宮﨑 康行

In recent years, a large space film structure having a thickness of several micro and a shape of several to several tens of meters attracts attention, and various storing methods have been studied. Considering the thickness of the film surface at the time of winding before launching, there is a problem that circumferential difference occurs inside and outside of the folded film surface. In order to solve this problem, a method of solving the difference between the inner and outer circumference by predicting the inner / outer circumferential difference arising from the film surface and the thickness of the device and managing the phase has been proposed. On the other hand, the point that the target value for adjusting the phase is unknown and empirical was pointed out, and as a result of adjusting the phase, the wave-like slack that occurred caused the unevenness of the film thickness in the circumferential direction. In this research, we derive target value of phase management analytically, compare with experiment, and verify.

A deployable structure serves an important function in small satellites as their performance is improved steadily. Larger deployable solar array paddles are expected as the satellite requires a larger amount of electric power. The synthetic aperture radar (SAR) system for small satellites requires a deployable antenna. Various deployable membrane structures have been proposed for the deorbit of micro/nanosatellites in low Earth orbit. The actuators and devices for a deployable structure continue to progress along with the development of a deployable structure. The hold-release mechanism has been becoming smaller and simpler. The shape control devices are actively researched for a high-precision deployable structure. There are several requirements imposed on the deployable structures, e.g., low cost, lightweight, small volume in stored configuration, reliable deployment, large aperture in deployed configuration, and accuracy/repeatability of a deployed shape. Dedicated efforts are made to satisfy these requirements in research and development of a deployable structure/device. This paper provides an overview of the past and current research and development of a deployable structure for small satellites.

© 2018 Univelt Inc. All rights reserved. In recent years, a large space film structure having a thickness of several micro and a shape of several to several tens of meters attracts attention, and various storing methods have been studied. Considering the thickness of the film surface at the time of winding before launching, there is a problem that circumferential difference occurs inside and outside of the folded film surface. In order to solve this problem, a method of solving the difference between the inner and outer circumference by predicting the inner / outer circumferential difference arising from the film surface and the thickness of the device and managing the phase has been proposed. On the other hand, the point that the target value for adjusting the phase is unknown and empirical was pointed out, and as a result of adjusting the phase, the wave-like slack that occurred caused the unevenness of the film thickness in the circumferential direction. In this research, we derive target value of phase management analytically, compare with experiment, and verify.

The gossamer structures combined with deployable truss structures and membrane structures are suggested for recently proposed large space structures. Several types of the boom are expected as the members of the deployable truss. Tape spring is one of the candidates because the rolled-up tape spring can extend automatically without any actuator by releasing the stored strain energy, i.e. it has self-extensibility. Bi-convex boom and the braid coated biconvex boom (BCONTM boom) are also candidates. If those booms are used for the structural member, the truss can deploy by employing the self-extensibility of the booms, i.e. it has characteristics of self-deployable structure. In this paper, the structural characteristics of those booms are described and the effect of the tension of braid mesh on the stiffness of the boom is investigated. The stiffness of the boom is estimated theoretically and experimentality. This paper the difference between those booms and the closed-section boom with the end of the boom fixed.

Recent expansion of space exploitation, especially advance of micro/nano satellite technology, leads to an increased demand for planer deployable membrane structure. One of the most important requirements is the simplicity and the realiability of the deployment as well as the stiffness of the structure. The self-deployable truss composed of bi-convex booms is one of the candidates that satisfy those requirements. In this paper, a planer membrane structure supported by the self-deployable truss is proposed and its structural characteristics is presented. The technical problems that must be solved for the realization of the proposed structure in space are discussed and the possible solutions of those problems are also given in the paper.

