山本 高行

© 2019 IAA This paper investigates the launch capability of the SS-520 as a CubeSat launch vehicle. The SS-520 was developed by JAXA originally as a two-stage, spin-stabilized, solid-propellant sounding rocket. With less than 2.6 tons in total mass and 10 m in length, the SS-520-5 successfully launched a single 3U-sized CubeSat into orbit on February 3, 2018. The SS-520-5 obtained its capability as a CubeSat launch vehicle by installing a 3rd stage solid motor in addition to the RCS between the 1st and 2nd stages. However, its launch capability was limited due to its rocket system configuration. In order to pursue the SS-520's launch capability, two effective modifications from the SS-520-5 are proposed: thrust enhancement of the 1st stage motor and installation of an additional RCS between the 2nd and 3rd stages. The framework of launch capability analysis is established by a multi-objective genetic algorithm (MOGA), where its two objectives are selected as the altitudes of perigee and apogee. The analysis reveals that the two proposed modifications to the SS-520-5 work effectively but differently. The 10% increase of the 1st stage enhancement is particularly effective when the target altitude of perigee is low (e.g., 200 km), whereas the installment of the additional RCS with 30 kg increases accessibility to a much higher altitude of perigee, even to circular orbit reaching altitudes of 550 km for a 1U-sized CubeSat and 280 km for a 6U-sized CubeSat. The simultaneous application of both modifications would result in launch capability able to deliver a 10-kg payload. From a more general perspective, the results in this paper suggest that it is possible for a very small launch vehicle (VSLV) of the 3-ton class and 10 m in length to deliver a 10-kg-class payload into low Earth orbit.

© 2020, Univelt Inc. All rights reserved. ISAS/JAXA has successfully launched the micro-satellite “TRICOM-1R” by the world’s smallest orbit rocket “SS-520 No.5” from Uchinoura Space Center on February 3rd in 2018. ISAS modified the existing sounding rocket SS-520 adding a small 3rd-stage solid-motor and the attitude control system. It flies spinning for the attitude stabilization in the flight. Therefore, we devised the rhumb-line control system with a new scheme. This rhumb-line system has the high-performance functions; the high-preciseness, the high-maneuver rate and the suppression of the unnecessary nutation angle generated at the RCS injection. This paper reports the development of the G&C system and the flight results.

<p>本稿では，観測ロケットを機体のベースとする超小型衛星打上げ機（SS-520 5号機）で実施した飛行安全について概説する．この機体は超小型であるため，搭載や重量における制約条件を受けたり通常の観測ロケットで用いている既存の地上設備を利用する上での制約条件を受けたりする中での飛行安全運用となった．そのため，本打上げ機は我が国の基幹ロケットに適用されている飛行安全基準を遵守しつつ，長年観測ロケットで培った飛行安全手法を最大限活用し，本打上げ機特有の制約条件を満足しつつ独自の飛行安全運用方法を適用し確実な飛行安全を行った．また，内之浦での軌道投入型ロケットの飛行安全運用はM-Vロケット以来となったため，新たな飛行安全管制システムが必要となった．今回新たに導入した飛行安全管制システムやシステム検証方法および実際のフライトにおいて新システムを適用した結果についても紹介する．</p>

<p>本稿では，観測ロケットベースの超小型衛星打上げ機による地球周回楕円軌道への軌道投入について，その飛行計画について概説する．本打上げ機による目標軌道は，遠地点高度約1,800 km，近地点高度約180 kmであり，近地点高度が低いために期待される軌道寿命は短い．飛行計画に対するミッション要求の一つとして，軌道寿命30日以上の軌道に衛星を投入することが挙げられる．機体誤差源や飛行環境の誤差が達成される軌道に対して大きく影響するため，これらの誤差が十分に小さくなるように管理しなければならない．ここでは観測ロケットをベースにして，どのように超小型衛星打上げ機としての要求を満足する軌道計画を立案したか，およびノミナル軌道に対する飛行分散や飛行安全に対する解析結果を示す．また飛行結果およびポストフライト解析を示し，将来的な能力向上の一案を紹介する．</p>

