Yuichi Tsuda

Hayabusa2 is the second Japanese small body sample return exploration mission, targeting the carbonaceous asteroid (162173) Ryugu. The spacecraft was launched with Japan's H2A launch vehicle from the Tanegashima Space Center in December 2014 and arrived at Ryugu in June 2018. After completing detailed remote sensing observations, two sampling operations, one kinetic impact experiment, and multiple deployments of robotic smaller probes, the spacecraft left the asteroid in November 2019. It was a six-year journey that the spacecraft traveled approximately 5.2 billion km. In December 2020, the spacecraft returned to the Earth with extraterrestrial materials. A special volume is developed as a primary reference to collect engineering efforts from mission planning through in-orbit operations that made Hayabusa2’s achievements. This chapter introduces a brief overview of this book.

Planetesimals—the initial stage of the planetary formation process—are considered to be initially very porous aggregates of dusts1,2, and subsequent thermal and compaction processes reduce their porosity3. The Hayabusa2 spacecraft found that boulders on the surface of asteroid (162173) Ryugu have an average porosity of 30–50% (refs. 4–6), higher than meteorites but lower than cometary nuclei7, which are considered to be remnants of the original planetesimals8. Here, using high-resolution thermal and optical imaging of Ryugu’s surface, we discovered, on the floor of fresh small craters (<20 m in diameter), boulders with reflectance (~0.015) lower than the Ryugu average6 and porosity >70%, which is as high as in cometary bodies. The artificial crater formed by Hayabusa2’s impact experiment9 is similar to these craters in size but does not have such high-porosity boulders. Thus, we argue that the observed high porosity is intrinsic and not created by subsequent impact comminution and/or cracking. We propose that these boulders are the least processed material on Ryugu and represent remnants of porous planetesimals that did not undergo a high degree of heating and compaction3. Our multi-instrumental analysis suggests that fragments of the highly porous boulders are mixed within the surface regolith globally, implying that they might be captured within collected samples by touch-down operations10,11.

Hayabusa2 is an asteroid sample return mission developed and operated by the Japan Aerospace Exploration Agency (JAXA). Hayabusa2 visited the C-Type asteroid Ryugu in 2018, stayed in the proximity of the asteroid for 1.5 years, and returned to Earth in 2020. Hayabusa2 succeeded in delivering three surface exploration robots to the asteroid surface, performing two landing and sample collection activities, generating one artificial crater impact, and deploying three small objects into orbit around the asteroid. The terrain of Ryugu was found to be unexpectedly harsh through the in-situ observations, and the operation strategy was obliged to be changed and aligned to the Ryugu environment. The project team overcame all the difficulties through tight and collaborative works between the team s scientists and engineers, and completed the planned missions perfectly. The total of 5.4 g of Ryugu sample was confirmed to contain in the returned capsule, which is now being analysed by specifically organized initial analysis teams, and will be delivered to international researchers through AO in 2022. This paper describes the entire flight result of the Hayabusa2 mission, and summarizes the engineering and scientific accomplishments of the mission.

© 2020, Springer Nature B.V. One of the primary goals of Hayabusa2 is to land on the asteroid Ryugu to collect its surface materials. The key for a successful touchdown is to find a promising landing site that meets both scientific and engineering requirements. Due to the limited availability of pre-arrival information about Ryugu, the landing site selection (LSS) must be conducted based on proximity observations over a limited length of time. In addition, Ryugu was discovered to possess an unexpectedly high abundance of boulders with an absence of wide and flat areas, further complicating the LSS. To resolve these problems, we developed a systematic and stepwise LSS process with a focus on the surface topography of Ryugu and the associated touchdown safety. The proposed LSS scheme consists of two phases: Phase-I LSS, a comprehensive survey of potential landing areas at the 100-m scale based on the global mapping of Ryugu, and Phase-II LSS, a narrowing-down process of the candidate landing sites at the 10-m scale using high-resolution images and a local terrain model. To verify the feasibility of a precision landing at the target site, we also investigated the landing dispersion via a Monte Carlo simulation, which incorporates the effect of the irregular surface gravity field. One of the major characteristics of the Hayabusa2 LSS developed in this study is the iterative feedback between LSS analyses on the ground and actual spacecraft operations near the target asteroid. Using the newly developed method, we chose a landing site with a radius of 3 m, and Hayabusa2 successfully conducted its first touchdown on February 21, 2019. This paper reports the methodology and results of the stepwise iterative LSS for the first Hayabusa2 touchdown. The touchdown operation results reconstructed from flight data are also provided, demonstrating the validity of the adopted LSS strategy.

