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2. Tsuda,Y., Saiki,S., Mimasu,Y., Funase,R., “Sun-Earth Based Spin Axis Determination for Interplanetary Missions and Its Application to IKAROS”, 2011 AIAA/AAS Astrodynamics Specialist Conference, AAS11-448, Girdwood, Alaska, GUIDANCE, NAVIGATION AND CONTROL OF 2011.8.2. WORLD’S FIRST SOLAR POWER SAIL IKAROS 3. Tsuda,Y., ”An Attitude Control Strategy for Spinner Solar Sail”, 17th IFAC Symposium on Automatic Control in

Aerospace, WE-P02, 2007.6.25-29 Toulouse 1 2 3 1 4. Saiki, T., Nakaya, K., Yamamoto, T., Tsuda, Y., Mori, O., Kawaguchi, J., “Development of a Small-spin-axis Controller Yuichi TSUDA , Tomohiro YAMAGUCHI , Hitoshi IKEDA , Yuya MIMASU , 1 1 1 and Its Application to a Solar Sail Subpayload Satellite”, Transaction of the Japan Society for Aeronautical and Spae Takanao SAIKI , Hiroshi TAKEUCHI , Masaki NAKAMIYA Sciences, Space Technology Japan, Vol. 7, pp.25-32, 2009. 5. Saiki, T., Tsuda, Y., Funase, R., Mimasu, Y., Shirasawa, Y., IKAROS Demonstration Team, “Attitude Operation 1Japan Aerospace Exploration Agency, Sagamihara, Kanagawa, Japan Results of Solar Sail Demonstrator IKAROS”, 28th International Symposium on Space Technology and Science, ISTS 2The Graduate University for Advanced Studies, Sagamihara, Kanagawa, Japan ISTS-o-4-11v, Okinawa, Japan, 2011. 3 6. Yamamoto, T., Mori, O., Shida, M. and Kawaguchi, J., “Development of Gas-Liquid Equilibrium Thruster for the Small Satellite,” 25th Japan Aerospace Exploration Agency, Tsukuba, Ibaraki, Japan International Symposium on Space Technology and Science, 2006-k-32, Kanazawa, June 4-11, 2006. 7. Yamamoto,T., Mori,O., Sawada,H. and Funase,R., “System Safety Activity for IKAROS Spacecraft,” 61st International Astronautical Abstract Congress, IAC-10.D5.1.10, Prague, Sep.27-Oct.1, 2010. This paper summarizes the guidance, navigation and control of the world’s first solar power sail 8. Funase, R., Shirasawa, Y., Mimasu, Y., Mori, O., Tsuda, Y., Saiki, T., Kawaguchi, J., "On-orbit Verification of Fuel-Free Attitude IKAROS. During the 1.5 years of its interplanetary flight, IKAROS has carried out the guidance, Control System for Spinning Solar Sail Utilizing Solar Radiation Pressure", Advances in Space Research, Vol.48, Issue 11, pp.1740-1746, 2 2011. navigation and control experiments using the large solar radiation force generated by its 200 m solar 9. Mori,O., Sawada,H., Hanaoka,F., Kawaguchi,J., Shirasawa,Y., Sugita,M., Miyazaki,Y., Sakamoto,H. and Funase,R., "Development of sail. Since solar radiation pressure is the main controllable force for a solar sail, its modeling is the key Deployment System for Small Size Solar Sail Mission," Transactions of Japan Society for Aeronautical and Space Sciences, Space factor for a successful guidance. A precise solar radiation pressure modeling for this spinning solar sail Technology Japan, Vol. 7, No. ists26, pp. Pd_87-Pd_94 (2009). has been performed in order to support the navigation and guidance using the large membrane. Due to 10. Sawada, H., Mori, O., Okuizumi, N., Shirasawa, Y., Miyazaki, Y., Natori, M., Matunaga ,S., Furuya, H. and Sakamoto, H., "Mission the complexity of the sail surface and shape, the refinement of the SRP model is done after the Report on The Solar Power Sail Deployment Demonstration of IKAROS," 12th AIAA Gossamer Systems Forum, AIAA 2011-1887, Denver, USA (2011). deployment in space with radiometric measurements. This solar sail navigation is also supported by the 11. Shirasawa, Y., Mori, O., Miyazaki, Y., Sakamoto, H., Hasome, M., Okuizumi, N., Sawada, H., Furuya, H., Matsunaga, S., Natori, M. precise delta-DOR (Differential One-way Range) measurements. These in-flight demonstrations with and Kawaguchi, J., "Analysis of Membrane Dynamics using Multi-Particle Model for Solar Sail Demonstrator IKAROS," 12th AIAA IKAROS enable the future deep space exploration with solar sailing technique. Gossamer Systems Forum, AIAA 2011-1890, Denver, USA (2011). 12. Miyazaki, Y., Shirasawa Y., Mori, O., Sawada, H., Okuizumi, M., Sakamoto, H., Matunaga, S., Furuya, H. and Natori, M., 1. Introduction "Conserving Finite Element Dynamics of Gossamer Structure and Its Application to Spinning Solar Sail IKAROS," 52nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, AIAA 2011-2181, Denver, USA (2011). Japan Aerospace Exploration Agency (JAXA) launched the solar sail demonstration spacecraft 13. Tsuda, Y., Saiki, T., Mimasu, Y., Yamaguchi, T., Ikeda, H., Nakamiya, M., Takeuchi, H. and IKAROS Demonstration “IKAROS” (Fig.1) on May 21, 2010. IKAROS was launched together with JAXA’s Venus climate Team, “Modeling of Solar Radiation Pressure Effect for Trajectory Guidance of Spinner Solar Sailer IKAROS”, 28th orbiter “AKATSUKI (Planet-C)” as an interplanetary piggy-back payload. The launch vehicle was H2A International Symposium on Space Technology and Science, ISTS ISTS-o-4-10v, Okinawa, Japan, 2011.6.11. and was launched from Tanegashima space center. 14. Tsuda,Y., Saiki,T., Mimasu,Y., Funase,R., “Modeling of Attitude Dynamics for IKAROS Solar Sail Demonstrator”,