<p>本解説では，超小型衛星打上げ機（SS-520 4，5号機）の機体システム開発の概要を示す．本ロケットの開発意義は，搭載した宇宙用機器に品質の高い民生部品を活用して超小型衛星打上げシステムを作り上げたことと，従来の開発手法に加え新たに取り組んだ民生品の品質保証の考え方を構築してフライト実証したことである．また，既存の観測ロケットに衛星打上げ能力を持たせるためには，いくつかの課題を克服する必要があった．抜本的な構造軽量化，搭載機器の小型軽量化，衛星とロケット一体となった機能の最適配分，誘導制御系の工夫，飛行安全，Test as Flyをベースとした検証試験等々，限られたリソースと開発期間の厳しい制約条件のなかで随所に創意工夫を施した．本解説では，その開発におけるポイントを総括した．</p>

© 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Solar electric propulsion is a key enabling technology that has improved the efficiency of space transport. With specific impulses that are typically ten times higher than the chemical counterpart, electric motors allow a considerable saving in propellant mass at the expense of longer times of flight. However, the length of the transfer process and the specific operational needs require to develop a different operational concept for the navigation and orbit control that can be sustained during the different phases of the mission. In this paper, a trade-off is performed among several operational concepts and solutions for multi-revolutions SEP transfers with application to the DESTINY+ mission. The GTO-to-Moon low-thrust transfer is first computed in a high-fidelity model with a tangential thrust strategy and later optimized with a five-order Legendre-Gauss-Lobatto collocation method. The impact of eclipses, radiation, thrust outages and misfires, and orbit tracking is analyzed in detailed and included in the transcript optimal problem as algebraic constraints where possible. Numerical results show that the driving factors for the optimal trajectory are related to the operations of the spacecraft rather than the final mass or time of flight.

On February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world's smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don't have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage bum-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.

Copyright © 2019 by the International Astronautical Federation (IAF). All rights reserved. DESTINY+ (Demonstration and Experiment of Space Technology for INterplanetary voYage, Phaethon fLyby and dUSt analysis) is a small-sized high-performance deep space vehicle proposed at ISAS/JAXA. The trajectory design of DESTINY+ is divided into several phases. First phase is an orbit injection into an extended elliptical orbit launched by the Epsilon rocket with the additional solid kick motor. Second phase is many revolutions transfer to raise apogee altitude by low thrust propulsion system to the moon orbit nearby. And at third phase, the distant flyby and the swing-by around the moon is designed to give DESTINY+ momentum to escape Earth gravitational field. At an interplanetary phase, DESTINY+ goes to an Asteroid Phaethon for flyby observation. After the Phaethon flyby, DESTINY+ is planned to go back toward Earth for gravity assist and go to another asteroid 2005UD which thought to have split from Phaethon. This paper discusses DESTINY+'s low-thrust trajectory design. As for the many revolution transfer phase, the low-thrust trajectory is optimized by the multi-objective optimization using genetic algorithm. In this phase, we minimize the time of flight, the passage of time of radiation belt, the work time of low thrust propulsion system and the maximum eclipse period. After the spacecraft reaches to the moon's orbit, it utilizes the moon swing-by several times to connect to the transfer trajectory for Asteroid Phaethon. From these studies, we can show the feasibility of the mission design of DESTINY+,.

© 2017, American Astronautical Society. This work explores the target selection and trajectory design of the mission candidate for ISAS/JAXA’s small science satellite series, DESTINY PLUS or DESTINY+. This mission combines unique aspects of the latest satellite technology and exploration of transition bodies to fill a technical and scientific gap in the Japanese space science program. The spacecraft is targeted to study the comet-asteroid transition body (3200) Phaethon through a combination of low-thrust propulsion and Earth Gravity Assist. The trajectory design concept is presented in details together with the launch window and flyby date analysis. Alternative targets for a possible mission extension scenario are also explored.

