Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28, pp. To_4_1-To_4_6, 2012

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Attitude Operation Results of Solar Sail Demonstrator IKAROS

1) 1) 2) 2) 2) By Takanao SAIKI , Yuichi TSUDA , Ryu FUNASE , Yuya MIMASU , Yoji SHIRASAWA and IKAROS Demonstration Team

1)The Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency, Sagamihara, Japan 2)Space Exploration Center, Japan Aerospace Exploration Agency, Sagamihara, Japan

(Received June 27th, 2011)

This paper shows the attitude operation results of Japanese interplanetary solar sail demonstration spacecraft IKAROS. IKAROS was launched on 21 May 2010(JST) aboard an H-IIA rocket, together with the AKATSUKI Venus climate orbiter. As IKAROS is the secondary payload, the development cost and period were restricted and the onboard attitude system is very simple. This paper introduces the attitude determination and control system. And as IKAROS is spin type spacecraft and it has the large membrane, the attitude control is not easy and it is very important to determine the long-term attitude plan in advance. This paper also shows the outline of the IKAROS attitude operation plan and its operation results.

Key Words: Solar Sail, Attitude Operation Result, IKAROS

1. Introduction control torque. Therefore, it is important to determine the long-term attitude control plan in advance and operate IKAROS is a solar power sail technology demonstration according to the plan. In this paper, we show the attitude spacecraft developed by Japan Aerospace Exploration Agency. operation plan and results of IKAROS. The spacecraft was launched on May 21st , 2010 together with the AKATSUKI Venus climate orbiter as the secondary payload. The concept of solar power sail has been proposed for years Full success in JAXA. The solar power sail is the hybrid propulsion system. (a half year) It combines the photon propulsion and the ion engine Minimum success (several weeks) propulsion by using the large power generated by flexible 5)Navigation & solar cells on the sail membrane. IKAROS is the precursor Orbit control mission to demonstrate the key technologies for the future 4)Acceleration using solar power sail missions. The main missions of IKAROS are Solar sail. 1) deployment of the large sail membrane in the interplanetary 1)Launched by H‐IIA. 2)Attitude control 3)Sail deployment. space, 2) power generation by the thin film solar cells on the Spin separation. Power generation. before deployment. sail, 3) confirming the acceleration by the solar radiation pressure and 4) demonstration of the navigation and guidance Fig. 1. Nominal operation sequence of IKAROS mission. Sail deployment was achieved on Jun 9th , 2010. of the solar sail type spacecraft. Fig. 1 shows the outline of the

IKAROS mission The successful deployment of the solar sail was achieved on Jun 9th and the acceleration by the solar radiation pressure was confirmed immediately following the sail deployment. And the power generation of the thin file solar cells was also verified. Fig. 2 is the picture captured by the deployable camera (DCAM) on June 14th, 2010. As the IKAROS spacecraft was the secondary payload, the development cost and period were restricted. Consequently, the spacecraft system is simple. Especially, the onboard attitude determination is incomplete. The attitude determination of IKAROS is possible only by collaborating with the ground systems. This paper shows the attitude Fig. 2. IKAROS solar sailing in the interplanetary space. This picture determination and control system of IKAROS. was captured by the deployable camara (DCAM) on June 14th, 2010. As IKAROS is the spin type spacecraft, the spin axis reorientation is not so easy because it requires the large

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Copyright© 2012 by the Japan Society for Aeronautical and Space Sciences and ISTS. All rights reserved.

