We have been studying long-lifetime helicon plasma thrusters as the Helicon Electrodeless Advanced Thruster (HEAT) project. Two important elements of the proposed helicon plasma thruster are a generation of a dense source plasma using a helicon wave, and an acceleration of the plasma by the Lorentz force using the product of the induced azimuthal current and static radial magnetic field. Here, in order to eliminate damage of electrodes, both generation and acceleration schemes are operated in non-contact condition between the plasma and electrodes. Acceleration schemes use two type of coils: rotating magnetic field coils and azimuthal mode number m = 0 ones. These studies have been carried out on the Large Mirror Device (LMD), which has two types the magnetic field source, permanent magnets and electromagnets, and the Small Helicon Device (SHD), which has small diameter discharge tubes. In this paper, current performances of acceleration schemes are reported.

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We propose a new thrust stand measuring thrust variation beyond the resonant frequency. The proposed thrust stand is based on null-balance thrust measurement method with acceleration-based compensation. A controller monitors pendulum motion and adjusts a solenoid actuator in such a way that the pendulum deflection angle is kept constant. Simultaneously, acceleration and solenoid-actuator current were measured during active control. Instant thrust is directly determined from the current and acceleration, because spring constant and damping factor originating from wires and the thrust stand can be ignored due to null-position active control. The prototype thrust stand successfully measured thrust with its variation in the frequency range from 0 to 60 Hz, whereas conventional null-balance-type stand yielded 370-% overestimation at a resonant frequency of 20 Hz. Beyond 60 Hz, the proposed thrust stand showed errors in the magnitude of evaluated thrust and phase shift. The deterioration in accuracy is caused by the translational flexibility in torsional hinges. Measurement of translational spring constant of the torsional hinge, and numerical simulation with a pendulum fixed with a hinge having rotational and translational flexibility yielded a correlation between frequency and accuracy, which is similar to that of the experiment.

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Hayabusa2 is the second asteroid sample return mission by JAXA. The ion engine system (IES) for Hayabusa2 is based on that developed for Hayabusa with modifications necessary to improve durability, to increase thrust by 20%, and to reflect on lessons learned from Hayabusa mission. Hayabusa2 will rendezvous with a near-earth asteroid 1999 JU3 and will take samples from its surfaces. More scientific instruments than Hayabusa including an impactor to make a crater and landers will be on board thanks to the thrust enhancement of the IES. An improved neutralizer with stronger magnetic field for longer life has been under endurance test in diode mode since August 2012 and has accumulated the operational hours of 25600 h ( > mission requirement: 14000 h) by July 2015. The IES flight model was developed within 2.5 years. The spacecraft was launched from Tanegashima Space Center in Kagoshima Prefecture on-board an H-IIA launch vehicle on December 3, 2014.

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A miniature microwave discharge plasma thruster for micro satellites is under development, with the target power consumption, thrust and specific impulse of 10 W, 1 mN and 1000 sec, respectively. Thrust was estimated from the measured ion beam current and ion energy distribution function using an ion collector and a retarding potential analyzer. Several parameters, including antenna configuration, length of the discharge chamber and configuration, were optimized for the improvement of the thrust performance. The overall thrust was estimated to be 82 μN at the incident microwave power of 10 W and the mass flow rate of 82 μg/sec.

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Research on a debris removal system that uses an electrodynamic tether (EDT) system has been conducted in JAXA. The EDT system requires an active electron emission device to drive a large electric current through the tether for obtaining adequate de-orbit thrust. A field emission cathode (FEC) is one good option for the electron emitter owing to its simplicity and potential performance. The FEC used in this study is comprises an emitter electrode with a carbon nanotube (CNT) coating and a gate electrode as the extraction electrode. In these EDT systems, it is expected that several FEC units will operate in parallel to fulfill the redundancy requirement. Since interactions between cathode units may cause instability or performance degradation, parallel operation experiments must be performed on the FECs prior to practical on-orbit operation. We conducted experiments on FEC parallel and single operations in both vacuum and plasma environments. Consequently, we found that the following control method for the FEC is effective for obtaining a maximum emission current and minimizing the gate current; the gate voltage is controlled in response to the change in the emitter potential by configuring the upper limit of the emission and gate currents. However, the emission currents during parallel operations of the FECs were still lower than those during operations of single FECs. The electron current passing through the gate in the plasma environment was found to be lower than that in the vacuum as the positive potential of the gate attracts electrons from the plasma.