<p>In this paper, anomaly detection that is configured as a combination of state observer and Mahalanobis-Taguchi (MT) method is proposed for real time fault detection of rapid and dynamic phenomena such as rocket engine operation. Real time anomaly detecting is recognized as one of the most important elements to realize advanced reusable space transportation system. Conventionally, bottom-up type anomaly detecting logic based on failure mode and effect analysis (FMEA) is usually used for this purpose, however, it requires large amount of time and labor. The proposed method can improve this process. In the present method, error values between calculated ones through rocket engine simulator constructed on autoregressive moving average model and extended Kalman filter (EKF) and measured ones are standardized with existing normal operation data of the rocket engine so as to compute Mahalanobis' distance, which expresses degree of anomaly. We performed engine hot firing tests in simulated anomaly conditions. The obtained data was processed with the present method, and the simulated anomaly in the tests was detected as expected.</p>

© 2016 Institute of Electrical and Electronics Engineers Inc.. All rights reserved. JAXA has successfully launched the SS-520 No.5 with micro-satellite 'TASUKI' on February 3rd 2018 at Kagoshima Space Center at Uchinoura in Japan. The base-line of the SS-520 sounding rocket is a two-stage rocket which has a capability for launching an 80kg payload to a maximum altitude of about 1000 km. and spun by 4 tails for attitude stability. Enhanced SS-520 No.5 is a three-stage rocket for the smallest-class launch system in the world, which has the orbit injection capability of a micro-satellite of a few kilograms by adding a high-performance third solid motor and advanced rhumb-line control system. Total length of the rocket is about 9.6 meters, the gross weight is 2.6 metric tons, and the reference diameter is 0.52 meters. The `TASUKI` has some experimental purposes for 'store & forward' communication on orbit and earth observation by some commercial cameras and others. The key points of this launch was to newly develop the rhumb-line control system, compact and high performance avionics, some lightweight structures, and the third motor made of CFRP. The rhumb-line control system established an attitude maneuver of about 70 degrees to inject the 'TASUKI' into the orbit of perigee altitude 180km and apogee altitude 2000km. This rhumb-line control system has some high performance functions. It has an angular momentum control function with high attitude maneuver rate, and the suppression function of nutation angle generated by the disturbance of RCS thruster injection during high spin rate of about 1.6Hz. We performed a Motion Table (M/T) Test 'Real-time Simulation Test' with flight models of the avionics for verification of the rhumb-line control design and the soft-wear in the loop test for verification of the flight soft-wear. An active nutation control (ANC) function is also equipped for the reduction of the residual nutation angle after the rhumb-line control. We show the outline of the rocket system and developments, especially the rhumb-line control system with the compact avionics system. Finally flight results are showed and we show one of the future enhanced ideas of SS-520 No.5 type launcher for 10 kilograms class satellite.

© 2018 Univelt Inc. All rights reserved. On February 3, 2018 at the JAXA Uchinoura Space Center, JAXA experimented SS-520 No. 5 launch with a 3U sized cube sat called TRICOM-1R aboard. After liftoff, flight of SS-520 No. 5 proceeded normally. Around 7 minutes 30 seconds into flight, TRICOM-1R separated and was inserted into its target orbit. And the launcher became the world’s smallest class satellite launcher. SS-520 launch vehicle is one of sounding rockets operated in JAXA/ISAS, and originally two-stage rocket. In this experiment, to make this vehicle put a satellite into orbit, the third stage motor is added. And this sounding rocket has four tail fins for spin stabilization, but usually don’t have an attitude control system during the flight. But in this mission, it is needed to control its attitude to ignite second and third motor toward horizontal after first stage burn-out. The gas jet system is installed into between the first stage and the second stage of the vehicle as a unique active attitude control system. The gas jet system can control the spin axis direction and the spin rate of the vehicle during the coasting fight. Because of this constraint, the apogee altitude after the burn out of the first stage motor almost correspond with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket-based Nano launcher is planned to put TRICOM-1R into the elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. Because the perigee altitude is relatively low, the orbit life is very short. One of the mission requirements is to make the vehicle an orbit insertion with more than 30 days orbital lifetime. The vehicle error or the environment error deeply affect the achieved trajectory. These errors must be small enough to put TRICOM-1R into orbit. This paper discusses about the trajectory design on how to manage the sounding rocket into a satellite launching vehicle, the effect of the orbital distribution depending on the various errors, the flight safety analysis, and finally flight performance evaluation.

