We report on the effect of feeding the propellant on the ion beam in porous ionic liquid electrospray thrusters and discuss it in the context of the ion dynamics in the ionic liquid. Ionic liquid electrospray thrusters are a promising, game-changing electric propulsion system that enables high ∆V even for small spacecraft. Especially, a porous emitter is beneficial for storing ionic liquids and is widely used in the development of electrospray thrusters. However, there are no guidelines for the design of propellant feeding in porous emitters. This is one of the bottlenecks in improving the thruster technology. To understand the effect of the pressure of the feeding, we applied feeding pressure to the emitter using a liquid column, using gravity to make it an actively fed system. This active feeding with positive pressure enhanced the emitted current. Based on the theoretical model in [Takagi et al., J. Appl. Phys, 135 (2024), 244502], applying feeding pressure increased the ion transport driven by the ion pressure gradient, resulting in an increase in the emitted current. The importance of controlling the feeding pressure for the porous emitter is thus clarified.

As number of microsatellites continues to increase every year, the demand for propulsion system is also expanding. As continuing to develop propulsion systems for microsatellites, chemical propulsion systems using low toxic propellants, commonly referred as green propellants, are emerging. This research team has been conducting research and development of a chemical propulsion system using 60wt% hydrogen peroxide (60wt%H₂O₂), which is relatively easy to handle, for microsatellites, called Microsatellite-Friendly Multi-Purpose propulsion system (MFMP-PROP). This paper reports the results of the fundamental evaluation of the performance of a valve injector that has been diverted from automotive/motorcycle to a space propulsion system to achieve an inexpensive and highly reliable propulsion system. The thruster was confirmed to operate as both mono- and bi-propellant mode, and its performance was evaluated.

This paper presents the development of a Functional Resonance Analysis Method (FRAM) network model to analyze the launch operations of various chemical rockets. This FRAM-based approach offers a systemic perspective for optimizing launch operations and for mitigating risks in complex aerospace systems. Functions that are necessary for the launch process, including stage assembly, electrical system tests, and countdown operations, were identified. Moreover, their mutual relations were defined. The resulting model represents high-risk and frequently used functions effectively, thereby providing insights into system performance variation and potential resonance effects. However, the analysis highlighted the need to incorporate additional functions related to ground safety and cyber incident management to enhance model completeness. Future improvements will be made to address these gaps to strengthen the model’s applicability. Additionally, this study demonstrated the necessity of exploring methods to improve launch system resilience, with specific examination of integrating a learning function to mitigate variation and to enhance operational robustness. An investigation into the effects of this learning function on network resilience is also planned.

The growing demand for Ka-band communications led Resolution 169 (WRC-19) to allocate additional frequencies to Earth Stations in Motion (ESIM) and establish technical requirements for their coexistence with terrestrial systems. To safeguard terrestrial services, Resolution 169 specifies per-aircraft limits on the permissible power flux density (PFD). Accurate estimation of the PFD at terrestrial base stations requires consideration of attenuation caused by the aircraft fuselage. This study evaluates the fuselage’s effect by comparing measured radiation patterns of an active electronically scanned array (AESA) antenna with equivalent isotropically radiated power (EIRP) values derived from the PFD limits set in Resolution 169. The fuselage reduces backlobe levels by up to 15.3 dB at 29.75 GHz, improving compliance with Resolution 169.

Given the recent rapid increase in the number of space objects, including space debris, in outer space, today, the establishment of proper governance for space activities in Earth's orbit has become a common agenda for all space-faring nations. For the time being, many countries, including the private sector, have been engaging in various efforts to establish national or regional Space Situational Awareness (SSA) capabilities as a foundational infrastructure to tackle Space Traffic Management (STM) and space debris issues. Now, we are facing the question: What kind of framework should we establish to achieve international cooperation and coordination among these SSA activities and systems? This paper conducts a preliminary study on the possible options and policy implications for considering the international cooperative framework for SSA by analyzing the features of existing international frameworks for global Earth observation and global navigation satellite system (GNSS). Based on the analysis, it is found that the preferences of leading actors, the flexible nature, and the participation of the private sector are important points for the future consideration of the framework.

In distributed space systems such as formation flying and constellations, relative orbit control is an essential technology. Thrusters are often used for this purpose; however, in low Earth orbit, aerodynamic forces can also be utilized. Unlike thruster-based control, aerodynamic control does not consume propellant, so the mission lifetime is free from propellant limitations. However, it cannot, in principle, increase orbital altitude, and the additional installation of aerodynamic control devices is instead expected to accelerate orbital decay. This paper investigates whether thruster-based control or aerodynamic control is more advantageous for relative orbit control from the perspective of mission lifetime. For thruster-based control, the operational duration is estimated through simulations assuming a flight-proven thruster suitable for small s atellites. For aerodynamic control, orbital decay is simulated for satellite con-figurations equipped with aerodynamic control paddles, and the feasible mission duration is estimated. These analyses are conducted across multiple classes of satellites to compare their mission lifetimes. As a result, the aerodynamic control paddle area and the scale of the relative orbit that make aerodynamic control advantageous over thruster-based control in terms of mission lifetime were clarified.

Small balloons provide an effective means of conducting in situ observations in the stratosphere. However, payload recovery has been challenging due to their passive and wind-dependent flight during both the ascent and descent phases. In this study, a parafoil-controlled payload module was developed to improve the recoverability of such payloads. The payload module consists of a parafoil with a 2.25 m wingspan and a control unit. To evaluate its performance, a balloon release experiment was conducted. The payload module was carried by a small balloon to an altitude of 4,522 m, where it was separated. After separation, the module glided under remote control. The mission was successfully concluded with a successful recovery in the designated sea area, establishing the feasibility of guided descent and demonstrating the potential for reliable payload retrieval. This manual control experiment serves as a proof of concept for future autonomous guidance systems, thereby expanding the technological scope of small-balloon missions.