© 2020, Tsinghua University Press. This paper describes the guidance and navigation technique used by Hayabusa2 for the asteroid rendezvous operation to reach Ryugu. The operation results, including the achieved guidance and navigation performance, are also summarized. Multiple assessment and navigation teams worked closely to provide reliable navigation solutions with a short solution delivery cycle. Although the uncertainty of the Ryugu’s ephemeris was considerable before Hayabusa2’s arrival, a combination of radiometric-optical hybrid navigation and a stochastic-constrained optimum guidance method was able to achieve an accuracy of less than 100 m and 1 cm/s, and the arrival was precisely timed.

© 2020 IAA Hayabusa2 arrived at the C-type asteroid Ryugu in June 2018. During one and a half year of the Ryugu-proximity operation, we succeeded in two rovers landing, one lander landing, two spacecraft touchdown/sample collection, one kinetic impact operation and two tiny reflective balls and one rover orbiting. Among the two successful touchdowns, the second one succeeded in collecting subsurface material exposed by the kinetic impact operation. This paper describes the asteroid proximity operation activity of the Hayabusa2 mission, and gives an overview of the achievements done so far. Some important engineering and scientific activities, which have been done in synchronous with the spacecraft operations to tackle with unexpected Ryugu environment, are also described.

© 2020 American Association for the Advancement of Science. All rights reserved. The near-Earth asteroid (162173) Ryugu is thought to be a primitive carbonaceous object that contains hydrated minerals and organic molecules. We report sample collection from Ryugu’s surface by the Hayabusa2 spacecraft on 21 February 2019. Touchdown images and global observations of surface colors are used to investigate the stratigraphy of the surface around the sample location and across Ryugu. Latitudinal color variations suggest the reddening of exposed surface material by solar heating and/or space weathering. Immediately after touchdown, Hayabusa2’s thrusters disturbed dark, fine grains that originate from the redder materials. The stratigraphic relationship between identified craters and the redder material indicates that surface reddening occurred over a short period of time. We suggest that Ryugu previously experienced an orbital excursion near the Sun.

© 2020 American Association for the Advancement of Science. All rights reserved. The Hayabusa2 spacecraft investigated the small asteroid Ryugu, which has a rubble-pile structure. We describe an impact experiment on Ryugu using Hayabusa2's Small Carry-on Impactor. The impact produced an artificial crater with a diameter >10 meters, which has a semicircular shape, an elevated rim, and a central pit. Images of the impact and resulting ejecta were recorded by the Deployable CAMera 3 for >8 minutes, showing the growth of an ejecta curtain (the outer edge of the ejecta) and deposition of ejecta onto the surface. The ejecta curtain was asymmetric and heterogeneous and it never fully detached from the surface. The crater formed in the gravity-dominated regime; in other words, crater growth was limited by gravity not surface strength. We discuss implications for Ryugu's surface age.

Carbonaceous (C-type) asteroids1 are relics of the early Solar System that have preserved primitive materials since their formation approximately 4.6 billion years ago. They are probably analogues of carbonaceous chondrites2,3 and are essential for understanding planetary formation processes. However, their physical properties remain poorly known because carbonaceous chondrite meteoroids tend not to survive entry to Earth's atmosphere. Here we report on global one-rotation thermographic images of the C-type asteroid 162173 Ryugu, taken by the thermal infrared imager (TIR)4 onboard the spacecraft Hayabusa25, indicating that the asteroid's boulders and their surroundings have similar temperatures, with a derived thermal inertia of about 300 J m-2 s-0.5 K-1 (300 tiu). Contrary to predictions that the surface consists of regolith and dense boulders, this low thermal inertia suggests that the boulders are more porous than typical carbonaceous chondrites6 and that their surroundings are covered with porous fragments more than 10 centimetres in diameter. Close-up thermal images confirm the presence of such porous fragments and the flat diurnal temperature profiles suggest a strong surface roughness effect7,8. We also observed in the close-up thermal images boulders that are colder during the day, with thermal inertia exceeding 600 tiu, corresponding to dense boulders similar to typical carbonaceous chondrites6. These results constrain the formation history of Ryugu: the asteroid must be a rubble pile formed from impact fragments of a parent body with microporosity9 of approximately 30 to 50 per cent that experienced a low degree of consolidation. The dense boulders might have originated from the consolidated innermost region or they may have an exogenic origin. This high-porosity asteroid may link cosmic fluffy dust to dense celestial bodies10.

