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EQUULEUS (EQUilibriUm Lunar-Earth point 6U Spacecraft) will be the world's smallest spacecraft to explore the Earth-Moon Lagrange point. It is being jointly developed by JAXA (Japan Aerospace Exploration Agency) and the University of Tokyo, and will be launched by NASA's Space Launch System Exploration Mission-1. The spacecraft will fly to a libration orbit around the Earth-Moon L2 point (EML2) and will demonstrate low-energy trajectory-control techniques within the Sun-Earth-Moon region for the first time by a nano-class spacecraft. EQUULEUS also carries three scientific observation missions: imaging of Earth's plasmasphere by extreme ultraviolet wavelength, lunar impact flash observation on the far side of the moon, and micrometeoroid flux measurements in the cis-lunar region. While all these missions have their own scientific objectives, they will also contribute to future human activity and/or infrastructure development in the cis-lunar region. Most parts of the spacecraft system use commercial off-the-shelf components, or are designed based on the experiences of various past space missions, with the exception of the newly developed water resistojet propulsion system. EQUULEUS uses X-band frequency for deep space telecommunication. Japanese deep space antennas (64-m and 34-m) will be nominally used for spacecraft operation, and support from the deep space network of JPL (Jet Propulsion Laboratory) is also being planned, especially for the initial phase of operation. The spacecraft will fly to EML2 in less than one year, and will remain there for scientific observations until shortly before the depletion of the onboard propellant, when the spacecraft will leave the orbit for space-debris compliance.

Recent low-thrust space missions have highlighted the importance of designing trajectories that are robust against uncertainties. In its complete form, this process is formulated as a nonlinear constrained stochastic optimal control problem. This problem is among the most complex in control theory, and no practically applicable method to low-thrust trajectory optimization problems has been proposed to date. This paper presents a new algorithm to solve stochastic optimal control problems with nonlinear systems and constraints. The proposed algorithm uses the unscented transform to convert a stochastic optimal control problem into a deterministic problem, which is then solved by trajectory optimization methods such as differential dynamic programming. Two numerical examples, one of which applies the proposed method to low-thrust trajectory design, illustrate that it automatically introduces margins that improve robustness. Finally, Monte Carlo simulations are used to evaluate the robustness and optimality of the solution.

The development of a robust navigation infrastructure in cis-lunar space is crucial for the coming new era of advanced lunar exploration. NASA has set a goal to develop an extensible and scalable lunar communication and navigation architecture called LunaNet. For the flexibility and robustness of the system, the architecture is desired to have an autonomous and decentralized operation capability. This paper presents and demonstrates a decentralized and autonomous state estimation algorithm for SmallSats that provides positioning, navigation, and timing service, each equipped with a GNSS receiver, chip-scale atomic clock, and inter-satellite communication module. In the framework, each satellite individually estimates its own state and clock offset with a decentralized Schmidt Extended Kalman Filter by processing weak GNSS signal and inter-satellite range measurements. The algorithm is validated with Monte-Carlo state estimate simulations of five satellites at the hybrid constellation of lunar frozen orbit and the Near-Rectilinear Halo Orbit (NRHO). The proposed filter converged in all Monte-Carlo cases with clock offset errors below 0.15 microseconds and 1.0 microseconds in the frozen orbit and the NRHO, respectively.

This paper presents the trajectory design for EQUilibriUm Lunar-Earth point 6U Spacecraft (EQUULEUS), which aims to demonstrate orbit control capability of CubeSats in the cislunar space. The mission plans to observe the far side of the Moon from an Earth-Moon L2 (EML2) libration point orbit. The EQUULEUS trajectory design needs to react to uncertainties of mission design parameters such as the launch conditions, errors, and thrust levels. The main challenge is to quickly design science orbits at EML2 and low-energy transfers from the post-deployment trajectory to the science orbits within the CubeSat’s limited propulsion capabilities. To overcome this challenge, we develop a systematic trajectory design approach that 1) designs over 13,000 EML2 quasi-halo orbits in a full-ephemeris model with a statistical stationkeeping cost evaluation, and 2) identifies families of low-energy transfers to the science orbits using lunar flybys and solar perturbations. The approach is successfully applied for the trajectory design of EQUULEUS.