Space development is shifting from an era led by national research institutions to one driven by private companies. Human exploration missions are now being scheduled from Earth orbit to the Moon and Mars, resulting in a diversification of gravity environments where people will reside—ranging from 0G to 1/6G and 1/3G. As commercial use of space facilities progresses, it is expected that people with less training than astronauts will begin living in space. This study aims to identify the challenges associated with evacuation behavior in a simulator that replicates lunar gravity, and to consider the conditions necessary to ensure evacuation safety in a 1/6G environment.

In-orbit servicing missions, such as space debris removal and satellite refueling, require precise and robust rendezvous maneuvers to ensure safe and effective operations. This paper proposes an integrated parameter-tuning framework for designing multi-objective controllers for spacecraft rendezvous, leveraging linear matrix inequalities (LMIs). The framework addresses the challenge of tuning multiple parameters within LMIs to satisfy diverse constraints, such as stability and performance, in continuous-time control systems. The proposed method utilizes Bayesian optimization to optimize the evaluation function efficiently. Monte Carlo simulations are used to calculate this function, which enables smooth relative orbit control with minimal tuning trials. Simulation results are used to compare the proposed Bayesian optimization-based approach with trial-and-error and random search methods, demonstrating its superior performance in achieving smooth relative orbit control.

Sputtering erosion of propulsion systems and experimental equipment is inevitable in electric propulsion systems using high energy plasma. Therefore, understanding the rate of sputtering is critical to assessing the equipment’s lifetime and preventing spacecraft contamination. This study aimed to evaluate the extent to which residual air in a vacuum chamber affects the sputtering rate of stainless steel and titanium and propose a model based on parameters such as background pressure and ion beam current. The sputtering rates of stainless steel, titanium, iron, and copper were determined using mass difference measurements using argon ion irradiation at 1500 eV. The results showed that the sputtering rates of stainless steel, titanium, and iron significantly decreased with increasing background oxygen partial pressure, whereas the reduction in the sputtering rate of copper was comparatively smaller. The model successfully explained the experimental results for stainless steel, titanium, and iron and accurately reproduced the experimental results of copper.

Electrodeless radio frequency (RF) electric propulsion systems are promising for future space propulsion due to the absence of grid erosion and necessity of neutralizer. To enhance thrust performance in RF thrusters, we are proposing an additional acceleration method using the Rotating Magnetic Field (RMF). The RMF method drives an azimuthal electron current, and the Lorentz force acts on electrons in the presence of a magnetic nozzle, enhancing a net plasma acceleration. In this study, a small-scale RMF thruster simulator is newly developed to evaluate the RMF acceleration effect in the small thruster scheme. As a preliminary study, we measured spatial profiles of plasma parameters, and an increase in ion Mach number was observed. Furthermore, the thrust was roughly estimated from these parameters, indicating a clear enhancement due to the RMF. We also evaluated the dependence of ion Mach number and static pressure on the RMF’s magnetic field strength.

Gravity compensation devices are indispensable for ground testing of deployable space structures. While various devices have been utilized, many lack versatility as they are custom-designed. Therefore, we propose a highly versatile gravity compensation device applicable to three-dimensional deployment. This paper focuses on the control system of a pneumatic cylinder, a key component of the proposed device. The system employs a differential pressure control method, supplying air to both cylinder chambers to achieve precise and continuous force adjustment, unlike conventional ON/OFF control using solenoid valves. The basic performance was experimentally evaluated. The results confirmed consistent pressure responses within the circuit but also revealed two critical issues: a significant response delay of approximately 1 second to step inputs, and a notable thrust hysteresis with a width of about 20 N, caused by internal static friction. This quantitative characterization is crucial for accurately modeling the system and developing future control strategies for its practical application.

Surface erosion was observed near the “μ10” gridded ion thruster onboard Hayabusa2 during its space operation. To investigate the interaction between the thruster plume plasma and spacecraft surfaces, laboratory experiments were conducted using the same type of thruster in a vacuum chamber. These tests revealed surface contamination not only near the thruster but also on nearby diagnostic instruments, likely caused by sputtered materials from beam target irradiation by high-energy ions. Based on these findings, follow-up ground experiments were conducted to assess the elemental characteristics and deposition effects of the contamination more precisely. Contaminants were sampled and analyzed through SEM-EDS and thickness measurements. Results indicated that the deposits were primarily composed of elements from the chamber wall materials, with a maximum thickness of approximately 500 nm. The contamination was estimated to lead to an underestimation of the incident ion flux by approximately 5% in the worst case. While this is not a drastic change, it is a factor that should be noted regarding measurement accuracy.

Electrospray thrusters have attracted attention as compact propulsion systems capable of precise thrust control. The thrust of electrospray thrusters depends on the ion beam current and ion mass, and the form of the emitted ions varies depending on the flow rate of the ionic liquid supplied to the emitter. Therefore, it is necessary to understand the effect of the flow rate on the beam characteristics in order to optimize the thruster performance. In this study, we have developed a pressure-based flow control system to vary the flow rate of the ionic liquid to externally wetted emitters, and measured the ion beam currents to evaluate their characteristics. We find that the beam current tends to increase with increasing supply pressure, and that the current–voltage characteristic depends on the flow rate.