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With over 272 attempted launches since 2000, CubeSat technology has exponentially increased as industries and universities have realized their potential. While this growth looks promising for space research possibilities, there are still a number of issues, with the largest being CubeSat maneuverability. The majority of CubeSats cannot orient or propel themselves, meaning mission functionality is limited and collision probability will increase as time goes on. CubeSat technology has been improving, and the mission of this technology has become increasingly more important in the development and advancement of new technologies. The Micro-propulsion and Nanotechnology Laboratory at The George Washington University has constructed a four-channel Micro-Cathode Arc Thruster (μCAT) micro-propulsion subsystem that allows these satellites to perform missions without reliance on their launch vehicles. The propulsion system has a volume of approximately 541 cm³ that can produce specific impulse values up to 3000 s. Each μCAT onboard is used for the CubeSat's attitude control, orbit change, de-orbiting, and movement. The μCAT system was integrated into the USNA's 1.5U CubeSat (BRICSat-P) to be used to perform three maneuvers while at an orbit of 500 km: de-tumbling, spin, and a deltaV that will attempt to change the orbit of the CubeSat relative to the orientation of Earth's magnetic field. The objective of this paper is to provide an overview of the thruster subsystem’s development and application for the BRICSat-P mission parameters. In addition, the μCAT subsystem’s circuitry, thruster head design, and development will be reviewed to provide the information used to reach CubeSat flight standards.

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High discharge voltage operation of a conceptual low-erosion magnetic layer type Hall thruster was modeled by an axial-radial two dimensional fully kinetic particle simulation code. No anomalous diffusion model was used to self-consistently capture the physics of electron cross-field transport and wall erosion of the thruster. For the verification of the size of the computational domain, simulation was conducted on two computational domains with different radial size. It was shown that the size of the computational domain had significant impact (>30%) on wall erosion prediction, though its influence on thrust performance was minor (<3%). The performance of the designed conceptual thruster was predicted to be >50% in anode efficiency and <0.3 mm/kh wall erosion rate inside the discharge channel. Furthermore, >1.5 mm/kh erosion rate was observed for the front wall, which was consistent with the experimentally observed pole-piece erosion of “Magnetic Shielding” Hall thrusters. The cause of the front wall erosion was shown to be the ions produced outside the channel, and the plasma potential structure leading to the ion acceleration toward the front wall.

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An innovative engine, designated “Volterra,” has been developed and the thrust performance, thrust, beam divergence angle, and ion energy distribution function were investigated using a 1 kW class magnetic-layer-type Hall thruster developed at Kyushu University. The thrust of this engine is superior to that of the thruster with 150 V constant voltage operation, but the ion energy distribution function is wider. Plume divergence is almost the same as for the thruster with 150 V constant voltage operation.

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This paper focuses on the variation in extraction-ion beam performances in miniature ion engines due to grid erosion. Because speed and cost are very important issues in microspacecraft development, we replace time-consuming grid wear tests with numerical simulations using the JAXA Ion Engine Development Initiatives (JIEDI) tool. By comparing the results obtained for a miniature ion engine with those of a μ10 ion engine, we observe significant extraction-ion beam performance degradation in the miniature ion engine. Because miniature ion engines tend to omit gimballing devices to save their masses, we also present a numerical simulation for predicting erosion-induced movement in the thrust-vector direction for the estimation of the accumulation of thrust-vector misalignment toque, which must be unloaded by firing low-Isp chemical thrusters. Using the JIEDI tool, we can successfully calculate its movement by grid erosion and obtain the averaged thrust-vector direction during the operation of the ion engine. This time-averaged thrust-vector direction provides the best initial thruster setting angle for each miniature ion engine.

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As a part of a Japanese collaborative research and development project on practical use of a high power anode layer type Hall thruster, a 5 kW class anode layer Hall thruster (RAIJIN94) has been developed and the thrust performance has been evaluated. The thrust was measured in the ion engine endurance test facility at ISAS/JAXA using a pendulum thrust stand developed at the University of Tokyo. The thrust performance at 3 kW operation was measured (xenon anode mass flow rate of 9.8 mg/s and xenon cathode mass flow rate of 0.5 mg/s); the thrust, specific impulse, and thrust efficiency were found to be 160 mN, 1600 sec and 0.42, respectively. The thrust performance depends on magnetic field configuration, that is, the strength of the magnetic field and the ratio of trim coil to inner/outer coil.

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In order to understand plasma properties of hollow cathodes, a numerical simulation code with Hybrid-PIC model has been developed, in which ions and electrons are modeled as particles and fluid, respectively. In this study, as a first step, the applicability of the model is demonstrated, and then the influences of the emitter temperature on the flow field are discussed for a discharge current of 30 A and a mass flow rate of 1 mg/s. The electron density for the maximum emitter temperature of 1900 K agrees well with the experimental data from JPL. The results also show that the electron density tends to be higher with lower emitter temperature due to the higher electron temperature inside the cathode tube. The higher electron temperature is caused by the energy loss suppression resulting from the higher sheath voltage on the emitter surface. It was also found that charge exchange collisions shift the location of the electron density peak upstream.

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A Hall effect thruster (HET) incorporates a magnetic circuit that must generate a specific flux density spatial distribution inside and near the outlets of the plasma channel. The first objective of the design process of this type of structure is to obtain a specific magnetic filed topology in the thruster channel by fixing radial and axial components of the magnetic field and a certain inclination of magnetic flux lines. The aim of this work is to develop a tool for solving this inverse magneto static problem, which is applied to the SNECMA PPS1350 ^® HET magnetic circuit in order to obtain a new “Low-Erosion” magnetic configuration.

