伊藤 琢博

Deci-hertz Interferometer Gravitational Wave Observatory (DECIGO) is the future Japanese space mission with a frequency band of 0.1 Hz to 10 Hz. DECIGO aims at the detection of primordial gravitational waves, which could be produced during the inflationary period right after the birth of the universe. There are many other scientific objectives of DECIGO, including the direct measurement of the acceleration of the expansion of the universe, and reliable and accurate predictions of the timing and locations of neutron star/black hole binary coalescences. DECIGO consists of four clusters of observatories placed in the heliocentric orbit. Each cluster consists of three spacecraft, which form three Fabry-Perot Michelson interferometers with an arm length of 1,000 km. Three clusters of DECIGO will be placed far from each other, and the fourth cluster will be placed in the same position as one of the three clusters to obtain the correlation signals for the detection of the primordial gravitational waves. We plan to launch B-DECIGO, which is a scientific pathfinder of DECIGO, before DECIGO in the 2030s to demonstrate the technologies required for DECIGO, as well as to obtain fruitful scientific results to further expand the multi-messenger astronomy.

© 2020 IAA Onboard computation of a fuel-optimal trajectory is an indispensable technology for future lunar and planetary missions with pinpoint landings. This paper proposes a throttled explicit guidance (TEG) scheme under bounded constant thrust acceleration. TEG is capable of achieving fuel-optimal large diversions with good accuracy and can find optimal solutions. Thus far, the TEG algorithm is unique as it offers an explicit and simultaneous search method for the fuel-optimal thrust direction and thrust magnitude switching in predictor-corrector iterations. Fast numerical search is realized with a straightforward computation of seven final states (position, velocity, and the Hamiltonian) from seven unknowns (six adjoint variables for position and velocity and one final time). In addition, global convergence capability is enhanced by implementing the damped Newton's method. A number of simulations of large diversions show the excellent convergence of the TEG algorithm within at most 15 iterations from a cold start. The experimental results of the runtime measurement of the TEG algorithm support its real-time feasibility on a flight processor. These features of the TEG are suitable for onboard guidance of pinpoint landings.

<p>本論文はSS-520 5号機のラムライン制御系開発および飛翔結果を論じる．ラムライン制御系は，第1段燃焼終了後から第2段燃焼開始前までに実施される唯一の姿勢制御期間において，第2段，第3段加速方向への姿勢変更およびニューテーション減衰制御を実施するシステムである．SS-520 5号機のスピンの高速性（ノミナル1.6 Hz）に加え，厳しい制御時間・消費燃料要求に対処するため，ラムライン制御系には，スラスタ噴射遅れの補償機能やラムライン制御中のヘディング方向の逐次修正機能による追従性向上，スラスタ印加力積を段階的に削減する制御方式による少消費燃料・短制御時間・高収束性の同時実現等，姿勢制御の高性能化に資するアイデアが取り込まれた．姿勢制御に関する飛翔結果は良好で，事前想定のノミナル結果とおおよそ一致した．SS-520 5号機で開発された姿勢制御技術が今後，高速スピンするロケットや宇宙機の高度な姿勢制御技術として活かされることに期待する．</p>

© 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.

<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>A Fault Detection, Isolation, and Recovery (FDIR) algorithm for attitude control systems is a key technology to increasing the reliability and survivability of spacecraft. Micro/nano interplanetary spacecraft, which are rapidly evolving in recent years, also require robust FDIR algorithms. However, the implementation of FDIR algorithms to these micro/nano spacecraft is difficult because of the limitations of their resources (power, mass, cost, and so on). This paper shows a strategy of how to construct a FDIR algorithm in the limited resources, taking examples from micro deep space probe PROCYON. The strategy focuses on function redundancies and multi-layer FDIR. These ideas are integrated to suit the situation of micro/nano interplanetary spacecraft and demonstrated in orbit by the PROCYON mission. The in-orbit results are discussed in detail to emphasize the effectiveness of the FDIR algorithm. </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>