Transonic buffet-like low-frequency oscillations are known to occur in the compressible low-Reynolds-number regime, where future Mars airplanes are expected to operate. Transonic buffet can cause structural damage on aircraft wings, and its careful charac-terization is needed for safe flight on Mars. In this study, to enhance the understanding of buffet in the low-Reynolds-number regime, its sensitivity to geometrical parameters like airfoil thickness and sweptback angle was evaluated through numerical simulations. As the airfoil thickness was increased, buffet in the low-Reynolds-number regime emerged and eventually saturated. For even thicker air-foils, non-stationary behavior was observed with massive separations on both sides of the airfoil. Despite such behavior, the presence of buffet was confirmed by Dynamic Mode Decomposition (DMD). At lower Mach number and lower Reynolds number, the onset of buffet occurred at increased airfoil thickness. Regarding the effect of sweptback angle, buffet in the low-Reynolds-number regime was consistently identified in all the test cases. While the presence of sweptback angle did not significantly affect the amplitude of the lift coefficient, it did have an impact on the drag coefficient. The findings of this study provide new insights into buffet in the low-Reynolds-number regime, valuable for Mars airplane design.

Formation flying has become a crucial technology that enables innovative space missions such as high-resolution Earth obser-vation, space-based interferometry, and distributed sensing. The deployment of multiple small satellites in precise formations offers advantages such as enhanced mission flexibility, improved system redundancy, and reduced overall mission costs compared to tra-ditional monolithic satellites. Furthermore, it is expected that autonomous decision-making capabilities in a satellite swarm will enable collective action such as the autonomous assembly of large structures like solar panels and antennas in orbit. This study in-vestigates cooperative control of satellite swarms using a coverage control framework. In coverage control, satellites autonomously distribute themselves within a predefined domain or partition the domain into subregions to meet specified constraints. To implement this, Voronoi diagrams are utilized, which partition space based on proximity and enable optimal spatial distribution by driving each satellite toward the centroid of its Voronoi cell. The coverage control method for three-dimensional satellite motion is demonstrated through numerical simulations, allowing for more realistic modeling and control of formation flying in orbital environments.

This study presents an Extended Kalman Filter (EKF)-based approach to estimating internal combustion states in hybrid rocket motors with varying port geometries. A total of eleven hot-fire tests have been conducted using 3D-printed PLA grains with twist angles ranging from 0° to 810°, and chamber pressure and orifice differential pressure have been measured during firing. The EKF is designed to estimate four key state variables—hydraulic diameter, oxidizer mass flow rate, fuel regression rate, and characteristic exhaust velocity efficiency—under three geometric modeling configurations, culminating in a full model incorporating end-face combustion and twist-induced surface expansion. Model performance is evaluated through time-averaged comparison against experimental data using Bland–Altman analysis, revealing that full geometric modeling reduces bias and variability significantly. Additionally, the effect of observation noise covariance on estimation behavior is examined via time-series comparisons and the coefficient of variation. Results demonstrate that higher confidence in model dynamics (larger observation noise settings) yields more stable estimates, though transient mismatches may emerge due to input model assumptions. These findings support the EKF's viability as a robust soft-sensor framework for onboard monitoring and control in hybrid rocket applications.

We propose the “laser-detonation expansion tube” concept to simulate the very low Earth orbit (VLEO) flow conditions with high-speed and rarefied characteristics in a ground-based test facility. A two-dimensional axisymmetric computational fluid dynamics (CFD) simulation was performed to validate its performance in generating the desired flow. The laser energy deposition was modeled within the expansion tube, which had two diaphragms separating the different pressure chambers. The CFD results show that a test flow with a Mach number of approximately 7, speed of 2,200 m/s, pressure of 350–800 Pa, and specific enthalpy of 2.4 MJ/kg could be driven after the laser-induced detonation wave passed through the second diaphragm. The flow duration was estimated to be approximately 25 µs, which is sufficient for conducting experiments on the air-breathing electric propulsion system for VLEO satellite operations.

To establish a presence in Mars exploration, the Japan Aerospace Exploration Agency (JAXA) is studying a phased approach, beginning with multiple ultra-small (20-30 kg) landers. This initial step will demonstrate a unique Entry, Descent, and Landing (EDL) strategy combining a deployable aeroshell for efficient aerodynamic braking and a penetrator for a simple, semi-hard landing. This concept creates a highly reliable EDL system without complex parachutes. This paper details the development of the lander's critical power system components, designed for the harsh Martian environment. We prototyped flexible solar array sheets using high-efficiency IMM3J cells integrated onto ZYLON fabric, the same material as the aeroshell. In the structural evaluation of the flexible solar array, the mechanical damage mechanisms were elucidated by comparing polyimide and woven ZYLON cores, revealing that crack generation under severe folding is highly dependent on local rigidity, relative thickness, and structural homogeneity. A flexible, amorphous-like Indium Tin Oxide (ITO) coating was successfully applied to the array's surface, demonstrating a significant dust mitigation effect. Evaluations using lunar and Martian simulants demonstrated the coating's fundamental effectiveness in preventing localized charge accumulation, while highlighting the necessity of considering macroscopic environmental factors like gravity and atmosphere, alongside terrestrial moisture artifacts. Furthermore, a custom lithium-ion battery was developed. This laminate-type cell can be charged and discharged at −30°C, retaining 73% of its room-temperature energy. It maintained a Coulombic efficiency of 99.6% and a capacity retention of 87.5% over 100 cycles at −30°C, confirming that irreversible reactions are minimal and establishing a fully viable baseline for the lander. It also proved its required shock resistance by withstanding a simulated landing impact of approximately 800 G without significant degradation. These successful results demonstrate the feasibility of the power system for the ultra-small Mars lander.

The amount of information handled in space has been increasing due to the growing number of satellites and the increasing volume of data processed by satellites. In order to effectively utilize the information acquired by satellites, a method for high-capacity communication is required. We have studied orbital angular momentum (OAM) in radio waves (GHz order) as a potential method for high-capacity communication. The reception of OAM in radio waves generally requires detecting all signals on a surface perpendicular to the radiation axis. However, the intensity distribution of the OAM radio waves spreads in a cylindrical pattern, and the radius of the ring increases with propagation distance. Consequently, it is difficult to receive useful signals at long distances because extremely large antennas are required. To address this reception problem of OAM, we propose a method of partial detection using a phase rotation process at the transmitter. In this work, we describe the principle of the phase rotation process that modulates the phase of the transmitted signal for partial detection and report the simulation results of transmitting the signal using a circular array. Our findings demonstrate that the proposed method is effective for measuring the state of OAM radio waves.