To_4_1 Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)

2. Attitude Determination and Control and System XMGA SSAS XLGA1 2.1. Attitude determination system of IKAROS IKAROS is a spin-stabilized spacecraft equipped with a large flexible structure. The typical spin type spacecrafts in the deep space are equipped with star sensors for the attitude determination. However, IKAROS does not have star sensors because the development cost and time are strictly restricted. Then the alternative attitude determination method is chosen. The attitude determination of IKAROS is realized by the Sun/Earth angle measurement. The sun angle is measured by a Spin Sun Aspect Sensor (SSAS). This sensor can observe the spin period and the angle between the spin axis and sun direction. The earth angle is measured by Doppler modulation RCS XLGA2 of the downlink RF. Fig. 3 shows the locations of ADCS Fig. 3. Locations of SSAS, antennas(XLGA1, XLGA2, MGA) and RCS. components; the SSAS, RCS and the antennas. IKAROS has Antennas are intentionally located offset from the spin axis. three antennas (XLGA1, XLGA2 and XMGA), and all of them are intentionally located offset from the spin axis. As a result, it can be possible to measure the Doppler shift of the spin motion. Fig. 4 shows the Doppler data of IKAROS. The spin period is detected by analyzing the period of the data. The earth angle can be estimated from the amplitude of the wave. The O-C (Observation – Calculation) value of the two-way Doppler data is written as = Ω+θ VAsin( t ) , (1) Fig. 4. Doppler modulation of downlink RF. Spin period and earth angle can be estimated from this data. where Ω is the spin period and A is the amplitude of the spin modulation. Then the earth angle can be estimated as

−1 Sun pulse Spin rate η = sin (Ar / (2ANT )) , (2) period

Sun sensor where rANT is the distance between the spin axis and the antenna. By combining this earth angle and the sun anble β , Sun angle Sun angle history Spin axis direction 2 the direction of the spin axis can be determined as follows; Spin axis direction 1 Earth angle

1−−xy cosβη cos Integrated Doppler ˆ ˆ ˆ ˆˆ Est. of amplitude A=+ xS yG ± ()SG×. (3) correction ˆˆ2 high-freq. 1(−⋅GS ) component Curve fitting Est. of phase Elimination of arbitrary Doppler Sˆ and Gˆ represent the direction of the sun and the earth data (1Hz) DC Polarization correction LOS Velocity viewed from the spacecraft, respectively. And x and y are component written as cosβ−⋅GSˆˆ cos ηη cos−⋅GSˆˆ cos β Fig. 5. Attitude determination system of IKAROS. Spin axis direction xy==, . (4) can be estimated by combining the date of sun sensor and Doppler data. −⋅ˆˆ22−⋅ˆˆ 1(GS ) 1(GS ) Eq. (3) corresponds to calculating the intersection of two circles on the celestial sphere. The centers of two circles are the sun and the earth, respectively and the radius are β and η . As Eq. (3) shows, two solutions can be found generally. In the IKAROS operations, the true solution can be chosen by using the phase information that can be obtained by handling the time information precisely. Fig. 5 shows the attitude determination system of IKAROS. The sun angle data is extracted from HK data. The Doppler data is given by the ground equipment and the ground-based attitude determination software process the data in quasi real-time. Fig. 6 shows the snapshot of IKAROS attitude determination QL. The spin axis direction is shown on the celestial sphere. Two solutions are found but the true one is chosen automatically by Fig. 6. Snapshot of IKAROS attitude determination QL. The spin axis processing the phase data. direction is shown on the celestial sphere.

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To_4_2 T. SAIKI et al.: Attitude Operation Results of Solar Sail Demonstrator IKAROS