AAS/AIAA Space Flight Mechanics Meeting, AAS11-112, 2011.2.14, New Orleans 15. Tsuda, Y., Saiki, T., Funase, R., Shirasawa, Y., Mimasu, Y., “Shape Parameters Estimation of IKAROS Solar Sail JAXA has been proposing a concept of “Solar Power Sail” for future deep space exploration (Ref.1, 2). Using In-Flight Attitude Determination Data”, 12th AIAA Gossamer Systems Forum, 53-GSF-5-2, Apr.11,2011, Denver It combines the concept of solar sail (photon propulsion) with a larger power generation by flexible 16. Funase,R., Mimasu,Y., Chishiki,Y., Shirasawa,Y., Tsuda,Y., Saiki,T. and Kawaguchi,J., "Modeling and On-orbit Performance solar cells attached on the sail membrane. IKAROS is a precursor mission to demonstrate key Evaluation of Propellant-free Attitude Control System for Spinning Solar Sail via Optical Parameter Switching", Advances in the technologies requisite for the solar power sail concept, which are (1) deployment of large sail in space, Astronautical Sciences, Vol.142 (Also Proceedings of the AAS/AIAA Astrodynamics Specialist Conference, Girdwood, Alaaska, USA, (2) solar power generation by means of thin film solar cells attached on the sail, (3) confirming the July31 - August4 2011), pp.1737-1754, 2012. acceleration by solar radiation pressure acting on the sail and (4) demonstration of the interplanetary

Fig. 1. A picture of IKAROS taken by DCAM (deployable camera) on June, 14, 2010, five days after the successful sail depoloyment. guidance and navigation of the solar sail spacecraft.

IKAROS successfully deployed a 20 m span sail on June 9, and is now performing an interplanetary solar-sailing mission taking advantage of the Earth-Venus leg of the interplanetary trajectory (Ref.3). The spacecraft mass is 307kg and is equipped with a rectangular solar sail which weights 16kg with the minimum thickness of 7.5um. The solar sail is deployed and kept extended by centrifugal force due to the spacecraft spinning. Thus it does not have any rigid member to support the extension of the sail, enabling to realize very light and simple sail support mechanism. The deployment process was measured and recorded by several onboard equipments, such as cameras, attitude sensors and some surface sensor on the sail (Ref.3).