This article presents a comprehensive overview of ground-based atomic oxygen testing facilities that are currently in operation worldwide. It begins by outlining the key characteristics of an ideal testing facility, followed by a detailed description of the principal methods used for atomic oxygen generation and beam formation. Each facility is then characterized by its main parameters, including atomic oxygen generation and beam formation techniques, kinetic energy, flux density, beam composition, and beam or sample size. In addition, the operational ranges of the facilities are visualized in terms of kinetic energy, flux density, and beam composition. The comparison of these operational ranges shows that no single facility is capable of fully replicating the conditions encountered in Very Low Earth Orbit. This finding highlights the need for improved facility characterization, greater standardization, and enhanced cooperation among different testing facilities.

Resonant orbits play a crucial role in trajectory design within the cislunar environment. In particular, resonant transition periodic orbits (RTPOs) offer promising pathways for low-energy transfers in the Earth-Moon system. RTPO is a periodic orbit whose trajectory contains (i) a segment shadowing a p:q resonant orbit and (ii) a segment evolving in the neighborhood of a libration point orbit (here, an L₁ Lyapunov orbit), so that a single periodic orbit spans both the resonant and libration-point regions in phase space. While it is known that these orbits reside in chaotic regions where resonances overlap, the specific geometric structures governing their location have remained unclear. This study investigates the formation mechanism of RTPOs in the Planar Circular Restricted Three-Body Problem, focusing on transitions mediated by the manifolds of Lyapunov orbits near the collinear libration point L₁. The manifold structures of resonant and Lyapunov orbits are analyzed using Periapsis Poincaré Maps to uncover their geometric interactions. We demonstrate that the invariant manifolds form characteristic rhombic structures, and RTPOs appear as periodic orbits confined within these manifold intersections. This finding establishes a clear geometric correlation between RTPOs and the invariant manifolds, offering a new perspective on their dynamical origin. Consequently, the rhombic structure is proposed as a key geometric indicator that characterizes the existence regions of RTPOs within the chaotic phase space.

One of the key challenges in future lunar exploration is the development of countermeasures against the accumulation of lunar dust on exploration systems. To address this issue, effective regolith cleaning technologies must be developed. One prominent solution is an Electrodynamic Dust Shield (EDS), which utilizes an electrostatic traveling wave to remove particles from surfaces. Although previous studies have demonstrated the effectiveness of EDS in removing lunar regolith simulants under various system parameters, the effects of additional charging methods, which are likely to influence the cleaning performance of the charged particles, have not been thoroughly investigated. In this study, we investigated the influence of incorporating ultraviolet (UV) light into an EDS system to improve the cleaning of regolith particles. Although previous studies have shown that UV assistance enhances the cleaning performance of EDS, investigations have been limited to a narrow range of experimental conditions, and the fundamental effects of UV irradiation have not been characterized sufficiently. Therefore, cleaning tests using EDS were performed under various key conditions with UV irradiation. For particles smaller than approximately 100 μm, the UV-assisted EDS exhibited a significantly higher removal efficiency than the non-UV cases under most conditions. In addition, the effects of the applied voltage, frequency, particle deposition amount, and UV irradiation duration were systematically characterized, providing a more comprehensive understanding of the influence of UV irradiation on the EDS cleaning performance.

Hall thrusters are a type of electric propulsion, typically operated with xenon (Xe) as a propellant owing to its high atomic mass and low ionization energy. However, xenon suffers from limited availability and high cost. Therefore, this study focused on carbon dioxide (CO₂) as an alternative propellant, given its low cost and suitability for low-pressure storage. Previous studies of Hall thrusters using CO₂ as a propellant investigated thrust performance and plume diagnostics using probes, aiming to improve overall performance. However, the achieved thrust performance has not yet reached the target level. To address this issue, Full Particle-in-Cell (Full-PIC) models have been developed to analyze plasma characteristics inside the thruster. While previous simulations have primarily considered only electron-impact ionization of CO₂, comparisons with experimental results have revealed the importance of including CO₂ dissociation and multi-step ionization processes. In this paper, the simulation model was improved by implementing CO₂ dissociation and multi-step ionization, and their effects on plasma characteristics and ion fractions were quantitatively evaluated. Understanding these micro-physical phenomena is crucial, as it forms an essential foundation for the future accurate prediction and optimization of the overall thrust performance. The results showed that the CO₂ dissociation accounted for more than 14% of the total reaction rate, representing a significant contribution from a single process. It was also confirmed that this dissociation was followed by the ionization and electron excitation of the resulting CO and O fragments. Regarding the ion species fractions, the simulation results showed improved agreement with experimental values. This is attributed to an increase in the production of CO⁺, O⁺, and C⁺, accompanied by a corresponding relative decrease in the CO₂⁺ fraction. However, the simulation still overestimated the ion beam currents for CO₂⁺ and O⁺. This discrepancy suggests the need to incorporate ion-neutral collisions downstream of the thruster and to investigate the influence of background pressure in the experimental environment.