GEO-X (GEOspace X-ray imager) is a 50 kg-class small satellite to image the global Earth’s magnetosphere in X-rays via solar wind charge exchange emission. A 12U CubeSat will be injected into an elliptical orbit with an apogee distance of ∼40 Earth radii. In order to observe the diffuse soft X-ray emission in 0.3-2 keV and to verify X-ray imaging of the dayside structures of the magnetosphere such as cusps, magnetosheaths and magnetopauses which are identified statistically by in-situ satellite observations, an original light-weight X-ray imaging spectrometer (∼10 kg, ∼10 W, ∼10×10×30 cm) will be carried. The payload is composed of a ultra light-weight MEMS Wolter type-I telescope (∼4×4 deg2 FOV

Accepted: 2018-08-05, 資料番号: SA1180110000

The hydrogen exosphere constitutes the uppermost atmospheric layer of the Earth, and its shape may reflect the last stage of the atmospheric escape process. The distribution of hydrogen in the outer exosphere remains unobserved because outer geocoronal emissions are difficult to observe from within the exosphere. In this study, we used the Lyman Alpha Imaging Camera (LAICA) onboard the Proximate Object Close Flyby with Optical Navigation (PROCYON) spacecraft, located outside the exosphere, to obtain the first image of the entire geocorona that extends to more than 38 Earth radii. The observed emission intensity distribution can be reproduced using our analytical model that has three parameters: exobase temperature, exobase density, and solar radiation pressure, which implies that hot hydrogen production in the magnetized plasmasphere is not the dominant process shaping the outer hydrogen exosphere. However, the role of the magnetic effect in determining the total escape flux cannot be ruled out

Interplanetary micro-spacecraft have been recently demonstrated for various missions. Orbital control of micro-spacecraft is challenging because the spacecraft systems have severe constraints. This...

Accepted: 2012-12-09, 資料番号: SA1004416000

The orbit of a solar sail can be controlled by changing the attitude of the spacecraft. In this study, we consider the spinning solar power sail IKAROS (Interplanetary Kite-craft Accelerated by Radiation Of the Sun), which is managed by Japan Aerospace Exploration Agency (JAXA). The IKAROS attitude, i.e., the direction of its spin-axis, is nominally controlled by the rhumb-line control method. By utilizing the solar radiation torque, however, we are able to change the direction of the spin-axis by only controlling its spin rate. With this spin rate control, we can also control indirectly the solar sail’s trajectory. The main objective of this study is to construct the orbit control strategy of the solar sail via the spin-rate control method. We evaluate this strategy in terms of its propellant consumption compared to the rhumb-line control method. Finally, we present the actual flight attitude data of IKAROS and the change of its trajectory.

Accepted: 2011-02-19, 資料番号: SA1003628000

This study aims to optimize a low-thrust trajectory for spacecraft in the event of stochastic operational anomaly. Most of the recent study have focused on the deterministic trajectory optimization for preliminary mission design. In the low-thrust operation, which conventionally use empirical (nonoptimum) operation strategies, the duty cycle of low-thrust propulsion is imposed throughout the trajectory. This paper presents the optimum value of duty cycle in the each operation cycle using Lagrange multiplier. The result shows that constant duty cycle constraint is optimal solution if surplus coasting period in the unit operation cycle is not used as thrusting period; in other words, the constant duty cycle throughout the thrust arc is not always optimum strategy if surplus coasting period is intensively used at the end of thrusting arc. Numerical example shows that the solution calculated numerically corresponds to the theoretical solution.

This paper describes design criteria for sail surface quality derived from requirements about attitude dynamics of spinning solar sail spacecraft. The method has been developed by an experience of the Japanese interplanetary solar sail mission IKAROS. The authors derived, in their previous work, a combination of a generalized spinning attitude model and a finite element model-based sail deformation analysis method to directly relate the sail surface wrinkling and the attitude disturbance via solar radiation pressure effect. Based on this work, this paper attempts to derive design criteria for the sail surface quality, which is useful for the mission analysis and the quality control in sail production. Specifically, this paper provides an acceptable deformation level as a function of scale length for given attitude stability requirements against the solar radiation pressure disturbance, which is then compared and verified by the post-flight analysis of IKAROS.