© 2018 Univelt Inc. All rights reserved. The Epsilon launch vehicle, the newest version of Japan’s solid propulsion rocket, made its maiden flight in September of 2013. The purpose of the Epsilon launch vehicle is to provide small satellites with responsive launching with low-cost, user-friendly and efficient launch system. The first flight was successfully finished, JAXA has been conducting intensive researches on a more powerful and lower cost version of Epsilon. In order to minimize technical risks and to keep up with demand of future payloads, JAXA plans to take a step-by-step approach toward Future Launch System. As the first upgrade toward Future Launch System, JAXA has started the development of the Enhanced Epsilon. This development is mainly the renewal of the second stage, and also includes each subsystem’s improvement. This paper describes the development and flight result of the Enhanced Epsilon’s Guidance and Control System.

Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. This paper investigates the launch capability of the SS-520 as a CubeSat launch vehicle. The SS-520 was developed by JAXA originally as a two-stage, spin-stabilized, solid-propellant sounding rocket. With less than 2.6 tons in total mass and 10 meters in length, the SS-520-5 successfully launched a single 3U-sized CubeSat into orbit on February 3, 2018. The SS-520-5 obtained its capability as a CubeSat launch vehicle by installing a 3 rd stage solid motor in addition to the RCS between the 1st and 2nd stages. However, its launch capability was limited (in target altitudes of perigee and apogee at 180 km and 1800 km, respectively) due to its rocket system configuration. In order to pursue the SS-520's launch capability, two effective modifications from the SS-520-5 are proposed: thrust enhancement of the 1st stage motor and installation of an additional RCS between the 2nd and 3rd stages. Furthermore, the framework of launch capability analysis is established by a multi-objective genetic algorithm (MOGA), where its two objectives are selected as the altitudes of perigee and apogee. The problem maintains its simplicity through the selection of only eight design variables of the six acceleration directions and two coasting durations. The analysis reveals that the two proposed modifications to the SS-520-5 work effectively but differently. The 10% increase of the 1st stage enhancement is particularly effective when the target altitude of perigee is low (e.g., 200 km), whereas the installment of the additional RCS with 30 kg increases accessibility to a much higher altitude of perigee, even to circular orbit reaching altitudes of 550 km for a 1U-sized CubeSat and 280 km for a 6U-sized CubeSat. Each modified configuration with the 1st stage enhancement and additional RCS installment enables carrying a payload about twice as heavy as that of the SS-520-5. The application of both modifications would result in launch capability able to deliver a 10-kg payload. From a more general perspective, the results in this paper suggest that it is possible for a very small launch vehicle of the 3-ton class and 10 meters in length to deliver a 10-kg-class payload into low Earth orbit.

Copyright © 2017 by the International Astronautical Federation (IAF). All rights reserved. On January 2017, JAXA/ISAS launched the Nano launcher based on the sounding rocket. Unfortunately, the rocket was not able to put a satellite into orbit. This paper discusses the trajectory design regarding how to manage the sounding rocket as a satellite-launching vehicle. At JAXA/ISAS, there are three types of sounding rockets. Two are single-stage rockets called S-310 and S-520, and one is a two-stage rocket called SS-520. These sounding rockets have tail fins for spin stabilization, but usually lack an attitude control system during flight. When attitude control is required to achieve the mission requirements, a gas jet system is available as an optional device. The gas jet system can control the vehicle's spin axis direction and spin rate during coasting fight. To enable the sounding rocket to put a satellite into orbit, a third-stage motor is added to the SS-520 two-stage sounding rocket. The gas jet system is a unique and active attitude control system installed between the first and second stages of the vehicle. Given this constraint, the apogee altitude after burnout of the first-stage motor almost corresponds with the perigee altitude of the elliptical orbit. In this mission, the sounding rocket based Nano launcher is planned to put a 3U-size CubeSat into elliptical orbit. Its targeted apogee altitude is about 1,800 km and its perigee altitude is about 180 km. As the perigee altitude is a relatively low altitude, the orbit life is very short. Thus, any vehicle or environment errors significantly affect the achieved trajectory. Such errors must be small enough to put a CubeSat into orbit. This paper also discusses the effect of orbital distribution depending on various errors.