In this study, we conducted an intercomparison of satellite cloud products derived from EarthCARE/MSI released in March 2025 and Himawari-9/AHI. The comparison results revealed that the cloud detection by MSI’s clear confidence level and AHI’s cloud mask has a good agreement with a high accuracy of more than 85%. Cloud properties such as cloud top temperature, cloud top height, and cloud effective radius were relatively consistent as well as cloud mask. On the other hand, it was found that MSI cloud optical thickness (COT) was exponentially overestimated for both water and ice clouds with reference to that of the AHI. The comparison result of reflectance and brightness temperature demonstrated that the over-trend of MSI’s COT derived from brighter reflectance of visible and near infrared band of MSI than those of AHI. Additionally, an experimental correction for visible band bias in the cloud retrieval algorithm supported the possibility that high visible band values affect the overestimation of COT. It suggested comprehensive evaluations will be essential for establishing characterization of the MSI cloud product performance.

This study investigates a novel orbit control method using solar sails called Discrete Locally Optimal Control. Conventional control methods require continuous attitude control, which achieves high performance but imposes significant operational load on the attitude determination and control system. To address this limitation, we propose a semi-analytical discretization of the control law that allows arbitrary maneuver frequency while preserving near-optimal performance. The proposed method was shown to closely approximate continuous control, indicating that the discretization provides an effective representation. A key insight is that increasing the number of discrete maneuvers improves performance, and beyond four to six maneuvers per orbit, the achieved performance closely approaches that of continuous control. This approach provides a practical balance between orbit control capability and operational effort, making it suitable for long-duration, propellant-free missions in planet-centered orbits.

Recently, multimode propulsions that use the same ionic liquids (ILs) for electric and chemical propulsion, such as ammonium dinitramide (ADN) and hydroxylammonium nitrate (HAN), have been developed. In this study, an IL electrospray thruster was operated with an ADN-based energetic IL, and its characteristics were compared with a conventional IL, 1-ethyl-3-methylimidazolium tetrafluoroborate (EMI-BF₄). The plume composition was analyzed using the time of flight (ToF), and the thruster performance was evaluated with an analytical balance. A broader mass distribution is observed for ADN-based energetic IL in ToF. In addition, the ADN-based energetic IL showed ranges of 225 – 415 s in specific impulse and 4.1 – 11.3% in thrust efficiency, while the conventional IL showed 446 – 876 s and 7.1 – 18.1%, respectively. Although these thruster performance obtained by the analytical balance include relatively large uncertainty due to several factors (e.g., instability in thruster operation and electrochemical reactions), the results confirm that AMU can operate as an electrospray propellant and is promising for multimode propulsion.

To meet the demands of future deep-space sample return missions and high-frequency reentry systems from low Earth orbit, a new type of thermal protection system (TPS) material has been developed. The design aim of this system incorporated the graded ablator and induced artificial controlled delamination around surface area for reducing heat transfer to the internal structure. The outermost layer is fabricated with high-density to reduce surface recession, while the base layer employs a low-density configuration to enhance thermal insulation and reduce weight. Basic thermal characteristics of the developed ablative material were obtained by conducting heating tests. The heating tests were carried out in the arcjet wind tunnel facility in the Sagamihara campus in JAXA for heat flux of 3.4~4.4 MW/m² and impact pressure of 11.6~12.7 kPa. The time variations of surface and in-depth temperature, the amount of mass loss of ablator were successfully obtained during the heating tests. Furthermore, in the specimen where the artificial controlled delamination was observed, the internal temperature was lower than that of the novel TPS. According to X-ray CT inspection conducted after the heating tests, unexpected delamination between layers was not observed inside the specimens. In addition, the present study showed the developed ablative material has a potential to reduce the TPS weight by 30% or more compared with the Hayabusa ablator.

This paper presents an overview of precipitation measuring missions previously conducted by Japan Aerospace Exploration Agency (JAXA), including the Tropical Rainfall Measuring Mission (TRMM), the Global Precipitation Measurement (GPM) mission, and the upcoming Precipitation Measuring Mission (PMM), which is scheduled for launch after Japan’s FY2028. These missions have contributed significantly to global precipitation monitoring using spaceborne radar systems. In anticipation of future scientific needs, this study investigates one of the concepts for a new precipitation measuring system based on a constellation of small satellites. The proposed architecture employs distributed radar techniques with formation flying, aiming to achieve high-resolution cross-track measurements through a synthesized large antenna aperture. We provide strategies of orbital design and station keeping, and present the results of detailed orbital analyses, including the effects of non-conservative perturbations, to evaluate the feasibility of the proposed configurations. Technical challenges related to the implementation of the proposed system for precipitation monitoring are also discussed.

Plasma particle simulations were conducted using a three-dimensional particle-in-cell simulation with a Monte Carlo collision method (PIC/MCC) to investigate instabilities that could cause cross-field electron transport in a magnetic nozzle. The plasma was located upstream of the divergent magnetic field, in order to reduce the computational cost while effectively reproducing the acceleration process in the magnetic nozzle. The simulation results were qualitatively consistent with experimental observations, showing that plasma instabilities develop in the downstream region. Furthermore, it was demonstrated that low-frequency waves contributed to the inward transport of electrons, leading to electron detachment from the magnetic nozzle.

The zero-velocity surface has been one of the most fundamental structures in many dynamical systems including gravitational N-body problems, chemical reaction models, and mechanical systems. It plays the role of a barrier placing bounds on the motion and energetically separating the regions of possible and impossible motion. Our previous work deduced that states associated with the barrier structures globally minimize an energy function. This global energy minimization property has been useful in identifying a new barrier structure in addition to the zero-velocity surface. The present paper extends the analysis method and explores the existence of other barrier surfaces in a circular restricted three-body problem. The comprehensive search in all the possible position, position-velocity, and velocity spaces unveils two novel barrier surfaces in addition to the already known barrier surfaces. The dynamically distinct characteristics of the newly identified barrier surfaces are analytically presented.