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Conventionally, neutralizers in ion thruster systems do not generate thrust force. Hence, the power consumption of a neutralizer negatively affects the thrust efficiency of the ion thruster system. Therefore, in this paper, a negative ion source that generates thrust force as well as neutralizes the positive ion beam was newly developed using fullerene as a propellant so as to realize a more efficient ion thruster system. To develop the negative ion source, two measurements were conducted. The first measurement was an E × B probe to identify the species of positive and negative ions. The second measurement was a magnetically filtered Faraday probe to measure quantitatively the negative ion currents. Based on the measurements, it is concluded that the negative current is not carried by electrons but by negatively charged fullerenes. Finally, the negative ion source was successfully coupled with a positive ion source. To the best of our knowledge, this is the first paper to report the demonstration of an ion thruster using a negative ion source instead of a cathode.

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A particle model of plasma was coupled with electromagnetic wave propagation to reproduce a filamentary structure and diffusive pattern in a microwave breakdown. The thrust performance of a microwave rocket was estimated by combining discharge calculation and computational fluid dynamics (CFD). An external magnetic field is applied to the discharge volume to improve thrust performance for low ambient pressure by combination of magnetic confinement and resonance heating technique. The magnetic field suppresses the propagation speed of the ionization front because the electron cannot cross the magnetic field at lower pressure, whereas propagation speed increases with decreased ambient pressure when no magnetic field is applied. The energy absorption rate increases when electron cyclotron resonance (ECR) occurs with the corresponding magnetic field, which improves the thrust performance of the microwave rocket.

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The present work describes key activities in the development campaign of the CAM200 low power Hall thruster from initial prototype testing to concurrent full performance mapping. The development program presented included proof-of-concept tests, experimental and numerical validation of physical mechanisms, wall material selection, performance testing, thruster engineering model structural, thermal and magnetic simulations followed by an engineering model production as well as full performance mapping. During the development campaign CAM200 demonstrated exceptional performance with anode specific impulse and anode efficiencies above 1500 sec and 43%, respectively, at a power level of 250 W. Future activities include thruster-cathode coupling test, full thruster unit lifetime qualification and full electric propulsion system coupling test, all part of the MEPS project, a joint European-Israeli endeavor.

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The European Commission Horizon 2020 funded DEMOCRITOS project (2015-2017, see under democritos.esf.org) will be primary focused to prepare preliminary design of the ground, core and space demonstrators and their test benches for the mega-watt class nuclear electric space propulsion INPPS flagship (International Nuclear Power and Propulsion System). In addition programmatic, organizational and funding aspects for international cooperation related to INPPS realization are sketched. The new project includes partners from Europe, Russia and the Brazilian guest observer IEAv and is the follow-up of the mega-watt class nuclear electric propulsion European-Russian MEGAHIT project (www.megahit-eu.org). Europe already established the high power nuclear MEGAHIT and the low power nuclear (20 to 200 kW NEP) DiPoP (www.DiPoP.eu) roadmaps. Because Europe has started the implementations for INPPS flagship in the 2030-2040 timeframe, both roadmaps will be also described – from MEGAHIT the INPPS technology options, the launcher, assembly and system architecture, space mission requirements, communications and public support. In case of DiPoP it will be explained the survey of European capabilities, technical options, potential space missions and the public acceptance as well.

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We present an electrostatic Particle-in-Cell simulation of a downscaled High Efficiency Multistage Plasma Thruster (HEMPT). The purpose of downscaling the HEMPT design is to reach the requirements of missions which have a need for low thrust (0.1...150 μN) and low noise (root of the noise spectral density ≤ 0.1 μN/√Hz). These are upcoming formation flying space missions like eLISA (evolved Laser Interferometer Space Antenna) or NGGM (Next Generation Gravity Mission). The aim of the here presented numerical simulations is to get an improved understanding of the thruster's physics especially in its downscaled configuration, in order to reach the design goals.

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The Inflatable Space Terrarium (IST) was launched as one of the experimental equipment of the Space Inflatable Membrane Structures Pioneering Long-term Experiments (SIMPLE) module attached to the Exposed Facility of the Japanese Experimental Module (JEM, Kibo). It aimed to verify the retainment of the pressurized membrane structure in which the atmospheric environment is simulated. It was deployed in orbit successfully on August 18, 2012 (JST), however, a slight gas leak then occurred and the pressure inside the IST has gradually decreased owing to this leak. The SIMPLE mission ended its regular operation in orbit at November 2014. Although the criteria for containment remained unattained owing to the leakage, we had monitored the conditions inside the IST by using cameras, and pressure and temperature sensors for about two years after deployment. In this paper, based on the pressure reduction profile we estimate size of a defect on the assumption that the air leakage occurred through one small orifice. And we analyze the relation between pressure and temperature using on-orbit data and numerical results.