2.2. Attitude control system of IKAROS 3. Long-term Attitude Control Plan of IKAROS The attitude control of IKAROS is done primarily by RCS. IKAROS is equipped with a newly developed Gas-Liquid Here shows the long-term attitude plan of IKAROS. Fig. 9 Phase-Equilibrium Thruster. This is a kind of the cold gas indicates the antenna coverage of IKAROS. XLGA1 is thruster, but the fuel is stored in liquid phase. attached on the top of IKAROS and XLGA2 is on the bottom Hydrofluorocarbon(HFC-134a) is used as fuel. The fuel is in panel. Both antennas have wide beams. But there is the liquid phase in the tank and it is in gas phase at the injection. possibility that the reflected radio wave cause harmful effects As HFC-134a is innoxious and it is not a high pressure to the communication, the antennas are purposely designed system, it is easy to handle and suitable for the secondary not to cover the sail directions (invisible zone). Consequently, payload. As the fuel loses heat when liquid phase is changed the link condition becomes bad when switching the antenna to gas phase, the temperature of the fuel tank and pipes are and the attitude should be controlled to avoid this invisible controlled to keep the phase transition and the control torque zone of antenna. is given intermittently not continuously. And it is important Fig. 10 shows the orbit of IKAROS viewed from the for this propulsion system to separate the gas from the liquid. Sun-Earth fixed frame. As this figure shows, during initial For this purpose, a porous metal is placed in the tank to keep three months after launch, XLGA1 is used for the the liquid in it. IKAROS has eight thrusters for spin up/down communication between IKAROS and ground stations. After and spin axis reorientation. that, the antenna is switched to XLGA2. The attitude control logics of IKAROS are quite simple. The conventional rhumb line control (RLC) and the active Spin axis nutation control (ANC) method are supported. In addition to these conventional logics, IKAROS is equipped with the XLGA1 extended control logics, Flex-RLC/Flex-ANC attitude control ± logics. Fig. 7 shows the outline of the Flex-RCL. The nutation 30deg Sail rate is monitored with the rate gyros attached on the central outside the coverage area body. If the control torque is expected to enlarge the nutation motion, the command is canceled (A). If the control torque is XLGA2 expected to attenuate the nutation motion, the control torque is Fig. 9 Antenna coverage of IKAROS. It is designed not to cover the sail given (B). In reality, as the nutation motion is heavily-damped directions to avoid the RF reflection due to the sail membrane. due to the structural attenuation of the sail, these control logics

were not so effective. 0.8 Earth IKAROS is equipped with another attitude control device, Venus 0.6 IKAROS RCD (Reflectivity Control Device). RCD is a liquid Sun crystal-based variable-optical property sheet. Fuel free attitude 0.4 control is realized because it is driven by electrical power. Fig. 0.2

8 shows the principle of the attitude control with RCDs. 0

-0.2 Y [AU] Fixed (S-E EC) (A) (B) (A) (B) Rhumb line -0.4 control window -0.6

Angular velocity -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 X [AU] (S-E Fixed EC) of nutation Fig. 10. Orbit of IKAROS (Sun-Earth fixed coordinate). It reached the closest point to Venus on Dec. 8th, 2010. Thruster command Fig. 7. Concept of Flex-RLC. The rhumb line control and nutation control are conducted at a same time.

② ⑤

① ③ ⑥ ④

⑦ ⑧

⑨

Fig. 8. Fuel free attitude control with RCD. 72 RCDs are attached at the edge of the membrane. Fig. 11. Long-term attitude control plan of IKAROS. The contour lines show the earth angle. White area is “out of communication area”.

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To_4_3 Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)

Fig. 11 shows the long-term attitude control plan control was performed at the 4th pass and 5th pass and considered before the launch of IKAROS. The horizontal axis in-plane sun angle became -18 degree and out-of-plane sun corresponds to time from the launch and the vertical axis angle became 13 degree. After 5th pass, the spin axis did not corresponds to the in-plane sun angle. And the contour line be controlled. Fig. 14 shows the total sun angle history. As the shows the earth angle. XLGA1 is used between the earth out-of-plane sun angle was controlled to be 13 degree, the angles of 0 to 60 degree and XLGA2 is used when the earth total sun angle during the deployment operation could be kept angle between 120 to 180 degree. White area of this figure is larger than 10 degree which was the requirement from the “out of communication area” where the link condition is bad. mission structure. The skeleton of each phase in the figure is shown below. 1) After separated from the rocket, the deployment Sun angle without operation is conducted. As the sun distance is large, the reorientation sun angle is required to be small to ensure sufficient power generation. The spin axis reorientation is impossible over the deployment operations, the spin axis direction should be controlled properly before that operation. 2) It is required to increase the in-plane sun angle to keep the communication link. This is the phase to demonstrate the solar radiation pressure acceleration. Three or more Ensure the power Constraint on sun generation by angle (derived from coasting passes are necessary for the orbit determination. controlling the the power generation 3) The spin axis direction is controlled to point the sun. The direction of spin axis requirement) antenna is switched to XLGA2 (First out of communication period) 4)-5) Same as phase 2). Fig. 12. Attitude control strategy before deployment. It is difficult to 6) Second out of communication period. The antenna is satisfy the sun angle requirement without attitude reorientation. switched to XLGA1 15 7) Wait until the communication link becomes good. 10