© 2020 The Japan Society for Aeronautical and Space Sciences. Hayabusa2 arrived at the asteroid Ryugu in June 2018, and as of April 2019, the mission succeeded in conducting two rovers landing, one lander landing, one spacecraft touchdown/sample collection and one kinetic impact operation. This paper describes the initial nine months of the asteroid proximity operation activity of the Hayabusa2 mission, and gives an overview of the achievements thus far. Some important engineering and scientific activities conducted synchronously with spacecraft operations in order to complete all planned operations in time against unexpectedly harsh environment of Ryugu are also described.

© 2019, Tsinghua University Press. This paper describes attitude dynamics properties of spinning, momentum-biased and zero-momentum solar sail spacecraft. The model called “Generalized Sail Dynamics Model” (GSDM) is introduced, which can deal with general and practical sail configurations, such as arbitrary optical property distribution, shape and surface wrinkles. Attitude stability criteria and other key dynamical characteristics are derived and compared by compact analytical equations induced from the GSDM. The newly derived zero-momentum sail dynamics is compared with that of spinning and momentum-biased sails. It is shown that the spinning and momentum sails have an advantage in terms of dynamical stability whereas zero-momentum sails are only statically stable. With this special property, angular momentum-stabilized sails can realize a sun-pointing stable attitude with almost zero-fuel, which are discussed with actual space flight experience of the JAXA’s two interplanetary missions, IKAROS and Hayabusa2.

© 2018 IAA The Japan Aerospace Exploration Agency launched the asteroid sample return spacecraft “Hayabusa2” on December 3, 2014. Hayabusa2 will reach the C-type asteroid 162173 Ryugu in 2018, and return back to the Earth in 2020. Sample collections from three sites, four surface rovers deployment and a 4 MJ-class kinetic impact crater forming are planned in the 1.5 years of the asteroid-proximity operation. The mission objective of Hayabusa2 has three aspects, science, engineering and exploration, all of which would be expanded by the successful round-trip journey. The objectives and technologies used in this mission is not a direct solution for the future planetary defense, but should contribute to this field by increasing general asteroid knowledge and enhancing human capabilities of small body-surface access/roving/sampling/impacting. This paper describes the outline of the Hayabusa2 mission, overviews the kinetic impact technology as an example of planetary defense-related technologies and the current flight status after the two and a half years of the interplanetary cruise.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. Hayabusa2 arrived at the C-type asteroid Ryugu in June 2018. During one and a half year of the Ryugu-proximity operation, we succeeded in two rovers landing, one lander landing, two spacecraft touchdown/sample collection, one kinetic impact operation and two tiny reflective balls and one rover orbiting. Among the two successful touchdowns, the second one succeeded in collecting subsurface material exposed by the kinetic impact operation. This paper describes the asteroid proximity operation activity of the Hayabuas2 mission, and gives an overview of the achievements done so far. Some important engineering and scientific activities, which have been done in synchronous with the spacecraft operations to tackle with unexpected Ryugu environment, are also described.

This paper describes a methodology to find almost-optimum trajectories which are robust against inflight stochastic events, such as navigation/guidance error and unexpected missed thrust due to temporal spacecraft malfunctions. A Monte-Carlo based solution search technique was developed which can generate robustness-increased trajectories by deoptimizing the original solution. Arbitrary practical control constraints can be imposed, and one can obtain a solution range in the neighborhood of the original solution which improves the stochastic events-robustness. The technique was applied to an asteroid sample-return mission Hayabusa2 to improve the missed-thrust recoverability, which are presented in detail in this paper.

© 2018 Univelt Inc. All rights reserved. This paper describes a methodology to find almost-optimum trajectories which are robust against inflight stochastic events, such as navigation/guidance error and unexpected missed thrust due to temporal spacecraft malfunctions. A Monte-Carlo based solution search technique was developed which can generate robustness-increased trajectories by deoptimizing the original solution. Arbitrary practical control constraints can be imposed, and one can obtain a solution range in the neighborhood of the original solution which improves the stochastic events-robustness. The technique was applied to an asteroid sample-return mission Hayabusa2 to improve the missed-thrust recoverability, which are presented in detail in this paper.