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This study proposes a finite element updating method using multiobjective optimization to consider multiple experimental conditions for estimating parameters. The method aims to minimize the root-mean-square (RMS) error of the deformation shape between the finite element analysis and experimental results. The proposed method is applied to the bread board model (BBM) of a tensionstabilized space reflector consisting of hoop cables and radial ribs, in which the rib is deformed into the prescribed shape by the cable tensions generated on deployment. The design requirement is to deform the rib into the prescribed shape by applying appropriate tension loads to the radial and hoop cables. Under actual conditions, the deformation shape deviates from the ideal shape because of uncertainties. Therefore, it is necessary to estimate the physical parameters with high accuracy, through a geometrically nonlinear finite element analysis, in order to investigate their effect on the deformation shape. To efficiently estimate the physical parameters, the satisficing trade-offmethod (STOM) is adopted as the multiobjective optimization method. Through numerical examples, the validity of the proposed method is demonstrated by comparing the analytical deformation shapes with experimental results.

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An extension demonstration of a space inflatable mast and subsequent long-term operation for 845 days on orbit were performed to show the advanced design technology and practical utility of space inflatable structures. The experiment verified the long-term availability of the inflatable extension mast and good establishment of the design technology, including development testing know-how for space-borne inflatable structures, by considering possible degradation effects simulated using a finite-element method analysis tool developed in parallel with the mission. The space inflatable extension mast achieved the extra success level set out in the mission plan.

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Piezoelectric ceramic actuators with less thermal strain were designed by symmetrically stacking two kinds of piezoelectric ceramics plates whose linear expansion coefficients were positive and negative, and their performance was evaluated. Three types of actuators were examined, which have a different stacking direction. The electromechanical material constants were evaluated by a one-dimensional model and the results were compared with a three-dimensional finite element analysis. The result showed that the piezoelectric constant of an actuator stacked in the perpendicular direction to the electric field was about three times larger than that of an actuator stacked in the parallel direction, even though the volume ratio between the two kinds of plates was almost the same between the stacking types. This is because most of the voltage in the latter type was distributed to the piezoelectric plate with low permittivity and low piezoelectric constant. Therefore, it is important to consider the stacking direction and the permittivity in the design of the actuators. The three-dimensional finite element analysis showed the linear expansion coefficient was not zero even though it was zero in the one-dimensional model. Therefore, three-dimensional analysis is necessary for precise design, although the one-dimensional model provides good estimation of the electromechanical property.

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This study clarifies the effect of thermal deformation in the reconfigurable space reflector system, through a series of experiments in thermostatic chamber using a prototype of reconfigurable space reflectors. This study then discusses possible countermeasures against precision degradation of such system due to the thermal deformation, exploiting a linkage model and a finite element method model. Though reconfigurable reflector systems have the mismatch in the coefficient of thermal expansion among their structural members, the effect of the mismatch may be minimized by using monometallic mechanism design, and a proper combination of several materials.

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Air density on Mars is much lower than that on Earth. To generate sufficient lifting force to fly, Mars-airplanes need to have a larger wing area than Earth-airplanes. The recently developed Mars-airplanes have multibody wings that can be folded and deployed to realize larger wing area and compactness. Aeroelastic analyses of the wings are necessary to avoid catastrophic behaviors, such as flutter or divergence. However, conventional aeroelastic analysis methods cannot be applied to the multibody wing because these wings have mechanical joints for connecting wing bodies, and thus, they differ significantly from conventional wings. In this paper, a new analysis method that can be applied to the multibody wing is explained. The method combines aerodynamics, multibody dynamics theory, and absolute nodal coordinate formulation. By using this method, we simulate the aeroelastic motion of multibody wings. We investigate the changes in aeroelastic motion when we change the number of the wing bodies and the structural parameters.

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An adaptive estimation method for nonlinear structural deformations is presented. The method is based on the ensemble Kalman filter (EnKF), which can effectively handle nonlinearities in structural models by using the Monte-Carlo simulation. In this study, a self-tuning algorithm for the system noise in the conventional Kalman filter is extended and applied for tuning in the EnKF. To verify the effectiveness of the presented method, a numerical experiment was performed for a deployable frame structure system that contains the typical nonlinearities of space deployable structures, that is, the geometrical nonlinearity in flexible members and the cable nonlinearity resulting from its slackened state.

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We investigate the effects of covering pressure and friction between convex tapes on the bending stiffness and natural frequency of a braid-coated biconvex tape (BCON) boom. We establish an analytical model by comparing experimental and analytical results of BCON booms because the braid-coated boundary conditions of the boom are complicated and the analytical model is difficult to construct. BCON booms will have various uses for regular and large-scale deployable space structures such as solar sails, solar array panels, and de-orbit mechanisms. Thus, the bending stiffness, natural frequency, and structural characteristics after deployment are experimentally measured. Several parametric analyses are also calculated by using the proposed contact analysis model. On the basis of these results, we discuss the effect of the braid coating and future work.