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

The Japan Aerospace Exploration Agency (JAXA) is studying the feasibility of using the solar power sail as a new propulsion engine for deep space exploration missions. In this paper, the sail shape and equipment layout for missions utilizing small-sized solar power sails are proposed. The two-stage deployment method of the sail is also proposed. The sail need to be deployed statically at the first stage, and two types of deployment mechanisms are introduced. On the other hand the second stage of the deployment can be performed dynamically, and the oscillating motion of the membrane is converged by tethers connecting the membrane to the main body. The deployment motions are analyzed by numerical simulations using multi-particle models in order to verify the deployment. They are compared with the results calculated by finite element method models. The numerical simulation results are discussed from the technological viewpoint of the sail deployment dynamics and mechanisms.

In this paper, the attitude motion and attitude control strategy of spinning solar sail are discussed. As the spinning type solar sail does not have any rigid structure to support its membrane, the impulsive torque by the RCS can introduce oscillatory motion of the membrane. Thus, an “oscillation free” attitude controller is needed, which takes into account the flexibility of the membrane and avoid unnecessary oscillatory motion. First, the dynamics model and numerical model were introduced, and the validity of these models and dominant out-of-plane membrane vibration mode is examined by membrane vibration experiment and comparison between both models. Then, based on the analysis of the dynamics of torque-free motion, it was shown that a spinning solar sail has three oscillation modes of nutation, one of which is equal to the spinning rate of the spacecraft. The dominancy of each nutation mode was analytically and numerically discussed. Then, we discussed the spin axis maneuver control using conventional RCS. It was analytically shown that continual impulsive torque synchronizing the spin rate can excite nutation velocity and that a controller is needed to damp the nutation while controlling the spin axis at the same time. The authors proposed new controller named Flex-RLC and improved one. Their effectiveness was verified by numerical simulations using precise multi-particle numerical model which can express higher order oscillatory motion of the flexible membrane, and it was found that the proposed method can control the attitude of spinning solar sail while drastically reduces the nutation velocity compared with conventional control logic. So, it can be said that the proposed method is promising fast and stable controller for spinning solar sail.

This study investigates the trajectory analysis of small solar sail demonstration spacecraft IKAROS considering the uncertainty of solar radiation pressure. Estimation of solar sail force model in space is the key factor for successful solar sail navigation because the solar sail have large uncertainty due to the flexible membrane. Since the sail wrinkles after the deployment and its surface will suffer from degradation, the solar sail force model is difficult to develop on the ground. In this paper, a practical analysis of estimating the solar sail force model from Doppler and range observable is investigated. This is demonstrated by orbit determination including parameter estimation of solar sail model. Some examples are described to investigate better parameters to estimate the solar sail force model.

The solar sailing spacecraft is one of the promising propulsion systems for the future deep space exploration mission. Japan Aerospace Exploration Agency (JAXA) has been studying the spin solar sail spacecraft which has a squared-shape type solar sail. One of the most significant objective to control the satellite orbit of the spacecraft is to estimate the thrust force induced by the photon, namely to establish the acceleration model before the launch. In a view point to use this model in orbit, the calibration of the acceleration model and evaluation of the dynamics on orbit are also important issue. This paper presents the way to construct the acceleration model of the Solar Radiation Pressure (SRP) on ground, and the calibration and evaluation strategy for this model by using the on-orbit data.

Solar sail is one of the promising propulsion systems for future deep space exploration missions as it does not require any fuel to acquire propulsive force. Although folding method and deployment mechanism of the sail have been intensively developed, attitude control of the solar sail, which is necessary for the orbital control by the solar sail, has not been much studied. This paper discusses the attitude dynamics and the control method of a spinning type solar sail spacecraft. The spinning type solar sail, where the sail equipped around the spacecraft hub is to be deployed and extended by centrifugal force, has no rigid structure supporting its membrane. This type of mechanism has the advantage in its simple and lightweight structure, however, the attitude control is difficult due to the flexibility of the membrane. In this paper, we introduced a mathematical dynamics model including first oscillation mode of the membrane which can handle coupled motion of a rigid spacecraft and a flexible membrane, and analytically developed a controller that can avoid unnecessary oscillatory motion. The performance of the controller was verified by numerical simulations using more precise multi-particle numerical model.