© 2017 IAA The Japan Aerospace Exploration Agency launched an asteroid sample return spacecraft “Hayabusa2” on December 3, 2014 by the Japanese H2A launch vehicle. Hayabusa2 aims at the round trip mission to the asteroid 162173 Ryugu. Hayabusa2 successfully conducted the Earth gravity assist on December 3, 2015, and now the spacecraft is flying toward Ryugu with the microwave discharge ion engine as the means of propulsion. As of September 2016, 1346 h of the ion engine operation has been achieved as planned. Three touch downs/sample collections, one kinetic impact/crater generation, four surface rovers deployment and many other in-situ observations are planned in the asteroid proximity phase. The operation team will perform extensive operation practice/rehearsal using a hardware-in-the-loop simulator in the year 2017 to be ready for the asteroid arrival in the summer 2018.

© 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. This paper describes the modeling, dynamical characteristics, and implementation of an attitude control method that actively uses solar radiation pressure. The theory behind this control method is called the generalized sail dynamics model, which was developed by the authors and successfully applied to Hayabusa2, which is a Japanese asteroid explorer launched in 2014. The quasi-stable property of the dynamics is proved, which enables the implementation of a fuel-free sun- Tracking attitude using only one reaction wheel. As of August 2016, the attitude of Hayabusa2 was maintained within 10 deg offset from the sun direction for 193 days in total without consuming any fuel. The auto-sun tracking, single-wheel, and fuel-free features were distinctive as compared to any other conventional control methods, such as three- Axis stabilization, and brought many merits to practical spacecraft operations. The theoretical background, the prelaunch evaluation based on a finite element model analysis, the identification process of the dynamics model carried out for the Hayabusa2 mission operation, and their effectiveness are presented in this paper.

© 2016 IAA The Japan Aerospace Exploration Agency launched the asteroid sample return spacecraft “Hayabusa2” on December 3, 2014. Hayabusa2 will reach the C-type asteroid 1999 JU3 in 2018, and return back to the Earth in 2020. Sample collections from three sites, four surface rovers deployment and a 4 MJ-class kinetic impact crater generation are planned in the 1.5 years of the asteroid-proximity operation. The mission objective of Hayabusa2 has three aspects, science, engineering and exploration, all of which would be expanded by the successful round-trip journey. This paper describes the outline of the Hayabusa2 mission and the current flight status after the seven month of the interplanetary cruise.

This paper describes a method of modeling general attitude dynamics of non spinning momentum-biased spacecraft under strong influence of solar radiation pressure (SRP). This model, called "Generalized Sail Dynamics Model", can be applied to realistic sails with non-flat surfaces that have non-uniform optical properties. A coarse Sun-pointing, momentum-biased sail spacecraft is especially focused, for which an approximate solution for the equations of motion is analytically derived. Stability and some other fundamental characteristics of momentum-biased sail spacecraft dynamics, as well as theoretical connections with the past representative sail dynamical models are discussed in detail. Furthermore, the unique behaviors predicted by the model are verified using flight data of the Japanese interplanetary probe Hayabusa2.

This paper describes a method of evaluating sail quality utilizing in-flight attitude behavior of spinning solar sailer IKAROS. Since the successful deployment of the sail, IKAROS has received SRP which strongly affects both translational and rotational motion of the spacecraft. The authors have derived the "Generalized Spinning Sail Model (GSSM)" to reproduce observed unique attitude behavior of IKAROS. Following the previous work, this paper attempts to relate the GSSM with sail quality such as sail shape and flatness. An optical FEM model is constructed to evaluate the precise SRP effect on the spacecraft, and some candidates of deformed sail shape is reproduced which is consistent with the observed attitude motion. We also conclude by the in-flight attitude behavior that the surface roughness of the IKAROS sail is 0.33% at minimum.

This paper describes an attitude determination strategy for spinner spacecraft based on the Sun and the Earth angles. This method realizes a complete spin vector determination using only one sun sensor. Thus this method is suitable for low cost, resource-limited spacecraft with a moderate attitude determination accuracy requirement. The method has been developed for and is actually used in IKAROS, which is a Japanese interplanetary solar sail demonstration mission. This paper introduces theoretical backgrounds of Sun-Earth based attitude determination and shows how the actual implementation was done in the IKAROS mission. Then the attitude determination performance achieved during the actual operation is evaluated.