In recent years, there are a lot of research on membrane space structure. It is essential to estimate the deployment behavior of these structure. However, It is difficult to use actual size membrane for experiment because it is affected by gravity and air drag. Therefore, prediction based on experiments has yet to be done. Accordingly, we make some sexperiment using caled-down model. But small scale models and large scale models do not always perform the same deployment behavior[1]. So if similarity deployment rule between large-scale model and small-scale models is established, we can predict the deployment behavior of the flight model membrane from ground experiments using small-scale membrane. This paper propose the similarity rule between large-scale and small-scale models at deployment.

As a system for construction of a large space structure and debris mitigation, deployable structural systems are indispensable. Gossamer structure realize a larger and lighter structure by storing and deploying using deformation and buckling of members. In particular, buckling easily occurs in the flexible structural member, and its large deformation enables smooth deployment. Furthermore, external impetus to deploy is suppressed to be small because deformation by the buckling progresses greatly only with the strain energy of the member. The paper proposes a new available deployable structure “Deployable Cube”, which is a bistable structure applying buckling actively. The authors developed a prototype of the Deployable Cube for a Cubesat. In order to investigate the structural properties, eigen mode analysis after deployment of the prototype model was performed, and the stiffness is indicated. Furthermore, dynamic buckling analysis using the original method proposed by the authors was carried out for the initial stage of the deployment, and it is indicated that the estimation of buckling mode is valid.

Deployable structures are necessary for spacecraft to challenge advanced missions. It is important in designing the deployable space structures that they are easily deployable and reliably repeatable. Conventional study for improving the repeatability was conducted by investigating errors and its effect to the deployment. However, there was no general numerical method to evaluate the reliability efficiently and quantitatively, not estimating any errors. Therefore, the authors proposed an original numerical method to evaluate quantitatively the reliability of the deployment by detecting buckling in a dynamic analysis of FEM. The repeatability is decreased due to occurrence of the buckling caused by the errors. Therefore, a structure not occurring the buckling should be selected for designing of a reliably repeatable structure. The buckling is detected by non-positive eigenvalues of a stiffness matrix of the structure in static analysis. However, detection of the buckling in dynamic analysis is difficult because the eigenvalue is also non-positive when the structure has rigid-body motion. This study solved the problem by proposing a method to discriminate the buckling from the rigid-body motion. In the recent study, we found a problem of false detection of the buckling by the conventional method, and we also found a method to solve the problem. The cause of the problem is using the eigenvalue for the judgement of the buckling because it cannot be evaluate that the sign of the eigenvalue derives whether from the rigid-body motion or from the deformation. Therefore, the modified method does not use the eigenvalue but uses the work of the deformation for the judgement of the buckling. Furthermore, a method to evaluate instability of the structure quantitatively is desired when the buckling occurs in the deployment of a structure for the spacecraft. When the buckling occurs, small disturbance sets off grave displacement. Therefore, this study proposed the method to evaluate the instability quantitatively by calculating disturbance force and buckling displacement as index values of the instability based on the equation of motion. Additionally, this study proposed a method to visualize structural instable area in an entire structure. By looking the instable area, developers of a spacecraft can take measures for the instable area, and improve the reliability of the structure. Finally, we applied the proposed method to an analysis of presumed actual space structure, and inquired about the buckling.

Technology of deployable space structures is necessary for spacecraft to challenge advanced missions. It is important in designing the deployable space structures that they are easily deployable and reliably repeatable. Traditional approach for improving the repeatability was conducted by investigating errors and its effect to the deployment. However, the traditional approach has a problem that results change depending on estimation of the errors. With that background, this study proposes numerical methods to enable selection of robust deployable structures against the errors. The repeatability is decreased due to occurrence of the buckling caused by the errors. Therefore, a structure not occurring the buckling should be selected for designing of a reliably repeatable structure. The buckling is detected by non-positive eigenvalues of a stiffness matrix of the structure in static analysis. However, detection of the buckling in dynamic analysis is difficult because the eigenvalue is also non-positive when the structure has rigid-body motion. This study solved the problem by proposing a method to discriminate the buckling from the rigid-body motion. Furthermore, a method to evaluate instability of the structure quantitatively is desired when only structures occurring the buckling are available for the spacecraft. When the buckling occurs, small disturbance sets off grave displacement. Therefore, this study proposed the method to evaluate the instability quantitatively by calculating disturbance force and buckling displacement as index values of the instability based on the equation of motion. Finally, it was confirmed that the proposed methods are appropriate by the dynamic analyses of truss arch.