© 1965-2011 IEEE. Recently, a variety of small spacecraft have been launched and used for interplanetary missions. Conventionally, reaction wheels (RWs) and thrusters are used for these attitude control systems in almost all interplanetary spacecraft. While these actuators are promising for attitude control, the lifetime of the spacecraft mission is limited by the extra fuel needed for the thrusters. Moreover, it is difficult to install thrusters in all small spacecraft due to low reliability and strict limitations on mass and power consumption. To obtain both fuel-free and available attitude control for small spacecraft, this study proposes an interplanetary magnetic attitude control system including attitude stabilization and angular momentum unloading based on an interplanetary magnetic field (IMF) Kalman filter. In the proposed method, an electromagnetic coil interacting with the IMF is used for an attitude control system. To achieve the proposed method, the faint magnetic field must be detected. However, the IMF is too weak to sense using only on-board magnetic sensors. To deal with the technical issue, this study proposes a magnetic attitude control system with an unscented Kalman filter using gyro measurements and generated magnetic moment by the electromagnetic coil to estimate the weak magnetic field. With the estimated magnetic field vector, the spacecraft can achieve fuel-free attitude stabilization and RW unloading under the constraints. This proposed system does not require fuel for attitude control. Furthermore, the simple structure and electronic circuits of the electromagnetic coil allow the spacecraft to achieve a simple and reliable attitude control system. Numerical simulations demonstrate the effectiveness of the proposed attitude stabilization and RW unloading methods.

© 2018 IAA This study proposes a solar sailing method for angular momentum control of the interplanetary micro-spacecraft PROCYON (PRoximate Object Close flYby with Optical Navigation). The method presents a simple and facile practical application of control during deep space missions. The developed method is designed to prevent angular momentum saturation in that it controls the direction of the angular momentum by using solar radiation pressure (SRP). The SRP distribution of the spacecraft is modeled as a flat and optically homogeneous plate at a shallow sun angle. The method is obtained by only selecting a single inertially fixed attitude with a bias-momentum state. The results of the numerical analysis indicate that PROCYON's angular momentum is effectively controlled in the desired directions, enabling the spacecraft to survive for at least one month without momentum-desaturation operations by the reaction control system and for two years with very limited fuel usage of less than 10 g. The flight data of PROCYON also indicate that the modeling error of PROCYON's SRP distribution is sufficiently small at a small sun angle (<10°) of the order of 10−9 Nm in terms of its standard deviation and enables the direction of the angular momentum around the target to be maintained.

Thanks to recent lunar exploration missions, high-resolution lunar surface observation data was obtained. In future lunar exploration, landing is being requested at a specific point having higher scientific interest than other areas. The SLIM project is demonstrating pinpoint landing technology, which entails a combination of “autonomous image-based high-precision navigation technology” and “autonomous guidance technology intended to generate a fuel-optimum landing trajectory.” This paper presents powered descending trajectory design in terms of trajectory optimization. As usually considered in general space mission development, an optimal solution in terms of minimum fuel consumption is the basis of investigation. This study addresses trajectory optimization considering specific objective functions derived from practical constraints regarding mission design, such as altitude, downrange length, and visibility from ground stations. In this paper, nominal trajectory design considering minimum fuel consumption is first presented, followed by parametric studies to identify the sensitivity to changes in initial conditions under which powered descending starts. Finally, trajectory optimization results with various types of objective functions are presented.

© 2017 The Japan Society for Aeronautical and Space Sciences. This paper describes development strategies and on-orbit results of the attitude determination and control system (ADCS) for the world's first interplanetary micro-spacecraft, PROCYON, whose advanced mission objectives are optical navigation or an asteroid close flyby. Although earth-orbiting micro-satellites already have ADCSs for practical missions, these ADCSs cannot be used for interplanetary micro-spacecraft due to differences in the space environments of their orbits. To develop a new practical ADCS, four issues for practical interplanetary micro-spacecraft are discussed: initial Sun acquisition without magnetic components, angular momentum management using a new propulsion system, the robustness realized using a fault detection, isolation, and recovery (FDIR) system, and precise attitude control. These issues have not been demonstrated on orbit by interplanetary micro-spacecraft. In order to overcome these issues, the authors developed a reliable and precise ADCS, a FDIR system without magnetic components, and ground-based evaluation systems. The four issues were evaluated before launch using the developed ground-based evaluation systems. Furthermore, they were successfully demonstrated on orbit. The architectures and simulation and on-orbit results for the developed attitude control system are proposed in this paper.

In this paper is presented a microgravity experiment system utilizing a high altitude balloon. The feature is a double shell structure of a vehicle that is dropped off from the balloon and a microgravity experiment section that is attached to the inside of the vehicle with a liner slider. Control with cold gas jet thrusters of relative position of the experiment section to the vehicle and attitude of the vehicle maintains fine microgravity environment. The design strategy of the vehicle is explained, mainly referring to differences from the authors' previous design. The result of the flight experiment is also shown to evaluate the characteristics of the presented system.