This paper describes a modeling of attitude dynamics of spinning solar sail spacecraft under influence of solar radiation pressure (SRP). This method is verified and actually exploited in the attitude and trajectory guidance operation of Japanese interplanetary solar sail demonstration spacecraft IKAROS. IKAROS shows a unique attitude behavior due to strong SRP effect. This paper introduces a new generalized dynamics model called Generalized Spinning Sail Model (GSSM), which clearly explains physics behind the observed phenomena with only three parameters. Precise understanding of attitude dynamics through the GSSM led to 6 months of world first interplanetary trajectory guidance of solar sail-craft toward Venus. The GSSM also contributed to realize a zero-fuel spin axis maintenance for most of the flight path before the Venus flyby. In this paper, an overview of IKAROS attitude and trajectory control operation in the cruising phase, derivation of the proposed model and its implementation to the actual IKAROS operation are shown.

This paper presents a method for approximating the state transition matrix for orbits around a primary body and subject to arbitrary perturbations. The primary objective of this method is to provide an accurate state transition matrix for orbits with realistic perturbations, which has a sufficiently simple form for implementation onboard spacecraft. The averaging method is employed to isolate the high and low frequency spectra of the perturbation terms, and construct a functional form of the approximate state transition matrix composed only of elementary analytic functions. In addition to the methodology of the approximation, it is shown that the symplectic property, which is a fundamental mathematical structure of Hamiltonian systems, can be incorporated into this method. This not only reduces the number of parameters required for approximations but also makes it possible to preserve the physically true structure of the state transition matrix. The resulting state transition matrix approximation is valid for tens of orbital revolutions without having to update the parameters. Numerical simulations show that this method is valid for arbitrary eccentricity orbits with semimajor axis ranging from LEO up to around 10 Earth radii when applied to Earth orbits. © 2010 Elsevier Ltd.

This paper describes a method of modeling attitude dynamics of spinning solar sail spacecraft under influence of solar radiation pressure (SRP). This method is verified and actually exploited in the operation of Japanese interplanetary solar sail demonstration spacecraft IKAROS. IKAROS shows a unique attitude behavior due to strong SRP effect. This paper shows a new attitude model of spinning sail, which is verified by flight data of IKAROS. It is also shown that the model proposed in this paper has a direct relation with the Generalized Sail Model.

IKAROS is the Japanese deep-space solar sail technology demonstration mission launched in 2010. IKAROS is a spinner spacecraft with the 20m-span solar sail kept extended by the centrifugal force. During its solar sailing flight from Earth to Venus, IKAROS showed a unique attitude behavior due to solar radiation pressure attracted on the sail. This paper proposes a generalized model of spinning sail-craft, which clearly explains the attitude behavior observed in IKAROS. Then it is shown that this behavior has a clear dependency on the sail shape and the optical property distributions on the sail. We show the estimation results of the on-orbit sail shape using this new model. Copyright © 2011 by Yuichi Tsuda.

JAXA launched the world's first deep space solar sail demonstration spacecraft "IKAROS" on May 21, 2010. IKAROS was injected to an Earth-Venus trajectory to demonstrate several key technologies for solar sail utilizing the deep space flight environment. IKAROS succeeded in deploying a 20 m-span solar sail on June 9, and is now flying towards the Venus with the assist of solar photon acceleration. This paper describes the mission design, system design, solar sail deployment operation and current flight status of IKAROS. © 2011 Elsevier Ltd. All rights reserved.

This paper describes a method of modeling attitude dynamics of spinning solar sail spacecraft under influence of solar radiation pressure (SRP). This method is verified and actually exploited in the operation of Japanese interplanetary solar sail demonstration spacecraft IKAROS. MAROS shows a unique attitude behavior due to strong SRP effect. This paper shows a new attitude model of spinning sail, which is verified by flight data of MAROS. It is also shown that the model proposed in this paper has a direct relation with the Generalized Sail Model.

This paper presents a method for approximating the state transition matrix for orbits around a primary body and subject to arbitrary perturbations. A generalized averaging method is employed to isolate the high and low frequency regions of the perturbation terms, and construct a functional form of the approximate state transition matrix composed only of elementary analytic functions. The resulting state transition matrix is expressed with a small number of constant parameter matrices and osculating orbit parameters at an initial epoch, and is valid for tens of orbital revolutions without having to update the parameters. Numerical simulations show that this method is valid for arbitrary eccentricity orbits with semimajor axis ranging from LEO up to around 10 Earth radii when applied to Earth orbits. This method has been developed for implementation onboard spacecraft for high accuracy formation flying missions. Furthermore, it is shown that the symplectic property, which is a fundamental mathematical structure of Hamiltonian systems, can be incorporated into the method. This not only reduces the number of parameters required for approximations, but also preserves the physically true structure of the state transition matrix and provides some important propertics that are practical and useful for onboard computation.