Radio waves carrying orbital angular momentum (OAM) possess modal degrees of freedom determined by the direction and spatial frequency of their helical phase fronts. In this study, we propose a novel method for transmitting and receiving structured radio waves formed through the superposition of multiple OAM modes. The state of a structured wave is represented on a Higher-Order Poincaré Sphere (HOPS), with a pair of orthogonal OAM modes assigned to the north and south poles. This representation enables applications in multiplexed communication systems, where information is encoded in spherical coordinates, and in radar imaging systems, where target scattering characteristics are analyzed with respect to the incident OAM mode (OAM polarimetry). We describe the design methodology based on spherical coordinate mapping and the corresponding transmission and reception system employing a circular antenna array. To validate the proposed method, we utilized the analogy between electromagnetic and acoustic wave phenomena to allow for direct observation of the complex spatial wavefronts. An 8-channel circular loudspeaker array with a radius of 18.5 cm was constructed to generate structured waves using superimposed OAM modes of ±1 at a carrier frequency of approximately 500 Hz—corresponding to the 2 GHz band via scale conversion. The Stokes parameters were computed from the received signals and visualized on the HOPS. Although the Mean Angular Error increased from 13.24 degrees at 10 cm to 70.26 degrees at 50 cm due to multipath propagation, the topological integrity of the transmitted states was preserved. These results demonstrate the feasibility of topologically robust communication and propagation-path-sensitive radar imaging using OAM-based structured waves. Finally, we discuss challenges related to long-distance propagation and radio-frequency implementation issues such as antenna mutual coupling.

This study proposes a “Laser-plasma-driven explorer for UNtouched low Altitude (LUNA)” to overcome the performance lim-itations of conventional air-breathing electric propulsion systems in the rarefied environment. The study aimed to maintain satellite flight at an altitude of 100–150 km on an orbit, called “extremely low earth orbit (ELEO),” using the LUNA, where no space device has achieved continuous flight. The rarefied flow at 100 km-altitude was simulated using the direct simulation Monte Carlo (DSMC) method. The contribution rate of the laser-heated particles to thrust generation was 7.90%, indicating the necessity of improving the geometrical design of the thruster. Parametric study results on the accommodation coefficient revealed that the specular thrust wall is suitable to achieve high thrust performance, as described by the thermalization of a particle. The required beam power for orbital maintenance in an elliptical orbit was estimated using the aerodynamic forces, thrust-supply duration, and power losses through laser transmission and energy transport in plasma. The study found that a flight in ELEO could be maintained using 57.1 MW of beam power.

The performance of upper-stage rockets in multistage solid rockets is an important factor in increasing the space transportation capability. Currently, the performance of Ammonium perchlorate-based composite propellants, which are the most widely used in solid rockets, can be improved by increasing the Al content. However, increasing the Al content increases the burning rate, which requires a change in motor design. In this study, a method for improving the performance of solid rockets without changing the motor design other than the solid propellant composition was investigated by increasing the propellant density and suppressing the burning rate by adding a negative catalyst. The burning rate reduction rate was calculated by changing the amount of negative catalyst added based on combustion test results. The theoretical propulsion performance was compared using chemical equilibrium calculations based on the test results. Theoretical calculations identified aluminum hydroxide, calcium carbonate, melamine, and oxamide as effective negative catalysts. Furthermore, it is shown that adding 5 parts of aluminum hydroxide, melamine, and oxamide would improve density specific impulse by approximately 1.5% without changing the motor design. In particular, adding 5 parts or less of melamine is expected to further improve.

Laser propulsion is a space transportation system that generates thrust by using a ground-based or satellite-based laser beam as an energy source. Fiber lasers are promising candidates for this application due to their high output power and excellent beam quality, which make them suitable for long-distance energy transmission. In this study, we investigated the thrust performance of a laser-sustained plasma (LSP) propulsion system driven by a fiber laser. Using argon as the working gas, we varied nozzle shape, mass flow rate, and pressure, and evaluated thrust efficiency by the sonic-flow method. Four nozzle throat models were developed: Model I (baseline), Model II (shortened distance between LSP generation and throat), Model III (narrow confinement section for enhanced heat exchange), and Model IV (a sub-chamber with an aperture upstream of the throat). Among these, Model III exhibited the best performance, with a maximum specific impulse of 139 s and a thrust efficiency of 21.5%. In Model IV, the LSP was successfully moved into the sub-chamber; however, thrust estimation was limited due to measurement constraints. These results demonstrate that nozzle throat configuration strongly influences heat exchange and propulsion performance in fiber-laser LSP thrusters.

As established, three-axis attitude control of spacecraft has conventionally been achieved using actuators such as gyroscopes, reaction wheels, and momentum wheels, which generate angular momentum by rotating a solid mass around its principal axis. In contrast, the VEKTOR-FDA (Vector Fluid Dynamic Actuator) introduces a novel approach to this challenge. Instead of relying on solid masses, this design utilizes a high-density fluid medium, specifically liquid metal [1]. Electromagnetic pumps are used to drive and circulate the fluid within the system [2]. In the VEKTOR-FDA system, the fluid is contained within a hollow spherical structure, where pumps actively circulate it by drawing it out through an outlet channel and reintroducing it via an inlet channel back into the sphere. This process generates a rotating fluid volume inside the sphere, producing an angular momentum vector aligned with the axis of rotation [3]. By implementing at least three pumps arranged orthogonally, fluid flow can be directed along each of the three principal axes, thereby establishing individual rotating fluid volumes. When these pumps operate concurrently, flow mixing occurs, enabling control over the resultant flow by adjusting the individual flow velocities of each pump. This modulation allows for the orientation of the rotation axis of the combined flow in any desired direction. Consequently, the angular momentum vector and the rotating fluid volume can be aligned spatially in any orientation. As a result, a single VEKTOR-FDA actuator can govern all three axes of the spacecraft, effectively functioning as a Multi Degree of Freedom (MDOF) actuator [4]. The VEKTOR-FDA project, initiated in October 2023 by the Chair of Space Technology at the Technical University of Berlin, aims to explore the feasibility of this innovative technology. The project’s primary objective is to assess its technical viability through the design and development of a technology demonstrator. This demonstrator will be subjected to rigorous testing and qualification in a simulated space environment to validate the technology’s functionality, with the project expected to be completed by the end of 2026. This paper provides a comprehensive overview of the project's motivation, concept, and technical foundations. It also describes the current development phase, highlighting insights gained from the experimental results of the development model, as well as the challenges encountered during the development process.

