A constraint of rubble-pile interior is an essential issue for better understanding of its formation scenario such as catastrophic disruption and reconfiguration. A goal of this study is to demonstrate a method to constrain unknown density distribution within a small solar system body. We calculated the gravity field on the surface of asteroid 25143 Itokawa using a 3D shape model. We assumed density inhomogeneity between two lobes of Itokawa; “Head” and “Body” of the sea-otter-like shape. Then, we estimated interior density distribution of Itokawa as an index of the gravity potential variance (and surface slope average). This study implied that the Head of Itokawa had a more monolithic structure of significantly higher density than the remaining Body part. The density of the Head was evaluated as ～ 2,750 kg/m³, which is much higher than the Itokawa's mean density of 1,950 kg/m³. Such a dramatic density anomaly corresponds to a great offset between the center-of-mass (COM) and the center-of-figure (COF) by ~ 16 m. This result implies interior density inhomogeneity within a rubble-pile asteroid.

Because thruster performance and the durability of electric propulsion differ between ground tests and space operations, it is important to evaluate the rarefied propellant flow within the vacuum test facility. Through numerical analysis for rarefied propellant flow within a vacuum chamber, results confirmed that vacuum pump arrangement and thruster location influenced the propellant flow, and that a conical target installation is somewhat useful. The average pressure within the chamber is proportional to the average distance between the thruster and the pumps.

Most of solid propellants have difficulty in controlling magnitude of their thrust or stopping their operation arbitrarily. Some solid propellants show a unique property of self-quenching in intermediate pressure range. This property could be utilized for the realization of actively controlled solid propellant rocket motor. However, little is known about the mechanism of this self-quenching phenomenon. In this paper, a simulation model of this phenomenon is developed and the simulation results are compared with the experiment results to determine the mechanism of this self-quenching phenomenon. The simulation results are in good agreement with the experiment results by assuming self-quenching phenomenon has stochastic nature and by selecting proper parameters.

The development of enhanced propulsion system for the next Epsilon rocket was progressed. The development of Enhanced Epsilon is mainly the renewal of the second stage, and also includes each subsystem's improvement. The second stage motor M-35 was newly designed and manufactured. In order to verify the design, the static firing test of the second motor M-35 under the condition of vacuum ambient was conducted in 2015. The JAXA successfully launched the first Enhanced Epsilon launch vehicle. All solid propulsion systems for the Enhanced Epsilon launch vehicle showed a very good behavior during the flight

Nakashimada Engineering Works, Ltd. and Tohoku University have developed a 1U-size CubeSat “FREEDOM”. The mission of FREEDOM is to demonstrate a thin film deployment mechanism called the De-Orbit Mechanism (DOM) for space debris prevention, and includes deployment of the thin film, early de-orbiting, and tracking of its orbit transition. FREEDOM was released into low Earth orbit from the International Space Station using the robotic arm of the Japanese Experiment Module “Kibo” on January 16, 2017. The initial altitude of FREEDOM was approximately 410 km. The transition of the orbit of FREEDOM was monitored by obtaining the publicly available two line elements. FREEDOM fell below an altitude of 250 km on February 6, 2017 and is considered to have re-entered into the Earth atmosphere approximately one day later. FREEDOM stayed in orbit approximately 22 days. This result agrees with the numerical estimates and proves that FREEDOM succeeded in the deployment of the film and early de-orbiting.

Numerical simulation of aerodynamics around the spaceplane was conducted in the equivalent condition to the transonic wind tunnel tests at ISAS/JAXA. Numerical results reproduced experimental results, and they are useful to discuss the experimental results. Two configurations, namely, the experimental model and the original model were investigated. In all Mach numbers, drag coefficient of the experimental model was larger than that of the original model. Especially the wing and the tail made large drag. The different cross section of the wing and the tail caused the larger drag in the experimental model.

The reference orbit implemented for the active TerraSAR-X mission works remarkably well for orbit control purposes, but an unexpected secular drift in the along-track separation between satellite and reference orbits has built up to a 60 s flight-time offset within 10 years of operation. The scope of this work is to understand the origin of the drift and to eliminate the effect for DLR's future repeat ground-track missions EnMAP and Tandem-L. The improved process of reference orbit generation is discussed and the underlying relations for the suggested inclination adjustment are derived. The improved process is successfully validated by means of 1-year numerical orbit control simulation. The presented process is generic and can be applied to any repeat-ground track mission.