This paper derives a new batch sequential estimator utilizing "Unscented Transformation." It is a natural extension of the Unscented Batch Filter(UBF), and it inherits advantages of the UBF, such as being able to deal with nonlinear system equations directly, and providing a higher convergence capability starting from poorer initial guesses compared with conventional Bayesian filters. This paper applies the proposed filter to a relative orbit determination for a multiple spacecraft formation flight mission using relative range measurement information. A major challenge of this problem is that, since the available initial guess of estimation and the dynamic range of the relative orbital motion itself are of the same order, relatively higher convergence capability is required for the estimator to realize an accurate guess. The Bayesian filter, which is generally applied to this kind problem, is found to have an insufficient convergence performance. Numerical simulations show that the proposed "Unscented Batch Sequential Filter" provides a better convergence performance without taking longer computational time than other conventional filters.

JAXA launched the world's first deep-space solar sail demonstration spacecraft "IKAROS" on May 21, 2010. IKAROS was injected to an Earth-Venus trajectory to demonstrate several key technologies for solar sail utilizing the deep space flight environment. IKAROS succeeded in deploying a 20m-span solar sail on June 9, and is now flying toward Venus with the assist of solar photon acceleration. This paper describes the mission design, system design, solar sail deployment operation and current flight status of IKAROS. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

Japan Aerospace Exploration Agency (JAXA) has developed the small demonstration solar sail spacecraft IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun), which will be launched in mid 2010. The main objective of this spacecraft is to deploy the 20m class sail membrane, and demonstrate the acceleration of a spacecraft by the solar radiation pressure (SRP) by means of that sail. It is important to model the SRP force adequately for the objective of navigation, especially for interplanetary spacecraft. In order to improve the model of the SRP torque induced by the sail membrane, the IKAROS project team plans to estimate the SRP torque parameters in orbit. In this paper, we present the approach to obtain the parameters needed for constructing the photon torque model through the analysis of the attitude dynamics.

This paper presents a numerical method for deriving a symplectic state transition matrix for high-fidelity Earth orbits subject to non-dissipative perturbation forces. By taking advantage of properties of Hamiltonian systems, this method provides an exact solution space mapping of linearized orbital dynamics, preserving the symplectic structure that all Hamiltonian systems should possess by nature. This method can be applied to accurate, yet computationally efficient dynamic filters, long-term propagations of the motions of formation flying spacecraft and the eigenstructure analysis of N-body dynamics, etc., when the exact structure-preserving property is crucial. We show the derivation of the numerical method of symplectic state transition matrix, and apply it to Earth orbits with perturbation forces based on real ephemerides. These numerical examples reveal that this method shows improvements in preserving the structural properties of the state transition matrix, and in the computational efficiency compared to the conventional linear state transition matrix with Euler or Runge-Kutta integration. © 2010 The Japan Society for Aeronautical and Space Sciences.

This paper is also submitted to AIAA Journal of Guidance, Control, and Dynamics. This paper presents a method for approximating the state transition matrix for orbits around a primary body and subject to arbitrary perturbations. A generalized averaging method is employed to isolate the high and low frequency regions of the perturbation terms, and construct a functional form of the approximate state transition matrix composed only of elementary analytic functions. The resulting state transition matrix is expressed with a small number of constant parameter matrices and osculating orbit parameters at an initial epoch, and is valid for tens of orbital revolutions without having to update the parameters. Numerical simulations show that this method is valid for arbitrary eccentricity orbits with semimajor axis ranging from LEO up to around 10 Earth radii when applied to Earth orbits. This method has been developed for implementation onboard spacecraft for high accuracy formation flying missions. Furthermore, it is shown that the symplectic property, which is a fundamental mathematical structure of Hamiltonian systems, can be incorporated into the method. This not only reduces the number of parameters required for approximations, but also preserves the physically true structure of the state transition matrix and provides some important properties that are practical and useful for onboard computation.