Penetrators are hard-lander-type instruments developed to establish observation networks through in-situ measurements in extreme environments. Currently, terrestrial penetrators are powered by primary batteries, and their demonstrated operational duration is limited to at most 90 days, which is a disadvantage for network-based environmental monitoring. To overcome this limitation, integrating with solar panels and secondary batteries is essential. This study investigates power resource management for penetrators through numerical simulations, focusing on extending operational duration. The results highlight the influence of solar panel size and deploying inclination angle on observational coverage, revealing that constraints in panel size and tilt necessitate improvements to the existing structural design. To address this, an “aft-body” configuration—allowing the solar panel to remain exposed above ground surface after penetration—was adopted, and penetration tests were conducted under simulated lunar and icy satellite conditions. Although the separation mechanism for the aft-body is still under development, achieving stable penetration with this design is expected to significantly enhance the feasibility of long-term observations.

Hollow cathodes are widely used for the discharge and neutralization of electrical propulsions. A number of studies have been conducted on hollow cathodes employing LaB₆ as an emitter, but the temperature characteristics of these cathodes when operating at higher temperatures than dispenser cathodes are not yet fully known. In this study, the temperature characteristics of the LaB₆ hollow cathode and the internal plasma characteristics in emitter region are investigated by simultaneously measuring the externally measurable temperature of the orifice plate and the temperature and plasma on the internal emitter region. Measurements at four orifice diameters from 0.5 to 3.0 mm showed a transition from small diameters operation mode of 1.0 and 2.0 mm where heating at the orifice is dominant to the large orifice diameter mode where heating at the emitter is dominant.

The MicroOrbiter-1 (MO-1) CubeSat mission demonstrates the potential of LoRa (Long Range) technology to enable international collaboration in the satellite-based IoT applications. This study summarizes the mission achievements attained through a store-and-forward satellite by viewing the data transmitted and retrieved between Ground Sensor Terminals (GSTs) and the MO-1 satellite from partners in Japan, Indonesia, and the United States. The GSTs operating in the 920 MHz band were applied to transmit seismic data, the very important data platform to assist Indonesia's disaster management of earthquakes, tsunami and volcanic activities. In the USA, it enabled encrypted data freely transmitted, its encryptions securely stored and forwarded via MO-1 without compromising user confidentiality. In Japan, GPS data, temperature and humidity data were sent and received despite limitations set against transmission power. This collaboration was instrumental in the validation of reliability using LoRa-based CubeSat communication. With the capability of transmitting and forwarding user-specific encrypted data, privacy has been ensured in that manner whereby only the originators of such data can interpret the information encoded. That thus leads to the satellite's role in enabling secure and decentralized IoT networks. It is a mission emblematic of the benefit of international collaboration, whereby each partner contributes to and benefits from the shared infrastructure. The results create the foundation for future multinational CubeSat missions in addressing challenges on a global scale, furthering IoT applications, and providing safe, efficient data exchange across borders.

Thanks to recent technological advancements, micro-spacecraft are becoming viable for deep space exploration. Given their limited propulsion capabilities and launch opportunities, utilizing rideshare launches to intermediate orbits such as halo orbits around Earth-Moon Lagrange points is a practical approach. This study proposes a trajectory design method for micro-spacecraft missions from the Lunar Gateway’s orbit, a near-rectilinear halo orbit (NRHO) around EML2, to a near-Earth asteroid (NEA). The method employs a lunar flyby to flexibly control the Earth escape direction while maintaining a low total delta-V (DV). The trajectory is divided into three phases: from the Gateway to an apogee, from the apogee to a lunar flyby, and from the lunar flyby to an asteroid flyby. Initial solutions are generated using a grid search, followed by trajectory optimization via a multiple shooting method. As a case study, trajectories to the asteroid Phaethon were designed. The results demonstrate that a spacecraft can reach the asteroid with a total DV of less than 20 m/s, which is feasible with current micro-spacecraft propulsion systems.

For future lunar habitation activities, securing water resources at appropriate locations on the lunar surface will be necessary. Thus, we believe that acquiring quantitative and high-spatial-resolution data on the water distribution on the lunar surface is essential. To realize this, we developed a prototype water probe that can frequently measure the water content in the regolith on the lunar surface with high sensitivity (0.1 wt%) and good reproducibility. In this study, we introduce the Tunable Diode Laser Absorption Spectroscopy (TDLAS) technology, which is a core technology showing precise and robust measurement performance. We report a method for measuring the water content in regolith using this technology and evaluate its performance.

Accurate simulation of supercritical and transcritical fluid flows remains challenging due to the occurrence of spurious pressure oscillations in fully conservative numerical formulations. Although alternative approaches such as pressure evolution equations and double-flux methods have been proposed to alleviate this issue, they often compromise energy conservation or require additional complexity in the governing equations. In this study, we evaluate a hybrid numerical method that dynamically switches between the fully conservative equation and the quasi-conservative equation based on local flow conditions. This approach aims to achieve a balance between pressure oscillation suppression and conservation fidelity without introducing extra governing equations. The performance of the hybrid method is evaluated through a series of benchmark problems, including one-dimensional advection, two-dimensional advection, and two-dimensional jet flows under supercritical and transcritical conditions. Numerical results demonstrate the capability of the evaluated method to alleviate spurious pressure oscillations while preserving energy conservation.