The first Epsilon rocket was launched successfully with a small payload 'HISAKI' on September 14th, 2013 in Japan. Epsilon has a new absorber structure in Payload Attach Fitting to reduce the vibration condition for payload. We designed the robust control logic to satisfy the compatibility of robust stability and response against various disturbances. The 3rd Stage under spinning has a Rhumb-line Control function which reduces the pointing error at separation and ignition of solid motor. We could insert the payload into the orbit precisely by 'LVIC' guidance, suitable for low thrust propulsion in Post Boost Stage. We will present the flight results of the Guidance & Control (G&C) system and dynamics of Epsilon rocket.

DESTINY+ is a small-seized and high performance deep space vehicle proposed for public offering small-sized plan space science mission of ISAS/JAXA. DESTINY+ is injected into an extended elliptical orbit launched by Epsilon rocket. The orbit is spiraled upward by the low-thrust of IES. And the swing-by is designed to give DESTINY momentum to Asteroid Phaethon flyby. After Phaethon flyby, DESTINY+ plan to go back to Earth for gravity assist and go to another asteroid. DESTINY+ has several mission objectives, including: demonstration and experiment of space technology of interplanetary voyage; Phaethon flyby with reusable probe; compact avionics as for Engineering mission, and the investigation of the process to the end of evolution of primitive body; the limitation of initial state and the evolution process of the meteor shower dust as for Science mission. This paper discusses DESTINY+'s low-thrust trajectory design and the related system analysis. As for the spiral upward trajectory phase, the low-thrust trajectory is optimized by the multi-objective optimization using genetic algorithm. In this phase, we minimize the time of flight, the passage time of radiation belt, the work time of IES and the shadow time. After the spacecraft reaches to the moon's path, it utilizes the moon swing-by several times to connect to the transfer trajectory for Asteroid Phaethon. Parallel to the trajectory design, the radiation effect analysis, thermal environmental analysis, attitude analysis and ground station visibility analysis for operation are achieved. From these study, we can show the feasibility of the mission design of DESTINY+.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The Epsilon launch vehicle, the newest version of Japan’s solid propulsion rocket, made its maiden flight in September of 2013. The purpose of the Epsilon launch vehicle is to provide small satellites with responsive launching with low-cost, user-friendly and efficient launch system. The first flight was successfully finished, JAXA has been conducting intensive researches on a more powerful and lower cost version of Epsilon. In order to minimize technical risks and to keep up with demand of future payloads, JAXA plans to take a step-by-step approach toward Future Launch System. As the first upgrade toward Future Launch System, JAXA has started the development of the Enhanced Epsilon. This development is mainly the renewal of the second stage, and also includes the each subsystem’s improvement. This paper describes the development status of the Enhanced Epsilon’s Guidance and Control System.

Copyright © 2016 by the International Astronautical Federation. All rights reserved. Although reusable launch vehicle's necessity and significance, being cost-effective, eco-friendly and reliable, have been recognized in a long time, practical system still has never been realized except the Space Shuttle. There are two main reasons in this. One reason is that reusable vehicle's recurring cost is high. The other reason is that reusable vehicle, especially that upper stage, have the problem of aerodynamic heating during re-entry. We are considering new upper stage reusable launch vehicle with solid rocket booster, which clear these problems concerning reusable launch vehicle. For the first problem on the recurring cost, the application of the auto inspection system which is cultivated in solid rocket motor's development and launch operation is being considered. That is expected to reduce the inspection cost drastically after the vehicle flight. For the second problem on the re-entry, challenging technologies are applied in the upper stage. Those are material and structure with heat tolerance and lightness, active-cooling system to share the hydrogen with the liquid propulsion system, advanced guidance and control system, and so on. On the other hand, to the lower stage or booster, application of solid rocket is considered. Since the challenging upper stage's size is expected to vary through the iteration of design cycles, the lower stage should be stable and flexible with the thrust level in development phase. Because solid motors of various sizes are developed in JAXA/ISAS since the first small solid motor started to be developed in 1954, those development method has been efficiently accumulated. Then this legacy's utilization for the new system is expected to be quite beneficial. On these technological backgrounds, this paper describes the system study for new upper-stage reusable launch vehicle with the solid rocket booster.

Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. A fully reusable sounding rocket is proposed in ISAS/JAXA. This vehicle adopts single stage and VTVL(vertical takeoff and vertical landing). The vehicle is decelerated by aerodynamic resistance and turn over for vertical landing in return phase. After that, the vehicle is decelerated by the main engine for soft-landing. To return the vehicle, there is a problem that the study of the guidance control method and securing both the launch ability and the propellant for soft-landing in addition to the ability required of conventional vehicle. It is necessary to design the airframe shape and control method after studying the motion characteristics at the time of return to clear these problems. This paper shows the result of motion analysis about the return system under consideration, and considers from the point of view of the vehicle specifications and control system. And it indicates the guidelines of the airframe and control system that is suitable for the return of reusable sounding rocket.

© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved. This study is devoted to explore space trajectory for DESTINY (Demonstration and Experiment of Space Technology for INterplanetary voYage), which was proposed to ISAS (Institute of Space and Astronautical Science) Epsilon-class small program in 2013 based on the \Space Science & Exploration Roadmap” which is proposed by ISAS and later approved by the government committee of space policy. In the DESTINY mission, spacecraft is first injected into a low elliptical orbit by the Epsilon rocket, and it raises the altitude to reach the Moon using an ion engine system (IES). After that it is injected into a transfer orbit of L2 Halo orbit of the Sun-Earth system through gravity assist of the Moon. While the spacecraft revolves around the Earth for several hundred times, it increases its altitude little by little, and thus, launch time and the thrusting profile must be chosen properly. It is very important to note that there are many conicting requirements such as reduction of fuel consumption, total flight time, and the maximum eclipse time and so forth. To satisfy these requirements, many-objective evolutionary computation is applied to search for a better orbital design.

This paper describes achievements of the IKAROS project, the world's first successful interplanetary solar power sail technology demonstration mission. It was developed by the Japan Aerospace Exploration Agency (JAXA) and was launched from Tanegashima Space Center on May 21, 2010. IKAROS successfully deployed a 20 m-span sail on June 9, 2010. Since then IKAROS has performed interplanetary solar-sailing taking advantage of an Earth-Venus leg of the interplanetary trajectory. We declared the completion of the nominal mission phase in the end of December 2010 when IKAROS successfully passed by Venus with the assist of solar sailing. This paper describes the overview of the IKAROS spacecraft system, how the world's first interplanetary solar sailer has been operated and what were achieved by the end of the nominal mission phase. © 2012 Elsevier Ltd.

The trajectory design of the interplanetary mission "Demonstration and Experiment of Space Technology for INterplanetary voYage, DESTINY" is discussed. The trajectory optimization of low-thrust spacecraft pose a difficult design challenge. Moreover, DESTINY mission requires many constraints for the orbital design. In advance of optimizing the whole transfer mission, we investigated the basic theory to design which can take into account such constraints.

The trajectory design of the interplanetary mission "Demonstration and Experiment of Space Technology for INterplanetary voYage, DESTINY" is discussed. The trajectory optimization of low-thrust spacecraft pose a difficult design challenge. Moreover, DESTINY mission requires many constraints for the orbital design. In advance of optimizing the whole transfer mission, we investigated the basic theory to design which can take into account such constraints.

The Venus explorer, Akatsuki, was launch on 20 May 2010. After 200-day journey through the interplanetary transfer orbit, it reached the Venus at the altitude of 550 km on 7 Dec 2010. However, it experienced a trouble of the explorer's propulsion system and was not able to be the Venus orbiter. It now orbits the Sun with the period of 203 days. In this paper, we discuss the trajectory design strategies for Akatsuki mission by introducing the constraints which come from the observation orbit and the spacecraft system. The details of planning and the results of orbital maneuvers are also shown in this paper.

This paper presents IKAROS extended missions. IKAROS entered its extended operation phase at the beginning of 2011. In the extended operation, the spin rate was decreased to observe the deformation of the sail under low centrifugal force environment. On Oct. 18, 2011, IKAROS transferred to the reverse spin to enhance the knowledge about the effect of stiffness of membrane against the solar radiation pressure. We investigated the change of the attitude motion by the reverse spin mission. At the end of 2011, IKAROS moved to hibernation mode because the Sun angle was increased. We searched for IKAROS considering the attitude and orbital motion during hibernation. On Sep. 6, 2012, we succeedcd in tracking IKAROS which came out of hibernation. A solar power sail can be a hybrid propulsion system with a solar sail by activating the ultra-high specific impulse ion engines with the power generated by thin film solar cells. This paper also introduces an advanced solar power sail mission toward Jupiter and Trojan asteroids via hybrid electric photon propulsion.©2012 by the International Astronautical Federation.