This paper describes a high precision computation method of deriving general symplectic state transition matrices. Respecting the symplectic structure of state transition matrices is important in such a field as accurate, yet computationally efficient dynamic filters, longterm propagations of the motions of formation flying spacecraft and the eigenstructure/manifold analysis of N-body dynamics etc. The method proposed in this paper is a post-processing to improve or add a symplectic property to arbitrary state transition matrices. Since this method can be combined with any state transition matrix derivation schemes, it works as a truncation error reduction and also provides a simple way to add the symplectic property to matrices derived with non-symplectic methods. We present the derivation, its applicability and numerical evaluations when applied to a twobody dynamics, an Earth orbit with perturbation forces based on the real ephemeris and a circular restricted three-body problem. © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Japan Aerospace Exploration Agency (JAXA) has developed the small demonstration solar sail spacecraft IKAROS (Interplanetary Kite-craft Accelerated by Radiation of the Sun), which will be launched in mid 2010. The main objective of this spacecraft is to deploy the 20m class sail membrane, and demonstrate the acceleration of a spacecraft by the solar radiation pressure (SRP) by means of that sail. It is important to model the SRP force adequately for the objective of navigation, especially for interplanetary spacecraft. In order to improve the model of the SRP torque induced by the sail membrane, the MAROS project team plans to estimate the SRP torque parameters in orbit. In this paper, we present the approach to obtain the parameters needed for constructing the photon torque model through the analysis of the attitude dynamics.

This paper presents a numerical method for deriving a symplectic state transition matrix for an arbitrary Hamiltonian dynamical system. It provides the exact solution-space mapping of the linearized Hamiltonian systems, preserving the symplectic structure that all Hamiltonian systems should possess by nature. The symplectic state transition matrix can be applied to accurate, yet computationally efficient, dynamic filters, long-term propagations of the motions of formation-flying spacecraft, eigenstructure/manifold analysis of N-body dynamics, etc., when the exact structure-preserving property is crucial. We present the derivation and key characteristics of the symplectic state transition matrix and apply it to the two-body dynamics, the circular restricted three-body problem, and an Earth orbit with perturbation forces based on the real ephemeris. These numerical examples reveal that this numerical symplectic state transition matrix shows improvements in preserving the structural properties of the state transition matrix as compared with the conventional linear state transition matrix with Euler or Runge-Kutta integrations. © 2009 by Yuichi Tsuda and Daniel J. Scheeres.

This paper presents a method for approximating the state transition matrix for orbits around a primary body and subject to arbitrary perturbations. A generalized averaging method is employed to isolate the high- and lowfrequency regions of the perturbation terms and to construct a functional form of the approximate state transition matrix composed only of elementary analytic functions. The resulting state transition matrix is expressed with a small number of constant parameter matrices and osculating orbit parameters at an initial epoch and is valid for tens of orbital revolutions without having to update the parameters. Numerical simulations show that this method is valid for arbitrary-eccentricity orbits with semimajor axes ranging from low Earth orbit up to around 10 Earth radii when applied to Earth orbits. This method has been developed for implementation onboard spacecraft for high-accuracy formation-flying missions. Furthermore, it is shown that the symplectic property, which is a fundamental mathematical structure of Hamiltonian systems, can be incorporated into the method. This not only reduces the number of parameters required for approximations, but also preserves the physically true structure of the state transition matrix and provides some important properties that are useful for practical onboard computation. © 2009 by Yuichi Tsuda and Daniel J. Scheeres.

This paper presents a numerical method for deriving a symplectic state transition matrix for an arbitrary Hamiltonian dynamical system. It provides the exact solution space mapping of the linearized Hamiltonian systems, preserving the symplectic structure that all Hamiltonian systems should possess by nature. The symplectic state transition matrix can be applied to accurate, yet computationally efficient dynamic filters, long-term propagations of the motions of formation flying spacecraft and the eigenstructure/manifold analysis of N-body dynamics etc., when the exact structure-preserving property is crucial. We present the derivation and key characteristics of the symplectic state transition matrix, and apply it to the two-body dynamics, circular restricted three-body problem and to an Earth orbit with perturbation forces based on the real ephemeris. These numerical examples reveal that this numerical symplectic state transition matrix shows improvements in preserving the structural properties of the state transition matrix as compared with the conventional linear state transition matrix with Euler or Runge-Kutta integrations.

This paper presents a method for approximating the state transition matrix for orbits around a primary body and subject to arbitrary perturbations. Primary objective of this method is to provide an accurate state transition matrix for orbits with realistic perturbations which has a sufficiently simple form for implementation onboard spacecraft. The averaging method is employed to isolate the high and low frequency spectrum of the perturbation terms, and construct a functional form of the approximate state transition matrix composed only of elementary analytic functions. In addition to the methodology of the approximation, it is shown that the symplectic property, which is a fundamental mathematical structure of Hamiltonian systems, can be incorporated into this method. This not only reduces the number of parameters required for approximations, but also makes it possible to preserve the physically true structure of the state transition matrix. The resulting state transition matrix approximation is valid for tens of orbital revolutions without having to update the parameters. Numerical simulations show that this method is valid for arbitrary eccentricity orbits with semimajor axis ranging from LEO up to around 10 Earth radii when applied to Earth orbits.