In this paper, a navigation and guidance technique for the solar sail is demonstrated with the IKAROS flight data. Section 2 describes the trajectory and attitude design concepts of the IKAROS spacecraft. Many constraints for the mission design are explained. In Section 3, three topics of the flight data analysis for the navigation and the guidance are described. Firstly, the solar sail acceleration is evaluated from the radiometric tracking data considering the attitude error due to the solar sail flexibility. Secondly, the guidance test is performed using the SRP force by changing the attitude. The navigation tests using the delta-DOR measurements are described afterwards. Finally, conclusions are given in Section 4.

2. Mission Design 2.1. Trajectory plan IKAROS was launched as one of the piggy back payloads of H-IIA flight #17, whose main payload was JAXA’s Venus explorer AKATSUKI (Planet-C). Hence the trajectory design of IKAROS was tightly 81 JSTS Vol. 27, No. 1

Fig. 1. A picture of IKAROS taken by DCAM (deployable camera) on June, 14, 2010, five days after the successful sail depoloyment. Fig. 2. IKAROS trajectory toward Venus. Left figure is drawn in J2000EQ inertial frame. Right guidance and navigation of the solar sail spacecraft. figure is drawn in Sun-Earth fixed frame.

IKAROS successfully deployed a 20 m span sail on June 9, and is now performing an interplanetary constrained by the primary payload, which was targeted to Venus. It was to take about solar-sailing mission taking advantage of the Earth-Venus leg of the interplanetary trajectory (Ref.3). six months to reach and fly by the planet, after which it was to continue its path by The spacecraft mass is 307kg and is equipped with a rectangular solar sail which weights 16kg with the orbiting the Sun while solar-sailing. The original ballistic trajectory injected by the minimum thickness of 7.5um. The solar sail is deployed and kept extended by centrifugal force due to launch vehicle was such that it exactly intercepted Venus (i.e. a typical Hohmann the spacecraft spinning. Thus it does not have any rigid member to support the extension of the sail, transfer). IKAROS attempted to escape from this Venus intercepting trajectory by means enabling to realize very light and simple sail support mechanism. The deployment process was of solar sailing (Fig. 2). measured and recorded by several onboard equipments, such as cameras, attitude sensors and some surface sensor on the sail (Ref.3). Although the trajectory was optimized exclusively for AKATUSKI, the Venus transfer orbit is quite beneficial for IKAROS, because (1) it is an interplanetary trajectory, in In this paper, a navigation and guidance technique for the solar sail is demonstrated with the IKAROS which the only major perturbation source is solar radiation pressure, thus pure solar flight data. Section 2 describes the trajectory and attitude design concepts of the IKAROS spacecraft. sailing is possible; (2) the spacecraft flies high solar-intensity region (Sun/Venus~0.7AU) Many constraints for the mission design are explained. In Section 3, three topics of the flight data compared with the Earth region (Sun/Earth~1AU), which makes it easier to evaluate the analysis for the navigation and the guidance are described. Firstly, the solar sail acceleration is solar sail performance. evaluated from the radiometric tracking data considering the attitude error due to the solar sail flexibility. Generally trajectory of solar sail spacecraft is coupled with attitude profile via sail Secondly, the guidance test is performed using the SRP force by changing the attitude. The navigation orientation with respect to the Sun. Thus the attitude profile of the solar sailer should be tests using the delta-DOR measurements are described afterwards. Finally, conclusions are given in designed to meet their trajectory requirements. IKAROS, on the other hand, did not have Section 4. a specific target to fly through, but it was rather optimized for the measurement and evaluation of solar sail performance (by experiencing various attitude w.r.t. the Sun. See 2. Mission Design Section 2.2). 2.1. Trajectory plan IKAROS was launched as one of the piggy back payloads of H-IIA flight #17, whose main payload was The V-infinity at the Earth departure was 4.4km/s. We had 17days of launch window JAXA’s Venus explorer AKATSUKI (Planet-C). Hence the trajectory design of IKAROS was tightly from May 17 through June 2, 2010 (UTC). The actual launch was on May 20, 2010 (UTC). 82

IKAROS was injected to the night side of the Earth. After experiencing 1.07AU apoapsis, the spacecraft flew to the daylight side of the Earth and reached the Venus on December 8, 2010. The fly-by V-infinity at the Venus was 3.3km/s, and the fly-by distance with respect to the Venus was 80,800km. IKAROS is continuing the interplanetary cruise after the Venus fly-by.