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This paper proposes improved methods for remotely operating orbital structures comprising connected flexible plates using real-time estimation. First, a hinge-connected plate model is introduced and a method of conducting estimation of dynamic behavior for operational spacecraft is defined. Then, a model-reduction method, which can be applied to targets with large-deformation behavior, is discussed. The proposed method is compared with the conventional one in terms of calculation cost and approximate accuracy. Next, an online parameter-identification method, that deals with the amplitude dependence of orbital flexible structures, is explained. A random decrement method is introduced to extract the natural vibration from the time history under the assumption of a random or environmental noise force. Using the extracted history, dynamic parameters are identified using an eigen system realization algorithm. Finally, these improved methods are introduced to the estimation system and a verification experiment is conducted using a ground experiment set that is representative of an extendable plate structure.

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In order to predict the sail membrane dynamics precisely, it is necessary to use the on-orbit measurement data. However, the on-orbit sail membrane measurement data are incomplete (missing spatio-temporal data) because of sensor placement limitations and sunlight reflection. This paper proposes an empirical data driven model in which past measurement data and current incomplete measurement data are used to estimate sail membrane dynamics and the confidence intervals of the estimated dynamics are obtained via a bootstrap method. Further, the applicability of the proposed empirical data driven model is evaluated via numerical and vacuum chamber experiments. The experimental model used comprises a spin deployable space membrane structure consisting of a square membrane, a center hub, and tethers.

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A quaternion representation is often used to describe the attitude of a spacecraft because it does not have any singular points. However, it becomes difficult to control the attitude described by a quaternion since a quaternion has four parameters despite that the attitude has only three degrees of freedom. In this paper, we employ the concept of port-Hamiltonian modeling to control systems with quaternions to introduce a general nonlinear control system synthesis method in aerospace engineering. It is also shown that the error quaternions are also naturally described by the port-Hamiltonian framework. Furthermore, the additional design parameter achieved by the proposed method is utilized for obstacle avoidance control. A numerical example exhibits the effectiveness of the proposed method.

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This paper presents attitude maneuver demonstration tests using control moment gyroscopes mounted in a three-axis free dynamics simulator that provides an on-ground test environment of free rotational movement around three axes using the air-floating method. The issue of torque error between the torque command and the actual output torque is discussed in relation to the singularity-robust based steering law for spacecraft maneuvers. The torque error is analyzed using singular value decomposition to determine its magnitude and direction. Based on the analysis, feedback control and steering laws are proposed to reduce the torque error. Before conducting the maneuver demonstration test, a preliminary simulation is performed using a simplified dynamics model, control law, and steering law. The maneuver test results indicate that the dynamics simulator successfully accomplishes agile, large angle, and rest-to-rest multitarget maneuvers as well as precision pointing.

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Satellite dynamics is described by a nonlinear differential equation. Most of recent studies about attitude control have used non-linear controllers. However, with these controllers, control performance is ignored in most cases. To overcome these problems, we applied Linear Parameter-Varying (LPV) control theory to attitude control problem. To avoid difficulties coming from the nonlinearity in satellite dynamics, we modeled dynamics of spacecraft as an LPV system and applied a Gain-Scheduled (GS) controller to this model using Linear Matrix Inequalities (LMIs). In this paper, by using two methods, GS controllers are designed to guarantee overall stability and to achieve H₂ performance with distinct Lyapunov solutions. By using these controllers, 3-axis attitude control of a spacecraft with Reaction Wheels (RWs) shall be achieved. To examine how the proposed approach improves the control performance, two proposed methods shall be compared with each other.

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This paper proposes a new guidance and control system for a re-entry vehicle. The vehicle flying through an extensive flight region can be expressed as a nonlinear system which has uncertain dynamic characteristics and physical constraints. In addition, to address considerations of mission abort and changes of landing site during a space transportation mission, sequentially generating flight trajectories from the guidance system is desirable. Therefore, in this study, we attempt to derive an online trajectory generation system by solving a two-point boundary value problem using the concept of flatness as an extension of exact linearization. Moreover, an attitude control system is designed to track a command signal corresponding to a generated trajectory. The system is employed as a constrained adaptive backstepping control method that can accommodate uncertain dynamic characteristics and physical constraints of a re-entry vehicle. The numerical simulations verify the effectiveness of the proposed guidance and control system.

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This paper considers line-of-sight (LOS) maneuver control of an underactuated spacecraft equipped with two skewed singlegimbal control moment gyros. To describe the spacecraft attitude, two parameters referred to as the W-Z parameters are used, the first of which represents the angle between the target LOS and the LOS of the mission-equipment, and the second of which represents the angle around the LOS. In order to stabilize the mission-equipment LOS to the target LOS, a two-step attitude control procedure is considered, which consists of feedforward control combined with feedback control based on a backstepping control method using the W-Z parameters. A numerical simulation is carried out to validate the proposed control procedure.