This paper derives closed-form solutions for the local deformation of a bi-convex boom under circular bending, and the resulting strain energy and self-extending force. Convex tapes and bi-convex booms that consists of a pair of convex tapes can be stored into a small volume and have high specific rigidity. They extert a self-extending force when stored cylindrically. Therefore, they have been proposed as members of deployable space structures. In this paper, two types of bi-convex booms are considered. In the first, the tapes of the bi-convex boom are bonded to each other at their edges; in the second, the tapes are wrapped in a cylindrical braid mesh. The latter is called a BCON (braid-coated bi-convex) boom. The tape of a BCON boom can slip on each other, and do not separate from each other because of the tension of the mesh net. Consequently, the BCON boom can be used in an ultralight self-deployable structure with quite high stowage volume efficiency and specific rigidity. However, structures using convex tapes or BCON booms have been designed and developed through a trial-and-error process because there is no appropriate formula for the self-extending force of convex tapes. This paper proposes a formula for the deformation of a convex tape that is initially bent into a circular shape. The deviation from the circular shape is obtained by solving the equilibrium equations. The deformation of a bi-convex boom is also derived by using the solution for a convex tape. Thus the theory described in this paper contributes to the design of space structures using convex tapes in bi-convex booms, as well as to the structural mechanics of flexible beams.

Recently, space science mission has been more and more complicated and advanced. Its scale has been larger and larger. Accordingly deployable structure grows increasingly important. The space verification is quite important or necessary for deployable structures that have not been proved in space. Thus more and more opportunities with low-cost are desired for the space verification. Nano-satellite in low earth orbit is most suitable for the test-bed of the small scale model of the structures. Moreover, such a deployable structure can be used as the de-orbit device, so that it can be a solution of the space debris issue. This paper discusses the importance and the methodology of the space verification of the key technologies for advanced large deployable structures by using nano-satellites.

大学宇宙工学コンソーシアム(University Space Engineering Consortium:UNISEC)は10年余にわたる国内高専・大学への実践的宇宙開発の支援を経て,宇宙ハンズオントレーニングの有用性・有効性を確信し,活動を宇宙開発新興国を含む他地域に広げるべく,2011年に国際委員会を組織した.国際向けの活動としては,超小型衛星のミッションアイディア国際コンテスト開催,海外の新興国への衛星開発教育の実践,超小型衛星シンポジウムの事務局運営などがある.UNISECでは,2020年までに100以上の国で大学生が実践的宇宙開発に参加できるような世界を作ろうという「VISION 2020-100」を発表し,世界各地にUNISECのような大学連携組織を作り,それらの組織を横断的につなぐUNISEC-Globalを設立しようという提案を国連の会議等で行い,2013年11月には,第1回UNISEC世界大会(The 1st UNISEC-Global Meeting)を実施し,UNISEC-Globalの設立が宣言された.本稿では,UNISEC国際展開の経緯と成果およびその将来展望,課題について考察する.

A space inflatable extension mast designed on an idea of structural rigidization simulation system has been projected and provided to verify its engineering technology and to obtain the structural property for a long-term operation in space environment. The inflatable mast extended successfully in orbit, and has sent the telemetry data for more than seven months, so far. The experiment progress meets both its minimum and full success criteria for this mast. The simulation model of this mast is made up based upon the ground test, and predicts the natural frequency in orbit. Lightweight extendible masts are fundamental and essential structural elements to construct space structures; therefore a pivotal first step of conditions to utilize space inflatable structures has been actually achieved.