2.2. Attitude plan The long-term attitude control plan of IKAROS was designed with consideration for its antenna coverage. IKAROS has four antennas and two of them (XLGA1, XLGA2) are mainly used in the operations. Fig. 3 illustrates the antenna coverage of IKAROS. XLGA1 is attached on the top of IKAROS and XLGA2 is on the bottom panel. Both antennas have wide beams, but the antennas are purposely designed not to cover the sail directions (invisible zone) because there is a possibility that the reflected radio wave cause harmful effects to the communication.

Fig. 4 illustrates the long-term attitude control plan considered before the launch. The fat red line shows the attitude plan. The horizontal axis is time from the launch and the vertical axis corresponds to the “in-plane sun angle.” The in-plane sun angle is the angle between the spacecraft spin axis and the orbital-plane components of the sun direction vector. The contour lines show the earth angle. White areas of this figure are “out of communication area” where the link condition is bad. The attitude plan was designed to minimize the invisible zone crossing. The dashed line shows the flight result of IKAROS attitude. As it was found after the launch that the actual antenna coverage area was wider than expected, the restriction on the attitude during the extended mission phase has been relaxed.

3. Solar sail flight results In this section, the on-orbit solar sail navigation and guidance are presented. First, solar sail navigation is discussed by the means of orbit determination using radiometric tracking data. Next, the guidance and control of IKAROS are investigated with the flight results. In the third part, the performance of the VLBI (Very Long Baseline Interferometry) tone generator on the IKAROS spacecraft is described and presents the further possibility for accurate solar sail navigation.

3.1. Solar sail navigation and SRP estimation IKAROS successfully deployed its 200-m2 membrane on June 9th, 2010. The continuous Doppler signal has been received at Usuda Deep Space Center of JAXA. The Doppler signals are evaluated as observed value minus ballistic value (Fig. 5). The inclination of the 2-way Doppler signal in the figure 83 JSTS Vol. 27, No. 1

IKAROS was injected to the night side of the Earth. After experiencing 1.07AU apoapsis, the spacecraft flew to the daylight side of the Earth and reached the Venus on December Spin axis

8, 2010. The fly-by V-infinity at the Venus was 3.3km/s, and the fly-by distance with XLGA1 respect to the Venus was 80,800km. IKAROS is continuing the interplanetary cruise ±30deg Sail after the Venus fly-by. outside the coverage area (invisible zone)

XLGA2 2.2. Attitude plan The long-term attitude control plan of IKAROS was designed with consideration for its antenna coverage. IKAROS has four antennas and two of them (XLGA1, XLGA2) are mainly used in the Fig. 3. Antenna coverage of IKAROS. It is Fig. 4. Long-term attitude control plan of operations. Fig. 3 illustrates the antenna coverage of IKAROS. XLGA1 is attached on the top of designed not to cover the sail directions to IKAROS. The red fat line indicates the IKAROS and XLGA2 is on the bottom panel. Both antennas have wide beams, but the antennas are attitude plan, and the black dashed line purposely designed not to cover the sail directions (invisible zone) because there is a possibility that indicates the flight result. the reflected radio wave cause harmful effects to the communication. reveals the solar sail force acting on the IKAROS spacecraft. The most critical timing of the

deployment event was operated with 1-way Doppler mode so as to minimize the possibility of signal Fig. 4 illustrates the long-term attitude control plan considered before the launch. The fat red line loss of the ground station receiver due to possible rapid attitude changes induced by the sail shows the attitude plan. The horizontal axis is time from the launch and the vertical axis corresponds deployment. The time that spacecraft is operated with 1-way Doppler mode is indicated by blank in the to the “in-plane sun angle.” The in-plane sun angle is the angle between the spacecraft spin axis and figure. The Doppler signal suddenly inclined after the scheduled deployment time of the solar sail. The the orbital-plane components of the sun direction vector. The contour lines show the earth angle. solar sail deployment is confirmed from this fact. The spacecraft continues its flight and it approached White areas of this figure are “out of communication area” where the link condition is bad. The Venus on December 8, 2010, one day after the Akatsuki’s Venus orbit insertion maneuver. This delay is attitude plan was designed to minimize the invisible zone crossing. The dashed line shows the flight the consequence of the solar sailing. During the interplanetary flight, JAXA ground station collected the result of IKAROS attitude. As it was found after the launch that the actual antenna coverage area was radiometric data not only for the tracking, but also for the investigation of the solar sail dynamics. wider than expected, the restriction on the attitude during the extended mission phase has been relaxed. A precise solar sail force modeling is the key factor of successful navigation for the solar sail spacecraft.

Many prelaunch experiments were performed to confirm the performance of solar sail deployment and 3. Solar sail flight results stability (Ref.4, 5). Since the prelaunch SRP model is based on ground-based measurements, solar sail In this section, the on-orbit solar sail navigation and guidance are presented. First, solar sail navigation force model needs to be refined after launch. This is because the condition of the solar sail may be is discussed by the means of orbit determination using radiometric tracking data. Next, the guidance and changed by the deployment of the sail, degradation during the interplanetary flight, and the attitude control of IKAROS are investigated with the flight results. In the third part, the performance of the control maneuver. Additionally, the attitude error due to the uncertainties of the Earth angle VLBI (Very Long Baseline Interferometry) tone generator on the IKAROS spacecraft is described and measurement makes the significant errors on the estimation of the SRP acceleration. For the IKAROS presents the further possibility for accurate solar sail navigation. attitude determination, the Earth angle is calculated from the spin modulation on the Doppler

measurement. The accuracy of this measurement becomes worse when the Earth angle becomes near 90 3.1. Solar sail navigation and SRP estimation deg. Unfortunately, in some periods during the IKAROS flight, the Earth angle becomes large in order IKAROS successfully deployed its 200-m2 membrane on June 9th, 2010. The continuous Doppler to fulfill the Sun angle requirements. signal has been received at Usuda Deep Space Center of JAXA. The Doppler signals are evaluated as observed value minus ballistic value (Fig. 5). The inclination of the 2-way Doppler signal in the figure 84

Fig. 5. Filtered Doppler signal at the solar sail deployment.

Fig. 6. Estimated SRP acceleration with conventional and hybrid method

The impact of the attitude uncertainty is observed in the SRP estimation using the flight data. The SRP estimation is evaluated using the Normalized SRP acceleration (NSRP), which should be constant for each arc. The NSRP is calculated by normalizing the Sun distance to 1AU and the Sun angle to 0deg using the estimated SRP parameters (area and reflectivity). The NSRP is calculated by the conventional method, however the irregular value is estimated in the arc with the Earth angle between 60-120deg (red squares of Fig. 6).

In order to resolve the proper SRP acceleration, a hybrid estimation of the orbit and the attitude is applied (Ref.6). This method estimates both the orbit and the attitude from the 2-way Doppler and range data with the support of the Sun angle measurements. Since the Doppler measurement delivers the Earth direction component of the SRP force, the Earth angle can be solved by the Sun sensor and the Doppler measurements. A coupled effect between the orbital and attitude dynamics due to the large SRP acting on the solar sail is utilized for the formulation. The SRP effect includes a solar sail deformation model, 85 JSTS Vol. 27, No. 1

which is found to play the key role on the attitude dynamics of the spinning sail (Ref.7). For the purpose of this hybrid estimation, we used a linearly deformed sail (represented by two deformation angles, “torsion” and “outer-plane deflection”) and any other smaller wrinkles are supposed to be represented by degradation of the optical parameters (Ref.6). By solving this coupled dynamics with given Sun angle measurements, range and Doppler measurements, the SRP-perturbed orbit and the attitude history can be estimated. The estimation also provides the average optical parameters of the sail surface and the two sail deformation angles. The improvement caused by the hybrid estimation is described in Fig. 6. The figure shows the estimated NSRP with respect to the Earth angle of the