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Effective computing methods are proposed for calculating the sensitivity matrices related to empirical acceleration parameters that are used to compensate for dynamic modeling deficiencies. Meanwhile, design and normal equation matrices are expressed in the form of partitioned sub-matrices. These methods of computing the sensitivity matrices and partitioning the normal equation matrices are extremely timesaving and require less storage, compared to conventional methods. The GRACE real flight GPS observation data has been used to evaluate the computing performance of setting up normal equation. The efficiency increased to 2 times when using code observations only, and increased to 3.5 times when using code and carrier phase observations together. Furthermore, we found that the empirical accelerations showed quasi-periodic characteristics with respect to time variable, which allowed Fourier series to be used to interpolate empirical accelerations. The fitted tangential empirical acceleration curves, which are added into deterministic force models, form enhanced-accuracy dynamic models that are applied to orbit prediction. For 3-day orbit pass, the prediction accuracy was better than 60 meters and an improvement of 2.3 times on average was achieved, compared to conventional dynamic models without considering empirical acceleration models. The proposed prediction schemes are beneficial to the establishment of advanced, possibly onboard, satellite navigation systems.

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A three-axis attitude control function module was assembled into a highly integrated cubic package (10³ cm³ dimension: 1U-sized). The key feature of the cubic module is its high-agility functionality, which is realized by a novel, compact electromagnetic braking mechanism combined with a ferromagnetic wheel. In this study, the proposed concept is compared with existing 1-axis wheel assemblies and the conventional CMG scheme in terms of momentum storage per mass/volume, maneuvering speed, and compactness. A prototype of the module was manufactured to demonstrate this innovative concept.

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Singularity is a problem in attitude control of a spacecraft using control moment gyros (CMGs). In this study, the internal singular states of two six-CMG systems; the hexagonal and twin-triangular type CMGs are investigated and one of the CMG arrangement types; the twin triangular type is proven to have only passable internal singular states. By using the passability, the gimbal angle calculation that satisfies the given angular momentum of CMGs is proposed and its effectiveness is compared with an existing method. A steering law for the gimbal angular velocities based on the gimbal angle calculation is also constructed. The attitude control of the spacecraft by using the proposed steering law is examined through numerical simulations.

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This paper discusses the performance of Wide-Field-Integration (WFI) of optic flow when it is applied to a guidance and navigation system for space probes near asteroid surface. WFI of optic flow is a state estimation method inspired by the visual processing system of compound eyes of flying insects. This method has attractive features: low computational cost, applicableness to a lowresolution image sensor, and robustness for unknown environments. Unlike the theory of WFI of optic flow, the region of an image sensor is limited in a real system. In this paper, the effects of the directions of optical axes on estimation accuracy are investigated in numerical simulations and experiments. The evaluation results in simulations indicate the proper directions of optic flow sensors. The results also indicate that under a limited field of view, one optic flow sensor has frequently difficulty to distinguish motion variables from WFI of optic flow. Finally, the results obtained in numerical simulations are examined in experiments by using two image sensors.

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New artificial Lagrange points for spacecraft with a tethered anchor are presented in this paper. The positions of conventional artificial Lagrange points generated by solar radiation pressure are restricted depending on the sail lightness number. Using a tethered anchor, that restriction can be relaxed. Around the L2 point, the new artificial Lagrange points for spacecraft with a tethered anchor are derived based on the circular restricted three-body problem. The dependence of the position of the new artificial Lagrange point on the parameters such as mass ratio and tether’s length is investigated for spacecraft with a tethered anchor. The trajectories of spacecraft with and without tethered anchors are calculated numerically. Results show that the tethered anchor can halve the offset of artificial Lagrange points. The effect of tethered anchor on the shift of artificial Lagrange point increases if a tether is long and the mass ratio m_anc /m_sc is large.

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In this study, the performance of rate damping control from a control moment gyro is investigated. Two types of wheel angular momentum (variable and constant) and two quantities of gimbal axes (one or two) are considered. When the angular momentum of the spacecraft body is fully absorbed by the wheel, the rate damping of the spacecraft can be achieved. In contrast, for cases of with constant wheel speed, the rate damping of the spacecraft body cannot be achieved. The steady state analysis of the rate damping control in these cases shows the stationary value of the angular velocity of the spacecraft. Numerical simulations verify the analytical results.

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Motion around small asteroids, given their weak gravity fields, is strongly affected by perturbation forces such as solar radiation pressure. Several types of orbits in such strongly perturbed environment have been proposed in past studies. However, these orbits usually have complex shapes and limited orbital geometries. To avoid these disadvantages, delta-V assisted periodic orbits (DVAPOs), which are made periodic by introducing a small impulsive delta-V within each period, are proposed. Stability analysis of DVAPOs is the main topic of this paper. The stability of DVAPOs can be analyzed by treating delta-V maneuvers as part of the natural motion and extending the conventional stability analysis method for natural periodic orbits. Moreover, the stability of feed-back controlled DVAPOs can also be discussed by introducing an augmented monodromy matrix. Based on an augmented monodromy matrix, stabilization strategies for DVAPOs are successfully established. Finally, the feasibility of a DVAPO for the Hayabusa 2 mission is shown.