In recent years, the use of membranes for space craft applications has attracted a great deal of attention, as so called gossamer structures. However gossamer structures do not have practical applications at present. One problem in space research is design time reduction. Currently, JAXA is working on a spin type solar power sail spacecraft. Computational analysis of the sail membrane's complex mathematical model is difficult and time consuming. Therefore this has a negative impact on design problems. The model reduction technique is required. With a feasible shortened computational analysis period, a trade-off would be possible when designing and determining the operation procedure. It is a necessary tool in order to put the gossamer structure to practical use. The authors have already suggested a possible benefici al effect of the empirical model reduction for gossamer stru crure. Unfortunately the method couldn't deal with the geo metric constraint, for example rigid body with cable constra int. The purpose of this research is to construct a reduction model of geometrical constrained gossamer structure. Generally, the equations of motion of gossamer structures are highly nonlinear and stiff differential equations. For this reason we employ the geometrically nonlinear FEM code. The code is based on the energy momentum method (EMM), so the numerical time integration is unconditionally stable. In this paper, we constructed a penalty method for the geometrical constraint based on the EMM. This permits expression of the geometric constraint in a model. It is possible to apply empirical model reduction techniques to this mathematical model. This can make constructing of a reduction model including geometric constraint. We will show the reduction model can approximate the full-order model with sufficient accuracy. Copyright (c) 2010 by JSME.

This paper reports the history of the student pico-satellite program in Nihon University, especially about how a small laboratory which had no knowledge on satellite technology has succeeded the launch and operation of the student pico-satellite. The author hopes that his method would be a helpful example for supervisors who will start the student satellite program.

This paper proposes a method to store a large solar-sail membrane while ensuring repeatability of its stored configuration. The feasibility and effectiveness of the method is verified through a series of sail-storage experiments using 10m-size membranes. Large membranes used as a solar sail should be stored compactly to save the launch volume; in addition, their stored configuration should be sufficiently predictable in order to guarantee reliable deployment in orbit. However, it is difficult to store a large membrane compactly with sufficient repeatability because of the finite thickness of the membrane. This paper classifies the existing and proposed folding patterns that can consider the finite-thickness of membranes. This paper then demonstrates the feasibility of "bulging roll-up" experimentally, and evaluates the repeatability of its stored configuration quantatively. © 2012 AIAA.

The Japan Aerospace Exploration Agency (JAXA) makes the world's first solar power sail craft IKAROS demonstration of photon propulsion and thin film solar power generation during its interplanetary cruise. The spacecraft deploys and spans a membrane of 20 meters in diameter using the spin centrifugal force. It also deploys thin film solar cells on the membrane, in order to evaluate its thermal control property and anti-radiation performance in the real operational field. The spacecraft weighs approximately 310kg, launched together with the agency's Venus Climate Orbiter, AKATSUKI on May 21, 2010. This paper presents the summary of development and operation of IKAROS.

The present study develops a new three-dimensional Timoshenko beam finite element whose length can be varied during transient dynamic analysis. The variable-length element enables the dynamic deployment analysis of flexible appendages with nonnegligible bending stiffness. In addition, the developed scheme employs an implicit time integration whereby energy and momentum in the system are properly conserved, and no artificial numerical dissipation is introduced. The developed beam element is then used in an finite element model of a solar sailcraft, and its deployment dynamics are analyzed allowing for the nonzero bending stiffness of the bundled membranes, as well as the effect of some realistic design imperfections.

Japan Exploration Agency (JAXA) launched a powered solar sail "Interplanetary Kite-craft Accelerated by. Radiation Of the Sun (IKAROS)" on May 21, 2010. One of the primal technologies demonstrated at IKAROS is the deployment of the sail whose diameter is 20m class. After the launch, the deployment operation was performed and successful expansion of the sail was confirmed. The deployment sequence in IKAROS consists of static first stage and dynamic second stage. In this paper, the flight data and observed dynamic motion during the first stage deployment are reported. These are compared with the results of numerical simulations using multi-particle model, and the accuracy and availability of this model is discussed. Copyright ©2011 by the International Astronautical Federation. All rights reserved.