Fig. 5. Filtered Doppler signal at the solar sail deployment. evaluated arcs. It is found that the hybrid method shows the constant NSRP even when the Earth angle is close to 90 deg. Using this hybrid estimation, the SRP acceleration of the IKAROS is determined as 88 % of the predicted value. The predicted value assumes the full deployment and the reflectance determined by the ground-based experiments. The difference may be caused by the wrinkles on the sail surface. These results emphasize the importance of the on-orbit evaluation of the SRP model for a solar sail spacecraft.

3.2. Guidance and control with solar sail In this section, the adaptive guidance approach for solar sail spacecraft is presented. In this approach, the guidance parameters such as the area and reflectivity coefficient of the sail are updated properly. This approach will be used for future solar sail missions.

Fig. 6. Estimated SRP acceleration with conventional and hybrid method In the IKAROS mission, we guide the spacecraft to the target point in the B-plane (target plane) of Venus by changing the attitude of the solar sail using the reaction control system (RCS) and the novel The impact of the attitude uncertainty is observed in the SRP estimation using the flight data. The SRP thin film reflectivity control device (RCD). Fig. 7 shows the guidance results for the six attempts before estimation is evaluated using the Normalized SRP acceleration (NSRP), which should be constant for the Venus flyby. The points on the line indicate the target points at each time, and the others are actual each arc. The NSRP is calculated by normalizing the Sun distance to 1AU and the Sun angle to 0deg data from the orbit determination. In these attempts, we conducted the guidance with an accuracy less using the estimated SRP parameters (area and reflectivity). The NSRP is calculated by the conventional than three thousand kilometers in the B-plane. method, however the irregular value is estimated in the arc with the Earth angle between 60-120deg (red squares of Fig. 6).

Fig. 7. Guidance history

On the other aspect of the guidance for the solar sail, we evaluate the feasibility of the guidance and control considering the attitude drift motion due to the solar radiation pressure (SRP) torque. The attitude drift motion is induced by the balance between spin angular momentum and the SRP torque (Ref.7). Fig. 8 and 9 show the attitude drift motion and the spin rate history in the actual flight operation of IKAROS.

As the result of the attitude motion, the position of the spacecraft is evaluated in the virtual B-plane. For the virtual B-plane, the x-axis is defined as the velocity direction at the certain time in the spacecraft’s orbit, the z-axis is the normal direction to the orbit plane, and the y-axis is defined as the right-handed coordinate system. Fig. 10 shows the schematic image of the virtual B-plane and the projected positions on the plane propagated from each orbit determination point. The origin of the virtual B-plane is defined as the propagated position from the initial orbit determination point (August 25) to October 11, 2011, with the planned attitude sequence. As shown in Fig. 11, the propagated positions are inside of the propagated error ellipsoid from the previous orbit determination point. This means that the navigation and guidance are properly performed within the considered error. In addition, the final orbit determination point (green rectangle) is within 1800km from the originally planned position (red circle).

30 2 ] g e 20 1.8 Predicted Spin Rate d [

e Actual Spin Rate l 1.6 g

n 10 ] A

1.4 m n p u r S 0 [ 1.2

e e t n a a

l 1 R

- p -10 t n i i 0.8 b Predicted Attitude p r o Att-Det in Sept. 3 - 17 S - f -20 Att-Det in Sept. 20 - 24 0.6 o - t Att-Det in Sept. 28 - Oct. 7

u 0.4

O -30 Planned on Sept. 17 0.2 Planned on Sept. 27 -40 0 -30 -20 -10 0 10 20 30 40 0 5 10 15 20 25 30 35 40 In-orbit-plane Sun Angle [deg] Date from Sept. 3 [day] Fig. 8. Predicted and actual attitude evolution Fig. 9. Predicted and actual spin rate history