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This paper considers position and attitude control of large flexible space structures composed of a number of subsystems(substructures) under the assumption of sensors and actuators collocation. The purpose of this paper is to propose a decentralized control scheme with fourth order local proper controllers using displacement/angle output, which makes both each closed-loop subsystem and an overall closed-loop system not only robustly stable but also optimal for quadratic cost functions. We first introduce a fourth order local proper controller for stabilizing each subsystem using only displacement/angle output. Then, we show each closed-loop subsystem becomes not only robustly stable against uncertain characteristic parameters such as mass, damping, and stiffness, but also optimal for a quadratic cost function by choosing parameters of each local proper controller appropriately. After stabilizing and optimizing each subsystem, we interconnect the closed-loop subsystems by flexible links to obtain an overall closed-loop system, and we show the overall closed-loop system also becomes robustly stable and optimal for a quadratic cost function. Finally, numerical examples are presented to show effectiveness of the proposed method.

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This paper presents an analytical formulation and a computational method of free-molecular flow effects incorporating the temperature variation over the orbital motion of a satellite. The rarefied aerodynamics interactions, so-called free-molecular flow, has been studied for decades and the complexity of the phenomena has been the bottleneck. A new aspect of analysis on the surface temperature contributions is implemented in the analytical form. It affects the magnitude of momentum exchanges and thus the resulting force at the satellite surfaces. The free-molecular interaction perturbs the orbit and the attitude of a satellite as well as of any other orbiting space objects below the exosphere of the Earth, including International Space Station and space debris. The enhanced fidelity and accuracy of the atmospheric perturbation model is expected to complement with the growing requirements on the orbit determination accuracy of space objects, especially gravimetry missions. Under exemplary orbit and attitude conditions, temperature variations together with the free-molecular flow effects are obtained and discussed. The result indicates the magnitude of the force coefficient jump is around 6%, which is induced by the surface temperature variation.

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In this paper, a real-time capable nonlinear model predictive controller is implemented for the attitude control of an upper stage launch vehicle with liquid propellant. A mass spring model is used as an analogy to simulate the disturbance generated by the sloshing propellant. For the implementation of the nonlinear model predictive controller, an optimal control problem is defined with finite time horizon. The objective function is minimized while satisfying constraints on the control inputs. The resulting optimal control problem is transcribed using single shooting method to parametrize the control inputs using uniform discretization points. The continuous control inputs are obtained by linear interpolation. A dedicated discretization algorithm in FORTRAN is coupled with a solver which used quasi-Newton algorithm to generate solutions fast. Approximation of the Hessian matrix is used to reduce computational requirements. Furthermore, the algorithm can perform parallel computation of the derivatives of the objective function with respect to optimization variables. This results in a real-time capability of generating solutions in the order of hundred milliseconds for each iteration. The algorithm is applied for attitude maneuver and disturbance rejection for the upper stage of a launch vehicle.

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A method for controlling a Coulomb force attractor spacecraft in the vicinity of an asteroid is presented. A Coulomb force attractor tows and deflects an asteroid through a combination of mutual gravitational and Coulomb forces. We show asteroid deflection distances with time before impact and the required fuel consumption for efficient mission design with limited resources. By considering the asteroid and the spacecraft as a single body, motion is represented with the separation distance between the spacecraft and the asteroid and two Eulerian angles. We also investigate linearized dynamics and identify the stability requirements using the Routh–Hurwiz stability criterion. Numerical simulations are also performed and the feedback law to stabilize the position of the spacecraft is investigated. By investigating the interaction between the separation distance and Eulerian angles, we propose and evaluate a method for independently controlling each motion.

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JAXA has developed a Jovian Trojan asteroid sample return mission using a solar power sail. Jovian Trojan asteroids are among the few remaining frontiers in our solar system and may hold clues to its formation and evolution. However, visiting Jovian Trojans is much more difficult than reaching near-Earth objects because of the large amount of fuel required to reach them. Moreover, large distance from the sun makes power generation difficult. Solar power sails offer an effective way of realizing such challenging exploration. This paper outlines a solar power sail spacecraft and discusses the design of a trajectory for a sample return mission to a Jovian Trojan asteroid. The time of flight is long, but a large payload can be delivered to the asteroid by using a solar power sail. Reducing the duration of a sample return mission is difficult, but it is possible for a one-way mission. This paper presents a trajectory design for such a one-way mission.