The stability of the dynamics of the spacecraft equipped with large flexible membrane is discussed. The dynamic stability of the object having a flexible structure is deteriorated from that of the rigid object in general. This paper provides a fully analytical solution for the stability criteria of the axisymmetric spinner spacecraft, equipped with a large flexible membrane. These criteria can be applied to the design of such missions as spinner solar-sail spacecraft, large solar-power satellite and other general spacecrafts having large axisymmetric structures. © 2008 by Yuichi Tsuda.

Japan Aerospace Exploration Agency (JAXA) is currently studying on the " Solar Sail " propulsion for future deep space explorations. JAXA has conducted extensive studies on spinner solar sail spacecraft system, utilizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize the 7.5μm-thickness and 50m diameter polyimide membrane, combined with thin flexible solar cells, as the photon acceptance surface. Based on the simulator development done through several space and ground experiments on solar sail dynamics, this paper discusses the attitude control of this spinning solar sail system via conventional RCS controller, and also focusing on the utilization of reflectivity control devices attached on the sail, to control the spinning axis and rate of the solar sailer.

Japan Aerospace Exploration Agency (JAXA) is currently studying "Solar Sail" populsion for future deep space explorations. One of the key technologies to realize a solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utlizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize a 7.5μm-thickness and 50m diameter polyimide membrane, combined with a thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In August 9, 2004, JAXA launched the S-310 sounding rocket, which tested two different shapes of membranes during the zero-gravity flight. The first type of membrane looked like a "clover-leaf", and another is like a "fan". These two membranes, both of them having 10m diameter, were unfolded sequentially during the zero-gravity flight under the free spin condition, and their behavior was observed by onboard cameras. This paper focuses on the "clover-leaf" solar sail, which was fully deployed successfully, and introduces the S-310-34 experiments, and then shows the flight results and postflight evaluations. © 2006 Lister Science. All rights reserved.

JAXA is currently planning the next generation magnetosphere observation mission called "SCOPE"(cross-Scale Coupling in Plasma universE). SCOPE aims at observing the Earth's magnetotail with 5 satellites flying in formation to fully resolve the temporary and spatial distribution of the magnetospheric phenomena. For this observation, the clock synchronization and relative distance measurement between the spacecrafts are essential. This paper describes an onboard relative ranging and clock synchronization algorithm, which applies a simplified formulation, using two-way and three-way phase differences as the filter inputs to construct the onboard system suit to the SCOPE mission.

Japan Aerospace Exploration Agency (JAXA) is currently studying on the "Solar Sail" propulsion for future deep space explorations. One of the key technologies to realize the solar sail is how light and how compact we can make the photon acceptance surface. JAXA has conducted extensive studies on utilizing centrifugal force to deploy the photon acceptance surface. The final objective is to realize the 7.5μmthickness and 50m diameter polyimide membrane, combined with thin flexible solar cells, as the photon acceptance surface that will be needed around the Jupiter orbit. In the August 9, 2004, JAXA has launched the S-310 sounding rocket, which tested two different shapes of membranes during the zerogravity flight. The first type of the membrane looks like a "clover-leaf", and another is like a "fan". These two membranes, both of them have 10m diameter, were unfolded sequentially during the zerogravity flight under the free spin condition, and their behavior was observed by onboard cameras. This paper focuses on the "clover-leaf" solar sail, which was fully deployed successfully, and introduces the S-310-34 experiments, and then shows the flight results and postflight evaluations.

On a NASDA's microsatellite named "μ-LABSAT," which was launched by H-IIA on December 14, 2002 (Fig.1), Communications Research Laboratory (CRL), National Aerospace Laboratory (NAL) and University of Tokyo (UT) have been jointly performing several orbital experiments as technology demonstration towards the future orbital servicing missions. In University of Tokyo's experiment which was held on May 14, 2003, the micro-satellite released a small object named "target," and its rotational motion was estimated by the images captured continually using a camera developed by CRL. Then satellite attitude control was performed by visual feedbacks of the target image position on the camera frame so that the target image may come to a certain point on the camera frame. This is a pre-experiment of so-called LOS (Line Of Sight) control, which will be indispensable during rendezvous and docking to the satellite to be serviced. In this paper, the objectives and procedure of these experiments, and the results will be described. Copyright © 2003 by the International Astronautical Federation. All rights reserved.