8) The spin axis is controlled and the earth angle becomes [deg]

Deployment less than 30 degree. 5 Plane) ‐ of

9) Communication link with MGA is established. ‐

(Out Attitude Control 0

Angle

4. Attitude Operation Result before Sail Deployment Sun ‐5 Launch The attitude operation result before deployment is shown ‐10 here. The sun angle during the deployment operation is ‐20 ‐15 ‐10 ‐5 0 5 10 Sun Angle (In‐Plane) [deg] required to be more than 10 degree. If the sun angle is less than 10 degree, the deployment structure is in the shade and Fig. 13. Spin axis direction until deployment. The spin axis reorientation its temperature becomes low. It is harmful to the deployment was performed at 4th and 5th pass. of the sail. There is another constraint on the sun angle. As the 25 sun distance is large during the deployment operation, the sun angle is required to be small to generate sufficient electric 20 power. 15 [deg] The attitude of IKAROS is sun-pointing with slow spin at Angle the separation from the rocket. As IKAROS is spin type 10 spacecraft, its sun angle continues to increase with 1deg/day Sun due to its orbital motion. Fig. 12 shows the attitude control 5 strategy before deployment. The dash line shows the 0 constraint on sun angle. It is derived from the requirement of 0 5 10 15 20 25 30 electric power generation for the deployment. As this figure Time [day] shows, it is difficult to satisfy this requirement without the attitude control because the sun angle becomes large. Then the Fig. 14. History of the total sun angle. It could be kept larger than 12 degree. spin axis reorientation control was performed to reduce the in-plane sun angle before the deployment operation, because it is difficult to control the attitude during the deployment phase. 5. Attitude Operation Result after Sail Deployment At the same time, the out-of-plane sun angle was controlled to be larger than 10 degree to warm the mission structure. 5.1. Spin axis drift of IKAROS Fig. 13 shows the spin axis direction with respect to the sun The final step of the deployment sequence (2nd deployment) direction until the sail deployment. The spin axis reorientation was performed on Jun 9th, 2010. The deployment was observed

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To_4_4 T. SAIKI et al.: Attitude Operation Results of Solar Sail Demonstrator IKAROS

by the Doppler measurement first. Fig. 15 shows the Doppler The similar effect was observed in the operation of measurement around the deployment. The slope change of this “HAYABUSA”. The diffusive reflectivity of the solar array data indicates that IKAROS was accelerated. The observed panels (SAP) causes the SRP force along the SAP surface. acceleration was 3.6e-6m/s2, which is almost as expected. And the location of SAP has an offset from the spacecraft The attitude behavior after the sail deployment has been center of gravity. As a result, the SRP torque perpendicular to affected by the SRP. Fig. 16 shows the history of the spin axis the spin axis was given and the circling motion of spin axis direction with respect to the sun direction. As IKAROS was occurred. “HAYABUSA” applied this SRP torque for an ordinary spin-stabilized rigid body before the sail fuel-free sun-tracking when it was operated as a spin type deployment, its spin axis is fixed with respect to inertial space. spacecraft. In this case, the in-plane sun angle is kept constant as Fig. 13 At first, it was considered that the spin axis behavior of shows. However, this situation suddenly changed after the sail IKAROS was same as HAYABUSA, but some points are deployment. The spin axis started circling around a certain different from HAYABUSA. First, the magnitude of estimated point. Additionally, the circling motion of the spin axis SRP torque of IKAROS is much larger than HAYABUSA became larger after the spin down control at 27th and 28th even considering the difference in the light receiving area. pass. The spin rate of IKAROS also has been affected by the Next, the shape of the spin axis trajectory was different. In the SRP. Fig. 17 shows the spin rate history after sail deployment. case of HAYABUSA, the shape was almost true circle, but in The spin rate before the sail deployment was constant. But the the case of IKAROS, the trajectory of spin axis seems spiral. spin rate began to decrease just after the sail deployment. And In addition, the spin rate change was not observed in the rate of spin rate change became larger after the spin down HAYABUSA’s case. After studying on these behaviors, we control. reached the conclusion that the non-uniform surface of the sail and optical property distribution that were very difficult to Second-stage2ᰴዷ㐿ታⴕ(9:36UTC) sail deployment(9:36UTC) control in the design phase affects the SRP torque. Here we consider the 1st order deformation as Fig. 18