In ISAS/JAXA, a fully reusable sounding rocket is proposed as one step for the future full-fledged reusable transportation system. This vehicle is a vertical take-off and vertical landing (VTVL) rocket vehicle. And this has the capability of ballistic flight to the altitude over 120 km and returning to the launch site. In the flight sequence, the vehicle takes off vertically and cuts main engine off about 100 seconds later, and reaches to an altitude about 120 km during the ballistic flight. After that, the vehicle turns back into a nose first attitude. During the return flight, the vehicle is guided to above the launch place. Then the vehicle makes a turnover maneuver to base first attitude form a nose first entry attitude. This makes it possible to achieve the deceleration and soft landing by its main engine thrust. It is considered that there are many technical concerns to realize this vehicle. To show the feasibility of the vehicle, technical demonstrations are under way in JAXA. One of the technical concerns is a turnover maneuver during the return flight. As for the inversion maneuver, it is considered for applying the aerodynamic turnover maneuver which is caused by the differences of pitching moment depending on the vehicle configurations. For example, vehicle configurations are changed by deploying strake, a kind of canard. To verify the turnover maneuver capability, it is considered to demonstrate the glide tests using the small sized vehicle model. In this demonstration, the technical problems for the turnover maneuver as for the vehicle dynamics and guidance control strategy will be investigated. In this paper, the turnover maneuver control for the nose first entry of the small sized vehicle is numerically simulated and the return flight guidance to a point is proposed. And turnover maneuver dynamics of the vehicle is investigated for practical use. The guidance and control method proposed here would be applied for the full scale reusable sounding rocket vehicle. Copyright © (2012) by the International Astronautical Federation.

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.

In this paper, the attitude dynamics of IKAROS, which is spinning solar sail, is presented. Multi Particle Model (MPM) and First Mode Model of out-of-plane deformation (FMM) are introduced to analyze the out-of-plane oscillation mode of spinning solar sail. Considering the thruster configuration of IKAROS, the force on main body and membrane by thruster plume as well as reaction force by thruster are integrated into MPM. The attitude motion after sail deployment or reorientation using thrusters can be analyzed by MPM numerical simulations precisely. The out-of-plane oscillation of IKAROS is governed by three modes derived from FMM. FMM is simple and valid for the design of attitude controller. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

A plume impingement analysis on the thruster deck and the membrane of IKAROS was performed based on available flight data. IKAROS employs gas-liquid thruster system as RCS for spin-rate adjustment and attitude control. Gas jet plume from the thruster nozzles leads to considerable undesired disturbance force on the spacecraft, because IKAROS has a large size of flexible membrane whose diagonal distance is 20m. To minimize thrust loss by means of improvement in maneuver planning and to obtain the local distribution of the momentum flux on the membrane for future solar sail missions design, it is important to evaluate the plume impingement on the solar and propose a plume distribution model. Therefore this paper presents actual values of mechanical load on IKAROS and provides a distribution of plume force on the surface. In proposal of a plume model, the conventional plume density distribution model [1] and plume impingement model [2] were used. And parameters in these models that describe the physical interaction between plume flow and the surface of membrane were identified based on fight data.