87 JSTS Vol. 27, No. 1

On the other aspect of the guidance for the solar sail, we evaluate the feasibility of the guidance and B 12000 R Planned Position control considering the attitude drift motion due to the solar radiation pressure (SRP) torque. The OD on Oct. 11 Propagated Positions 9000 with the actual sequence Only Spin Rate Control attitude drift motion is induced by the balance between spin angular momentum and the SRP torque BS BT 6000

(Ref.7). Fig. 8 and 9 show the attitude drift motion and the spin rate history in the actual flight operation ] m

k 3000 [ n of IKAROS. o i

t Aug. 25 c e

r 0 i

D Sept. 3

e n a l -3000 Sept. 28 p

As the result of the attitude motion, the position of the spacecraft is evaluated in the virtual B-plane. For Final OD - t i b Point OD Point 3 r o

- -6000 the virtual B-plane, the x-axis is defined as the velocity direction at the certain time in the spacecraft’s OD Point 2 f o - t

True Orbit u Deep color line: Error ellipsoid propagated covariance of OD orbit, the z-axis is the normal direction to the orbit plane, and the y-axis is defined as the right-handed e O Light color line: Error ellipsoid considered attitude error lan -9000 -P Propagation with SRP force Dashed-line: Error ellipsoid considered SRP parameter l B coordinate system. Fig. 10 shows the schematic image of the virtual B-plane and the projected positions ua irt in the planned sequence -12000 V -12000 -9000 -6000 -3000 0 3000 6000 9000 12000 on the plane propagated from each orbit determination point. The origin of the virtual B-plane is defined OD Point 1 In-orbit-plane Direction [km] as the propagated position from the initial orbit determination point (August 25) to October 11, 2011, with the planned attitude sequence. As shown in Fig. 11, the propagated positions are inside of the Fig. 10. Image of projected positions in the Fig. 11. Propagated positions from each OD point propagated error ellipsoid from the previous orbit determination point. This means that the navigation B-plane and guidance are properly performed within the considered error. In addition, the final orbit determination point (green rectangle) is within 1800km from the originally planned position (red 3.3. Precise Delta-DOR measurements circle). To determine the orbit under the continuous big influence of the non-gravitational perturbative force (i.e. solar radiation pressure), Delta-DOR (Delta-Differential One-way Range, Ref.8) observation is effective because sky plane position of the spacecraft can be directly and instantaneously measured by Delta-DOR observables without (or with less dependence on) a priori assumption for solar radiation

30 2 pressure model. In order to effectively perform Delta-DOR measurements, a signal generator of DOR ] g e 20 1.8 Predicted Spin Rate d

[ tones, which consist of multiple tones whose spanning bandwidth is about 28MHz, was developed and

e Actual Spin Rate l 1.6 g

n 10 ]

A installed to the spacecraft (Ref.9). A digital backend system for the ground stations which has

1.4 m n p u r S 0 [ 1.2

e e maximum output performance of 4-Gbps has also been developed to sample these wideband DOR t n a a l 1 R

- p -10 t n i i 0.8 tones. A total number of 24 inter-continental Delta-DOR experiments were carried out during July and b Predicted Attitude p r o Att-Det in Sept. 3 - 17 S - f -20 Att-Det in Sept. 20 - 24 0.6 o - t Att-Det in Sept. 28 - Oct. 7 August in 2010. Measurement accuracy of DOR observables for IKAROS was confirmed to be u 0.4

O -30 Planned on Sept. 17 0.2 typically 50-pico second level during this observation period. This is a 20 times improved precision Planned on Sept. 27 -40 0 -30 -20 -10 0 10 20 30 40 0 5 10 15 20 25 30 35 40 compared to the JAXA’s conventional deep space spacecraft such as Hayabusa and Akatsuki. In-orbit-plane Sun Angle [deg] Date from Sept. 3 [day]

Fig. 8. Predicted and actual attitude evolution Fig. 9. Predicted and actual spin rate history Fig. 12 shows the DOR and the quasar VLBI delay residuals observed on the Usuda-Canberra baseline on July 29 2010. Two radio quasars were used as calibrators (second quasar is weaker than first quasar but angularly nearer than the first one) to validate the delay bias cancelation effects by quasars. VLBI delay residuals are gradually changed and those trends are consistent for both quasars but a 7.8 nano seconds (corresponding to 2.3m) of systematic delay bias is clearly seen between quasar and IKAROS observables. This bias is the Delta-DOR observable, which reflects on the error 88

Fig. 12. Measured Delta-DOR observables on the Usuda-Canberra baseline of IKAROS’s reference orbit which is determined by Range and Doppler observables without Delta-DOR observations. A more detailed description of the newly developed system and results of combined orbit solution in which all of range, Doppler and Delta-DOR observables are included will be discussed in future papers.