When a meteoroid impacts the Moon at several tens of km/s, a brilliant flash, referred to as a lunar impact flash, can be observed at the point of impact by ground-based telescope. Lunar impact flashes observed from the ground are biased due to atmospheric extinction, background illumination of earthshine, and short observation windows, typically a few hours per day during a period of approximately 10 days. NASA's meteoroid impact program took 10 years to detect 400 lunar impact flashes. EQUULEUS will demonstrate low-energy trajectory control techniques, such as multiple lunar flybys, proposed by the University of Tokyo and JAXA, within the Earth-Moon region. The spacecraft will be launched by NASA-SLS in 2019. The DELPHINUS camera system will be placed onboard EQUULEUS to observe lunar impact flashes while the spacecraft stays in halo orbit around an Earth-Moon L2 point. Thus, the lunar distance from the spacecraft is approximately one tenth that from ground-based observation. We estimate that DELPHINUS will detect 1,607, 2,699, and 4,534 lunar impact flashes during its six-month mission phase by assuming the limiting magnitude of its camera to be the 4.5th, 5.0th, and 5.5th magnitudes, respectively. The present study describes the DELPHINUS camera system and the first method for space-based observation of lunar impact flashes.

A thermal lattice-Boltzmann model of a van der Waals gas was used to check its applicability to specific challenges in the numerical simulation of transient cryogenic two-phase flow in rocket engine feed systems. Three test cases were chosen to prove the model is capable of capturing the underlying physics. Overall correct representation of incompressible flow should be demonstrated by a lid-driven cavity. The capability of the model to handle shocks and supersonic flow is shown in a shock tube configuration. The last test case was chosen in order to show instantaneous evaporation by the formation of a single vapour bubble at a heated surface.

For all common satellite attitude determination sensors, star trackers provide the most accurate measurement. However, these devices can be both large and expensive, and for some CubeSat applications it would not be suitable. Star trackers have in the past been successfully made for CubeSats. This paper investigates star tracker algorithms, implemented with a smartphone, so it may be used for testing attitude determination on a CubeSat. By testing through a proposed implementation, star centroids should be found by the moment method, stars should be identified by planar triangles, and QUEST should be used for attitude estimation. Smeared star images should be avoided and blurred images provide greater accuracy. Using these techniques, a star tracker using a smartphone may be constructed for attitude determination testing and software development, applied in the lost-in-space situation. This may be applied to QKD CubeSats, which require an attitude precision below 0.01°.

Ablative performance of High-density carbon phenolic (HD-CFRP) is investigated by exposing to an extremely cold temperature of -160 degree Celsius, followed by rapid heating conditions ranging from 5.5 to 12.1 MW/m² of JAXA/ISAS arc jet tests to identify the thermal integrity of the material as a future re-entry capsule heat shield material with demanding system requirements, In this paper, we introduce our activities and the outcomes by the investigation.

Pointing mechanism installed on GOSAT/TANSO-FTS is one of the most essential techniques to realize optimal observation patterns on the earth and a long-term operation period. In orbit, the continuous increase of angular rate sensor cable torque was occurred, and this induced the limited pointing angle operation for the along track direction. In this paper, we report reproducing experimental results for the degradation of the pointing mechanism in orbit.

Taking action for evacuation immediately is the most important and effective measure when tsunami attacks in the aftermath of a large earthquake. In this paper, a concept of delivering the least amount of emergency warning information and location based information possible is presented with a design of indoor/outdoor seamless positioning and emergency warning system that uses GNSS-based signals. This system includes an indoor positioning system equipped with a positioning method with GNSS-based signals and broadcasts a small volume of data including emergency warning information indoor and outdoor as well as the navigation message. A facility with large indoor space in coastal areas has GNSS antenna, receives and relays the emergency warning message from GNSS to transmitters of indoor positioning system via local area network. The system also has a function to provide indoor evacuation control information from building security room. The data communication speed of indoor signals compatible with GNSS is low as well as GNSS signals. Although the available data volume for emergency warning information is limited, we have confirmed the system is able to provide marginal information on an actual case of indoor positioning system, the indoor messaging system (IMES).

The performance improvement and cost reduction of solid rocket propellants has been a major concern of the solid rocket community. The use of low melting temperature thermoplastic as a fuel binder is one of the solutions. This paper presents a molding method of LTPs (Low Melting Temperature Thermoplastic Propellant) and the result of tests with a new mixing device.

Mid-density carbon phenolic (MD-CFRP) is a promising candidate for heat shield materials for future re-entry capsules because it could provide us with much lighter heat shields than the conventional high-density CFRP and also it could be possible to be free of delamination even if it was heated in the cross-ply direction, Here in this paper, we introduce our in-house R&D activities on the MD-CFRP which is currently underway.

This study considers the minimum time orbit-raising problem of geostationary spacecraft with low-propulsion thrusters. This problem is equivalent to determining an appropriate thrust direction during orbit raising. This study proposes a closed-loop thruster steering law that determines the thrust direction based on the optimal feedback gains and control errors of each orbital element. The feedback gains of the steering law are assumed to be the functions of orbital elements and are optimized by a meta-heuristic method. The orbital semi-major axis, eccentricity, and inclination are considered as independent variables for expressing the gains. Numerical simulations show that whichever orbital elements is selected as an independent variable, the same performances are obtained. Therefore, regardless of the initial orbital elements, by selecting the independent variable appropriately, the proposed method can always solve the minimum time orbit-transfer problem.