In recent years, numerous formation flying missions with multiple nanosatellites have been actively proposed. Due to strict constraints on mass, volume, and power consumption, minimizing propellant use is preferable for maintaining satellite formation in the disturbance environment of Low Earth Orbit (LEO). However, conventional active relative orbit control requires propellant to suppress disturbances, specifically J2 perturbations, in order to maintain formation in the out-of-plane direction. To address this technical issue, multiple studies have investigated a propellant-free method for maintaining satellite formations by utilizing the Lorentz force generated through the interaction between charged satellites moving at orbital velocity and the geomagnetic field. Conventional satellite charging methods primarily employ electron guns, which require large power consumption to emit electrons. While conventional charging methods have focused on high-power electron guns to emit electrons from satellites, this research exploits space plasma environment in LEO by applying active voltage control to satellite panels, thereby enabling collection with electrons with low power consumption. This research investigates the charging characteristics of satellite panels under high-voltage application through analytical modeling and numerical simulations.

In recent years, numerous missions based on the formation flying of multiple small satellites have been proposed. Maintaining formation flying requires accurate relative position determination, which is typically achieved using GPS and cameras. However, these methods are limited by available orbits and observation directions. Therefore, this research focuses on relative position determination using radio waves for inter-satellite communication, as this approach is less constrained by orbit or direction. In addition, this method has the advantage of low power consumption and the ability to reuse existing on-board equipment. Since the received radio wave strength varies due to antenna deflection, on-orbit calibration is necessary to improve the accuracy of relative position determination. As satellites cannot arbitrarily change their orbit and attitude, reference points are unavailable, which complicates calibration across all directions. This study addresses this issue by focusing on the direction where antenna directivity is less affected by deflection and by representing the directivity using a Legendre polynomial.

Despite looming concerns over the long-term sustainability of Earth’s orbital environment, rapid satellite deployment continues to heighten collision risks and traffic congestion, while effective economic incentives remain underdeveloped. This study investigates a fiscal mechanism that integrates a launch tax with subsidies to accelerate Active Debris Removal (ADR). Through a system dynamics (SD) modeling framework, we conduct counterfactual ("what-if") scenario analysis of historical debris trajectories and extend the assessment to a long-horizon setting to examine the implications of different policy introduction timings and transmission mechanisms. The model dynamically adjusts tax rates to meet ADR financing needs in response to launch activity, debris accumulation, and removal targets. Results demonstrate that early policy implementation significantly curbs debris accumulation, whereas delayed intervention results in steeper tax burdens and greater orbital risk. Long-horizon simulations and policy decomposition analyses further show that launch taxation alone materially moderates the growth of object accumulation, while the addition of ADR improves debris outcomes, although earlier debris suppression can partially restore launch incentives by easing the congestion-related launch restraint. These findings indicate that policy delay matters most clearly for long-run debris-risk control, but the effect of earlier intervention is mediated by endogenous interactions among launch taxation, ADR, and launch-restraint feedbacks.

On the lunar surface, dust contamination presents a critical challenge to photovoltaic systems, as fine regolith particles adhere strongly and degrade performance. To address this issue, this study investigates a hybrid dust mitigation approach that combines passive polydimethylsiloxane (PDMS) surface coatings with active electrostatic removal using ultraviolet (UV) irradiation, electron beam exposure, and a DC-biased grid. Under controlled vacuum conditions (< 5 × 10⁻³ Pa), experiments were carried out in two phases: dust deposition and cleaning. During deposition, silica microsphere dust was electrostatically charged and deposited onto solar cells. Subsequently, in the cleaning phase, the dust was charged by a 300 eV pulsed electron beam under continuous UV illumination, while a +500 V grid generated an electrostatic field that attracted and removed the charged particles. The results demonstrated that the active technique alone restored only ~50% of the original power, whereas the hybrid method with PDMS coating achieved up to 92% recovery and an image-based cleaning efficiency exceeding 90%. Crucially, the PDMS coating reduced both adhesion and contact forces, allowing the hybrid technique to remove a much larger fraction of dust even under identical exposure times and parameters.

The understanding of hypersonic rarefied gas dynamics is important for very low earth orbit (VLEO) satellites and so forth, and the development of ground test equipment is crucial so as to improve the reliability of numerical results for rarefied gas flows. A hypersonic rarefied wind tunnel (HRWT) has been developed at JAXA (Japan Aerospace Exploration Agency),. The HRWT is capable of generating hypersonic flows in rarefied flow regime. However, test conditions are currently limited to near-continuum flow regime, where Knudsen number with 5 mm is basically lower than 0.3. In order to improve the rarefaction level in HRWT, we first studied nozzle core flow properties with a variation of mass flow rate and ambient pressure in this work. Second, we investigated test conditions with low mass flow rates by carrying out pressure measurements in HRWT.

Minimizing total ∆V is crucial for efficient spaceflight mission planning, formulated mathematically as an ℓ¹-optimal control problem. Traditional approaches discretize the trajectory and apply Newton methods to find locally optimal controls. However, these methods often fail in non-convex spaces. Prior efforts based on Newton methods produced high-fidelity trajectories that were locally optimal, but did not achieve global optimization. To address this issue, the present paper proposes a novel global search algorithm grounded in Newton methods, interpreting the algorithm as a dynamical system. By tuning parameters to operate near the stability-instability boundary, the algorithm enables stochastic transitions across local optima, facilitating global search. Experimental results on optimal transfer problems between Earth-Moon Lagrange points demonstrate that the semi-stable update rule enables escape from suboptimal basins and convergence to multiple locally optimal trajectories.