4. Conclusions This paper summarized the mission design, navigation and guidance technique for the world’s first solar sail demonstrator IKAROS. Successful mission design allows IKAROS to survive more than 1.5 years in interplanetary space. This is three times more than the nominal mission phase. The solar sail performance of the IKAROS is evaluated using the hybrid estimation in order to correct the attitude error due to the systematic error of the attitude determination system. The results described the importance of the in-flight calibration of the SRP model for a solar sail spacecraft. JAXA is pursuing the realization of the “Solar Power Sail” technology for future interplanetary explorations, and the experience we got from IKAROS flight operation is the big step for realizing this new concept.

Acknowledgement We would like to appreciate the help of S. Taniguchi of Fujitsu Ltd. for providing us with the tracking data.

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Arakawa, M., and Sugita, M., “Static deployment of large membrane of spinning solar sail using a Fig. 12. Measured Delta-DOR observables on the Usuda-Canberra baseline balloon,” Advances in the Astronautical Sciences, Vol. 127 PART 1, Advances in the Astronautical

Sciences, 2007, pp. 1029–1040. of IKAROS’s reference orbit which is determined by Range and Doppler observables without 6. Yamaguchi, T., Mimasu, Y., Tsuda, Y., and Yoshikawa, M., “Hybrid Estimation of the Delta-DOR observations. A more detailed description of the newly developed system and results of Solar Radiation Pressure for a Spinning Solar Sail Spacecraft”, Engineering note, Journal combined orbit solution in which all of range, Doppler and Delta-DOR observables are included will of Spacecraft and Rockets, doi: 10.2514/1.A32387 , 2013 be discussed in future papers. 7. Tsuda, Y., Okano, Y., Mimasu, Y., and Funase, R., “On-orbit Sail Quality Evaluation Utilizing

Attitude Dynamics of Spinner Solar Sailer IKAROS,” 22nd AAS/AIAA Space Flight Mechanics 4. Conclusions Meeting, Charleston, USA, February, 2012, Paper AAS 12-211; Also Advances in the This paper summarized the mission design, navigation and guidance technique for the world’s first solar Astronautical Sciences, Vol.143, pp.1609-1626, 2012. sail demonstrator IKAROS. Successful mission design allows IKAROS to survive more than 1.5 years 8. Delta-DOR - Techinical Characteristics and Performance, Report Concerning Space Data System in interplanetary space. This is three times more than the nominal mission phase. The solar sail Standards, CCSDS 500.1-G-1. Green Book. Issue 1. Washington, D.C.: CCSDS, May 2013. performance of the IKAROS is evaluated using the hybrid estimation in order to correct the attitude 9. Takeuchi, H., Horiuchi, S., Phillips, C., Edwards, P., McCallum, J., Dickey, J., Ellingsen, S., error due to the systematic error of the attitude determination system. The results described the Yamaguchi, T., Ichikawa, R., Takefuji, K., Kurihara, S., Ichikawa, B., Yoshikawa, M., Tomiki, A., importance of the in-flight calibration of the SRP model for a solar sail spacecraft. JAXA is pursuing the Sawada, H., and Jinsong, P., Delta-DOR Observations for the IKAROS Spacecraft, in: Proceedings realization of the “Solar Power Sail” technology for future interplanetary explorations, and the of the 28th International Symposium on Space Technology and Science, 2011-o-4-14v, 2011. experience we got from IKAROS flight operation is the big step for realizing this new concept.

Acknowledgement We would like to appreciate the help of S. Taniguchi of Fujitsu Ltd. for providing us with the tracking data.