In ISAS/JAXA, a fully reusable sounding rocket is proposed as one step for the future full-fledged reusable transportation system. This vehicle aims to be low in cost and to be high frequency of experimental opportunities by its reusability. The enlargement of the flight profile flexibility could acquire the high qualities of the experimental conditions. This vehicle is a fully reusable vertical take-off and vertical landing (VTVL) rocket vehicle. And this has the capability of ballistic flight to the altitude over 120 km and returning to the launch site. In me flight sequence, the vehicle takes off vertically and cuts main engine off about 100 seconds later, and reaches to an altitude about 120 km during the ballistic flight. After that, the vehicle entries into atmosphere and decelerated by aero braking, and vertically lands to the launch place. In landing phase, this vehicle makes a turnover maneuver from a nose-first entry attitude. This makes it possible to achieve the deceleration and soft landing by its main engine thrust. As for the inversion maneuver of the vertical landing vehicle, the aerodynamic turnover maneuver is considered. After turnover maneuver, control torque of a reaction control system (RCS) stabilizes the vehicle in a tail-first attitude before a re-ignition of main engines. To verify the turnover maneuver capability, it is considered to demonstrate the glide tests using the small sized vehicle model. In this demonstration, the technical problems for the turnover maneuver as for the vehicle dynamics and guidance control strategy will be investigated and the guidance and control method will be modified for the practical use to achieve the system requirements. In mis paper, the turnover maneuver control for the nose first entry of the vehicle is numerically simulated and then the guidance strategy for the vehicle is considered. It is important for the guidance that the vehicle attitude could be stabilized after the turnover maneuver. By the given moment torque of the RCS, it is needed for the vehicle to maintain the attitude at 180 degree of angle of attack.

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.

IKAROS (Inter-planetary Kite-craft Accelerated by Radiation Of the Sun) is a Small Solar Power Sail Demonstrator which deploys the membrane and generates solar power by means of thin film solar cells. IKAROS was launched by H-IIA rocket from Tanegashima Space Center on 21st May 2010 as a piggy back spacecraft of Planet-C "AKATSUKI" Venus climate orbiter. IKAROS (and also AKATSUKI) is the first spacecraft managed by JAXA (Japan Aerospace Exploration Agency)'s system safety activity at ISAS (Institute of Space and Astronautical Science). IKAROS itself has to be managed to avoid occurring the critical hazard not only to worker, rocket and facility but also to the main spacecraft AKATSUKI because IKAROS is a piggy back spacecraft. GSE (Ground Support Equipment) and operation procedures are also objects for system safety. IKAROS has several hazardous functions like a separation mechanism, an unfolding mechanism, batteries, a pressure system and so on. There is also a problem to solve as for SCC (Stress Corrosion Cracking) of aluminium alloys used for structure material. There is three or four phase for system safety management before it is allowed to bring the spacecraft to the launch site. In this paper, it is shown that how IKAROS demonstration team manage the system safety activity, which is, the design modification, the operation management, the verification experiment for hazard control. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun) is a small demonstration spacecraft of Solar-Sail which deploys the sail-membrane in the space to be accelerated by radiation of the sun. IKAROS has a reaction control system (RCS) to spin up itself before the centrifugal deployment of sail-membrane and control its attitude. Because of strict safety requirement, very short manufacture schedule, and relatively low manoeuvre (delta V) requirement of IKAROS, IKAROS RCS has adopted the gas-liquid equilibrium propulsion system which stores chlorofluorocarbon alternative HFC-134a as liquid-phase in the tank, extract the vapor of HFC-134a from the tank, and eject the vapor from the thruster nozzle. Cold gas through the nozzle realizes simplified thruster system, therefore can shorten manufacture schedule, but resultant 1sp is lower than that of hot gas system (ex.: Hydrazine system). However this system can achieve higher manoeuvre than cold gas system (ex.: GN2 system) because of higher density in the tank. These are why it is suitable for small satellite propulsion system. There are two major technical problems during its development. One is how to avoid the thrust degradation due to the condensation of the equilibrium gas flow in the nozzle. The other is how to extract the vapor of propellant from the tank. This paper shows the design method to solve these problems, the confirmation test results of the condensation in the nozzle, and the verification results of the vapor extracting device. Copyright ©2010 by the International Astronautical Federation. All rights reserved.

The Japan Aerospace Exploration Agency (JAXA) will make the world's first solar power sail craft demonstrate for both its photon propulsion and thin film solar power generation during its interplanetary cruise. The spacecraft deploys and spans its membrane of 20 meters in diameter taking the advantage of the spin centrifugal force. The spacecraft weighs approximately 315kg, launched together with the agency's Venus Climate Orbiter, PLANET-C in 2010. This will be the first actual solar sail flying an interplanetary voyage.