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The current paper presents a concept of the non-conventional method of handling/transfer of the space objects/payloads, using single and multiple cooperative robotic manipulators with extremely flexible arms. Practical cases may involve launch only (as in the case of the ejection of the spacecraft), capture only (as in the case of the capture of the dysfunctional spacecraft or debris) or launch and capture, combined together (as in the case of a payload transfer from one platform to another). This method is seen as having numerous advantages over the traditional methods, employing propulsive systems on the payloads. These include: simplicity,reliability, re-usability, ability to handle passive payloads, and small power requirements. In this study, various operational scenarios of the launches and captures of the payloads are considered and designed in detail. In particular, we demonstrate the feasibility of the launch of the outgoing rotating payloads (called "frisbees") using a highly elastic robotic arm, capable of transferring its potential energy of the pre-deformed shape into the kinetic energy of the payload. Using the co-rotational FEM, we firstly simulate pre-launch phases, coiling the elastic elements in different shapes, for example, "U", "S" or even more complex shapes, and then propose and simulate the scenarios of ejecting the payloads utilising the kinetic energy of the elastic members (playing role of catapults). We also demonstrate the feasibility of the ejection of the payloads with required dynamic parameters, using cooperative robotic manipulator arms, performing coordinated throwing manoeuvres. This is done in compliance with the kinematics of the system and dynamics laws. The cases of the ejection of the payload is simulated using fully non-linear formulation, employing so-called co-rotational FEM,which enables to deal with large deformations, large rotations and large translations of the simulated elastic robotic arms. Numerical simulation allows to observe, to analyse and to suppress the transient strains and stresses in the flexible arms. In the cases of the payload capture, the co-rotational FEM method is used to simulate the process of de-spin, slowing and stopping the spinning object. This is also supplemented with the analysis of the strains and stresses on the members of the robotic arm manipulators. Various study cases are illustrated with the animations of the representative cases in Virtual Reality.

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This paper presents relative position control of spacecraft formation flying. A control law based on the Tschauner-Hempel (TH) equation is usually used to reconfigure a spacecraft formation. However, the solution of the TH equation is not valid when the formation size is large. This paper focuses on the effects of the relatively long distance between two spacecraft. By taking the differences in orbital elements between two spacecraft as state variables, a state equation is obtained with long-distance effects as perturbations. Then, the optimal control input is derived by the Hamiltonian perturbation theory, including the long-distance effects. Numerical simulations show the usefulness of the control input from the perspective of the accuracy of the relative position and velocity. The obtained control input is also compared with the optimal control input numerically obtained by nonlinear programming.

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Lagrangian points are libration points in the circular restricted three-body problem. Three of them are collinear libration points, and known to have unstable periodic and quasi-periodic orbits in the vicinity of themselves. This study aims to design periodic and quasi-periodic orbits from the center manifold theory standpoint. In this paper, a new approach is proposed to design periodic and quasi-periodic orbits around the collinear libration points based upon the successive approximate method. First, the verification of the proposed approach is discussed by evaluating obtained orbits. Then the relation between the initial condition and the type of resulting orbits is investigated. In conclusion, this paper reveals that the proposed method significantly reduces the effort to obtain bounded orbits in the circular restricted three-body problem. The proposed method needs neither an initial guess for the periodic orbit nor the complex algebraic manipulation unlike the conventional methods designing periodic or quasi-periodic orbits. Moreover, a guide to choose the initial condition to obtain desired orbits is given.

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The purpose of this study is to design orbits for transporting passengers and cargo regularly from the Earth to the Moon and back. To accomplish this, so-called “cycler orbits” are used. A cycler orbit is an orbit circulating regularly between two astronomical bodies and can be achieved through gravity-assist swingbys. This study consists of two steps. The first step exploits the mechanics of the “double lunar swingby” to realize a cycler orbit without expending propellant. The second step employs the primer vector theory to design low-cost transfer orbits and calculates the minimum fuel for the mission. The proposed method of this study can realize an Earth-Moon transportation system which operates every two months with a significantly lower cost than the previous method.

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The purpose of this work was to develop escape trajectories from distant retrograde orbits in the Sun–Earth circular restricted three-body problem. Previous studies have suggested installing a space port outside the gravitational sphere of influence of Earth for future deep space exploration. This paper proposes the placement of a space port on a Sun–Earth distant retrograde orbit (SE-DRO), which is stable over a long period. The characteristics and fuel consumption efficiency of escape trajectories from the space port on the SE-DRO using one or two impulses ΔV were investigated and compared with those obtained in previous studies that considered a space port located in the vicinity of the Sun–Earth L2 Lagrangian point. The connection of the escape trajectories from an SE-DRO with Earth gravity-assist (EGA) trajectories was then investigated based on the results of the analysis with the impulse ΔV. Finally, a series of trajectories from SE-DRO to EGA were shown to have a high efficiency from the perspective of both ΔV and the flight time.

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Station-keeping maneuvers are necessary for the precise formation flying of spacecraft, because various disturbing forces deviate their relative motion. The objective of this study is to derive a two-impulsive control law for the formation flying of two spacecraft in an orbit with a small eccentricity under the J₂ perturbation, especially one that suppresses the relative position deviation during the control period. Because of the J₂ perturbation, the relative position and velocity have secular terms. However, it is possible to compensate for the secular terms of the relative motion and create an artificial periodic motion by conducting the impulsive control. In this study, a state transition matrix of the relative motion based on the osculating orbital elements with the first order of J₂ is utilized, in order to calculate the secular and periodic terms of the relative motion, to derive an impulsive control law, and to determine the appropriate initial conditions for suppressing the deviation of formation flying. The numerical results of the proposed control law are compared with those of another method, which is based on the relative mean orbital elements, and it is shown that the proposed control law successfully suppresses the formation deviation.

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