Period1way䊄䉾 of䊒䊤 data䊷䊝 䊷loss䊄 dueಾ䉍 ᦧto䈋 deployment䈮䉋䉎䊂䊷䉺 shows as it is very difficult to consider the arbitrary surface operationᰳ៊㗔ၞ and reflective property of the sail. The detail is omitted in this paper, the equation of spin axis motion as written as −φ&&& +Ωθ =−p(cosγφ sin cos(θθ− ) − sinγ sin(θθ − )) ss, (5) &&& θ+Ω φ =−p(sin γ sin φ cos( θθss− ) + cos γ sin( θθ − )) Fig. 15. Doppler date around the sail deployment process. The slope change indicates that IKAROS was solar sailing. where θ and φ are in-plane and out-of-plane phase angle 15 of spin axis respectively, p is the magnitude of SRP, γ is the direction parameter of SRP torque and Ω is the spin rate 10 of spacecraft. θs is the phase angle of the sun, then θ −θs [deg] means the out-of-plane sun angle. The p and γ are 5 Plane) ‐

of nd determined by the magnitude of ξ and η . From the ‐ 2 Deployment (Out 0 Spin Down equation of motion, it is found that the spin axis draws the

Angle (2.5rpm‐>1.1rpm) spiral trajectory. The center of the spiral motion called

Sun ‐5 “equilibrium point” here is written as ΩΩ ‐10 φ =− θ&&cosγ , θθ= − θsinγ . (6) ‐10 ‐5 0 5 10 15 20 pps ss Sun Angle (In‐Plane) [deg] The solution of characteristic equation can be written as Fig. 16. Spin axis direction behavior after sail deployment. The out-of-plane sun angle started to reduce just after the deployment. ppsinγγ cos λ =− ± i . (7) 2.6 ΩΩ 2.4

2.2 This means that the sign of sinγ determine whether the

2 spiral motion is stable or not. In the case of IKAROS, the 2nd Deployment 1.8 spiral motion is stable and the spin axis approaches the Rate[rpm]

1.6 equilibrium point. Spin 1.4 Spin Down 1.2 1 η h 15202530 ξ Time [day]

Fig. 18. First order deformation of sail. Deflection angle and torsion Fig. 17. Spin rate behavior after sail deployment. The spin rate also angle are considered. started to reduce just after the deployment.

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To_4_5 Trans. JSASS Aerospace Tech. Japan Vol. 10, No. ists28 (2012)

5.2. Attitude operation results axis reorientation control was performed. The spin axis The attitude operation results after sail deployment are motion in the phase 3) is shown in Fig. 21. The spin axis shown here. As shown in Fig. 11, in the phase 2), it is required direction approached the equilibrium point near the sun as to increase the in-plane sun angle. Before the launch, it was expected. As a result, IKAROS could save the fuel of the expected that the spin axis was almost fixed in the inertial reorientation. After this phase, we took the full advantage of frame and the in-plane sun angle increased naturally due to the this motion and fuel consumption could be reduced. orbital motion. However, as shown before, the SRP torque was quite large and the spin axis moved dynamically. The 20.00 simplest strategy to keep the spin axis in the inertial frame is 15.00 to control the spin axis actively by using the RCS. However, 10.00 [deg] as the phase 2) was the acceleration demonstration phase, the 5.00 0.00 Plane) long coasting for the precise orbit determination was ‐ necessary. Then the attitude control strategy shown in Fig. 19 of ‐5.00 ‐ (Out was applied. As Eq. (6) indicates, when the spin rate is large, ‐10.00 ‐15.00 Angle

