Y., Mori, O., Tsuda, Y., Saiki, T. and Kawaguchi, J.: On-orbit Verification of Fuel-Free Attitude Control System for Spinning Solar Sail Utilizing Solar Radiation Pressure, Advances in Space Research, 48 (2011), pp.1740-1746. 9) Greschik, G. and Mikulas, M. M.: Design Study of a Square Solar Sail Architecture, J. of Spacecraft and Rockets, 39 (2002), Fig. 25. Extended solar power sail craft system. pp.653-661. 10) Hinkle, J. D., Warren, P. and Peterson, L. D.: Geometric Imperfection Effects in an Elastically Deployable Isogrid Column, J. of Spacecraft and Rockets, 39 (2002), pp.662-668. L4 Trojan asteroids 11) Tsuda, Y., Mori O., Takeuchi S. and Kawaguchi, J.: Flight Result (+10 years) and Analysis of Solar Sail Deployment Experiment using S-310 ⑥ Sun Sounding Rocket, Space Technol., 26 (2006), pp. 33-39. Earth 12) Nishimaki, S., Mori O., Shida M. and Kawaguchi, J.: Stability and (1AU) Control Response of Spinning Solar Sail-craft containing A Huge Mainbelt ~ 3AU ⑤ (+3 years ) Membrane, 57th International Astronautical Congress, 2006, ① - ④ Jupiter ~5.2 AU IAC-06-C1.1.07. (+5 years) 13) Sawada, H., Mori, O., Okuizumi, N., Shirasawa, Y., Miyazaki, Y., I. Cruising phase to and beyond Jupiter II. at Jupiter and Trojan asteroid Natori, M., Matunaga, S., Furuya, H. and Sakamoto, H.: Mission ① IR astronomy (CIB and zodiacal light obs.) ⑤ In-situ Jovian science (plasma observation) ② Dust observation ⑥ Rendezvous observation of Trojan asteroid(s) Report on The Solar Power Sail Deployment Demonstration of ③ High-E astronomy (Gamma-ray burst) IKAROS, 12th AIAA Gossamer Systems Forum, 2011, ④ Mainbelt asteroid flyby observation AIAA-2011-1887. Fig. 26. Mission sequence of extended solar power sail. 14) Matunaga, S., Inagawa, S., Nishihara, T., Kimura, S., Sawada, H., Mori, O., Kitamura, K. and Structure Team IKAROS: On-Orbit Demonstration of Deployable Camera System for Solar Sail 6. Conclusions IKAROS, 28th International Symposium on Space Technology and Science, 2011, 2011-o-4-07v. This paper presents summary of the development and 15) Yamamoto, T., Mori, O., Sawada, H. and Funase, R.: System operation of IKAROS. IKAROS was launched on May 21, Safety Activity for IKAROS Spacecraft, 61st International Astronautical Congress, 2010, IAC-10.D5.1.10. 2010 and it deployed the membrane successfully and 16) Kishino, Y., Tamura, M., Yamamoto, T. and Mori, O.: demonstrates photon propulsion and guidance, navigation and Development of Gas-Liquid Equilibrium Propulsion System for control using solar sail during its interplanetary cruise. IKAROS RCS, 61st International Astronautical Congress, 2010, IKAROS flew by Venus on December 8, 2010 and continues IAC-10.B4.5.3. 17) Yonetoku, D., Murakami, T., Gunji, S., Mihara, T., Sakashita, T., the first actual solar sail flying. The extended solar power sail Morihara, Y., Kikuchi, Y., Takahashi, T., Fujimoto, H., Toukairin, mission toward Jupiter and Trojan asteroids is planning based N., Kodama, Y., Kubo, S., and IKAROS Demonstration Team: on the results of IKAROS mission. Gamma-Ray Burst Polarimeter (GAP) aboard the Small Solar Power Sail Demonstrator IKAROS, Publ. Astron. Soc. Japan, 63 (2011), pp.625-638. References 18) Yano, H., Tanaka, M., Okamoto C., Hirai T., Ogawa N., Hasegawa S., Iwai T. and Okudaira K.: Cosmic Dust Detection by the 1) Mclnnes, C. R.: Solar Sailing, Technology, Dynamics and Mission IKAROS-Arrayed Large-Area Dust Detectors in Interplanetary Applications, Springer-Praxis, 1999. Space (ALADDIN) from the Earth to Venus, The 42nd Lunar and 2) Leipold, M. and Garner, C. E.: Solar Sails – Exploiting the Space Planetary Science Conference, 2011, 2647. Resource of Solar Radiation Pressure, ESA Workshop on Space 19) Takeuchi, H., Horiuchi, S., Phillips, C., Edwards, P., McCallum, J., Exploration and Resources Exploitation - Explospace, 1998. Dickey, J., Ellingsen, S., Yamaguchi, T., Ichikawa, R., Takefuji, 3) Kawaguchi, J.: A Solar Power Sail Mission for A Jovian Orbiter K., Kurihara, S., Ichikawa, B., Yoshikawa, M., Tomiki, A., and Trojan Asteroid Flybys, 35th Scientific Assembly of the Sawada, H. and Jinsong, P.: Delta-DOR Observations for the Committee on Space Research, 2004, COSPAR04-A-01655. IKAROS Spacecraft, 28th International Symposium on Space 4) Funase, R., Mori, O., Tsuda, Y., Sawada, H., Saiki, T. and Technology and Science, 2011, 2011-o-4-14v. Kawaguchi, J.: Development of Solar Power Sail System for Future

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