This paper considers a rest-to-rest spacecraft attitude maneuver using two variable-speed control moment gyros. Two methods are used to derive angular momentum and gimbal angles over time subject to physical constraints. The first is an analytical method using variational calculus: from the analytical solution the dynamic characteristics of the attitude maneuver are easily understood and feasible trajectories are derived. The second is a numerical method which uses a Chebyshev pseudospectral method. The maneuvering time is minimized by the numerical method. In this second method, it is difficult to obtain a solution without appropriate initial values. However, a solution can be derived by using an analytical solution as initial values. The analytical and numerical methods are executed in all directions of the attitude maneuver, thus validating the effectiveness of the successful completion of the intended maneuver. Finally, ground experiments are conducted using the optimal trajectories obtained by the numerical method.

Satellite charging causes anomalies and its effect still accounts for the main part of failure that occurs in the space environment. It is important to measure plasma particles in order to elucidate the interactions between such charging and the space environment. For this purpose, the Avalanche PhotoDiode (APD) type sensor is used to minimize instrument size and weight. The APD basically measures photons, but can also be used to measure electrons. This paper reports the performance of the APD sensor, and presents the instrument design results for flight investigation.

Space debris poses a serious threat and the problem must be solved. Space debris should be removed to maintain the space environment. One of the techniques to remove it is an electrodynamic tether system (EDTS). To improve the toughness of the tether used in EDTS, this study investigated the tether shape of a hollow cylinder instead of the conventional solid cylinder. When the mass and length of a hollow tether are the same as those of the solid tether, the hollow cylindrical tether has a larger diameter. Thus, the toughness against small debris collisions will be improved. To investigate the damage to the hollow tether following a debris collision, hypervelocity impact experiments were performed using a two-stage light-gas gun. Through the impact experiments, the relationship between the damage to the hollow tether and the impact condition of the debris was revealed. It was confirmed that when the impact angle increased the damage area increased as well. The experiments validated that a hollow tether is superior to a solid tether in terms of resistance to debris collisions. In addition, the importance of considering the influence of the impact angle was shown by an evaluation of the lifetime of the hollow tether.

This paper describes the “2016 KOSEN Space Camp,” which is an introductory education program on space technology for students enrolled at the various KOSEN colleges of technology in Japan. The camp was conducted twice, the first time on September 3–6, 2015 and the second time on September 1–4, 2016 at the Marine Park Niihama. Approximately 40 students and more than 10 teachers participated in the camp on each of the two occasions when it was held. The activities in the camp included a lecture by an eminent scientist; the study of basic rocket theory; model rocket experiments; and the development, analysis, and presentation of CanSat satellite model experiments. This unique education program provided by aerospace scientists and engineers is a KOSEN concept and product.

The Martian Moons eXploration (MMX) mission is now under study by the Japan Aerospace Exploration Agency (JAXA). Its scope includes the world's first landing on one of the Martian moons, collecting samples from the surface, and returning to Earth. This paper describes the orbit design for MMX. Nominal and backup trajectories for launch in 2024 and 2026 are discussed. The Mars orbit insertion (MOI) sequence using 3-impulse maneuvers is introduced. A new scheme, the robust MOI, is also proposed as a contingency to enhance the robustness of the mission sequence. A method to design a robust MOI trajectory and examples are presented.

Failure of the orbit insertion maneuver has a significant impact on the entire mission, for the trajectory of a spacecraft is largely deflected by swing-by. The risk can be reduced by targeting a point on the B-plane where the spacecraft reaches the free-return (FR) trajectory with the target body in the case of insertion failure. The backup orbit must also satisfy conditions suitable for the mission. We investigated the type of orbit insertion that is both robust to failure and reasonable for the mission requirements. We call this method FR ensured orbit insertion. Among various failure modes of the orbit insertion maneuver, we focus on the complete maneuver failure. The impact parameters on the B-plane to achieve the orbit insertion are formulated based on the geometry of velocity vectors at swing-by. The necessary deflection angle α_FR at swing-by must be smaller than the possible maximum deflection angle α_max for the target body. When we introduce proper scaling factors, the relation of α_max and α_FR is characterized by a single parameter λ. Using polar orbit insertion as an example, maps which show the reachability of FR trajectory after the insertion failure for each approaching condition are presented. The derived maps can be used as a tool to assess the applicability of the method in the mission design. Finally, as an application to practical mission design, we demonstrate the use of FR ensured orbit insertion in JAXA's MMX mission.