the equilibrium point is far away from the sun. It means that ‐20.00 the spin axis is almost fixed in the inertial space and the Sun in-plane sun angle increase naturally. The spin rate decreases ‐25.00 due to the SRP torque as the Fig. 17 shows without the active ‐30.00 ‐20.00 ‐10.00 0.00 10.00 20.00 30.00 control. Then the spin axis reorientation control is performed. Sun Angle (In‐Plane) [deg] It becomes easier to change the direction of spin axis because the spin rate is small on this moment. The spin axis control Fig. 21. The spin axis trajectory between Sep.11, 2010 and Nov.30 2010. In this period, only the spin rate was maintained by RCS. result in phase 2) is shown in Fig. 20. The spin axis was controlled as Fig. 19. 6. Conclusion

3) Spin axis 1) Increase the in-plane This paper shows the attitude determination and control sun angle by spin-up reorientation and spin up control system of IKAROS. Different from the other spin satellite, IKAROS use the Doppler modulation of the downlink RF for Sun Angle the attitude determination. Although it is very simple method (in-plane) it works very well. And the new Gas-liquid phase-equilibrium 2) Spin rate decreases thruster system works all right in space. And this paper naturally and the out-of-plane sun angle outlines the spin axis motion of IKAROS and the operation Sun Angle(out-of-plane) Sun becomes small results. Although the SRP torque is larger than we expected Fig. 19. The spin axis control strategy in the phase 2). It is possible to and the spin axis change dynamically, we can use this drift ensure the long coasting time. motion to save the fuel. IKAROS is solar sailing even now and it continues to give us important data 20

15 References 10 [deg] 5 1) Tsuda, Y., Mori, O., Funase, R., Sawada, H., Yamamoto, Saiki, T., Plane) ‐ Endo, T. and Kawaguchi, J.: Flight Status of IKAROS Deep Space of ‐ 0 Solar Sail Demonstrator, 61st International Astronautical Congress,

(Out 0 10 20 30 40

‐5 2010, IAC-10-A3.6.8. Angle 2) Tsuda, Y.: An Attitude Control Strategy for Spinning Solar Sail, ‐10 Sun 17th IFAC Symposium on Automatic Control in Aerospace, 2007, WE-P02. ‐15 3) Saiki, T., Nakaya, K., Yamamoto, T., Tsuda, Y., Mori, O. and ‐20 Kawaguchi, J.: Development of a Small-spin-axis Controller and Its Sun Angle (In‐Plane) [deg] Application to a Solar Sail Subpayload Satellite, Transaction of the Fig. 20. The spin axis control result in phase 2). The spin axis was Japan Society for Aeronautical and Spae Sciences, Space controlled as shown in Fig. 19. Technology Japan, 7,(2009), pp.25-32. 4) Funase, R., Mori, O., Tsuda, Y., Shirasawa, Y., Saiki, T., Mimasu, Y. and Kawaguchi, J.: Attitude Control of IKAROS Solar Sail In the phase 3), it is required to control the spin axis to the sun Spacecraft and Its Flight Results, 61st International Astronautical direction. During operation of the phase 2), we could understand Congress, 2010, IAC-10.C1.4.3. the spin axis drift mechanism by observing the spin axis trajectory. 5) Kawaguchi, J. and Shirakawa, K.: A Fuel-Free Sun-Tracking Attitude Control Strategy And the Flight Results in Hayabusa It was expected that the equilibrium point of the spin axis drift (MUSES-C), AAS Flight Mechanics Conference, 2007, motion was located near the sun direction and the spin axis AAS07-176. would approach the equilibrium point. Consequently, in this period, only the spin rate was maintained by RCS. No spin

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