The Solar Sail "IKAROS" (Interplanetary Kite-craft Accelerated by Radiation Of the Sun), which is the first solar sail on orbit, consists of 20m diagonal length square membrane made of 7.5mm thickness polyimide film. The sail membrane is deployed by centrifugal force due to spinning motion of the spacecraft. The wrapping fold is applied to realize stable deployment property. The manufacturing process of the large membrane with the folding techniques are described in detail. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

The paper will try to discuss the issue of reducing the time to calculate the nonlinear dynamics simulation of the gossamer structure, especially for the solar sail, because one of the problems in the gossamer structure is design time reduction. Therefore the present paper describes the low-order model of a spin type solar sail dynamics. The model reduction technique is required to shorten the design period. However there is no model reduction methodology about the solar sail dynamics. In this paper the full-order model of a spin type solar sail craft is modeled by geometrically nonlinear FEM based on the energy momentum method (EMM), so the numerical time integration is unconditionally stable. The low-order model is constructed by an empirical model reduction method. To conserve the geometrical relation of constraint conditions on the low-order space, we constructed a penalty method for the geometrical constraint based on the EMM. We will show the reduction model of the spin type solar sail model can approximate the full-order model with sufficient accuracy. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Japan Exploration Agency (JAXA) launched a powered solar sail "Interplanetary Kite-craft Accelerated by Radiation Of the Sun (IKAROS)" on May 21, 2010. One of the primal technologies demonstrated at IKAROS is the deployment of the sail whose diameter is 20m class. After the launch, IKAROS performed the deployment sequence and have confirmed that the membrane was successfully expanded. In this paper, the flight data and observed dynamic motion via deployment are reported. These are compared with the results of numerical simulations using multi-particle model, and the accuracy and availability of this model is discussed. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Japan Aerospace Exploration Agency (JAXA) launched the solar power sail orbiter "Interplanetary Kite-craft Accelerated by Radiation Of the Sun: IKAROS", on May 21st, 2010. IKRAROS was launched by H-IIA 17th vehicle with "Venus Climate Orbiter: AKATSUKI". IKAROS demonstrates a new propulsion technology of utilizing photons from the sun for deep space exploration, which is called the Solar Power Sail technology. In a case of the solar system exploration, an ion-propulsion engine is effective as a main propulsion system because it has high specific impulse and it can provide a continuous acceleration. However, the ion-engine needs high electric power in proportion to its performance. The solar power sail technology can be a hybrid engine, which can provide high electric power generated by very thin flexible solar arrays attached on the solar sail, while obtaining acceleration generated on the solar sail by the sun radiation. IKAROS succeeded in deployment the solar power sail in an interplanetary orbit, on June 9th, 2010, the first in the world, and we could obtain various flight data of the solar power sail deployment mission. We report the details of the mission system of IKAROS that applying a new deployment method, and the flight data obtained actually from IKAROS in inter-planetary orbit. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.

Japanese spin-type solar power sail KAROS was launched in May 21 st, 2010 by JAXA, and the 20m-sized sail membrane with 7.5μm thickness was successfully deployed in last June. During the design and development process of the sail membrane structure in last three years, the nonlinear finite element elasto-dynamics code named NEDA has been improved for the prediction of the deployment motion of the sail membrane of IKAROS. The formulation of the dynamics in NEDA is based on the energy momentum method (EMM), which preserves the total energy, the linear momentum and the angular momentum. In this paper, the theory of the finite element dynamics implemented in NEDA is summarized and the mathematical model of the deployment dynamics of IKAROS is shown in detail. The numerical result is compared with the flight data, which shows the proposed numerical method is available for the prediction of the nonlinear motion of gossamer structures. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc.