The rapid expansion of rideshare launches has increased the demand for orbital transfer vehicles (OTVs) and kick stages capable of providing flexible last-mile transportation for small spacecraft. Such propulsion systems require moderate thrust, high operational safety, restart capability, and compatibility with compact satellite platforms. Hybrid rockets are a promising option because they offer a balance between the simplicity and safety of monopropellant systems and the high performance of bipropellant engines. This paper introduces recent research and development activities on a hybrid kick motor aimed at ridesharing small spacecraft applications. A high-altitude test facility was developed to accurately evaluate thrust performance under low-pressure conditions, and a diffuser design based on normal shock theory was validated experimentally. A restartable ignition technique using an electrically conductive polymer was proposed, demonstrating low-power ignition suitable for small satellite power budgets. The applicability of hydrogen peroxide as a green oxidizer was also investigated. Combustion experiments showed that stable operation can be achieved without catalyst decomposition even at reduced oxidizer concentrations by employing a CAMUI-type fuel configuration. Furthermore, regenerative cooling using liquid oxidizer successfully suppressed graphite nozzle erosion during long-duration firing. These results indicate that hybrid propulsion systems using green oxidizers can provide a practical solution for safe, efficient orbital transfer in rideshare missions, enabling long-duration operations while maintaining propulsion performance and system simplicity.

This paper presents research on chemical propulsion devices for small satellites conducted by the author’s group. Conventionally, hydrazine-based monopropellant or bipropellant thrusters are widely used as onboard propulsion systems, while electric propulsion is applied to spacecraft requiring relatively large total impulse. Solid propellant motors are commonly used as apogee kick motors; however, they are rarely employed as onboard propulsion systems. Today, small spacecraft with masses as low as 100 kg are being developed by start-ups and universities. Furthermore, from the viewpoint of the space environment, post-mission disposal (PMD) has become increasingly important, as operators are required to remove end-of-life satellites to prevent the growth of space debris. Collision avoidance maneuvers (CAM) have also attracted attention, as several spacecraft have reportedly been at risk of in-orbit collisions. These circumstances have intensified the demand for compact, lightweight, and environmentally friendly thrusters capable of enabling flexible maneuvers for small satellites. The author’s group has therefore developed solid propellant thrusters, liquefied-gas thrusters, and electric-chemical dual-mode thrusters.

This paper reviews the recent development of a novel magnesium-water hybrid micropropulsion system for small spacecraft. Designed to overcome the toxicity and high-pressure storage limitations of conventional chemical thrusters, this system utilizes a safe and green metal-water reaction. First, we investigate the fundamental combustion phenomena of magnesium wires in a water vapor environment. Due to extreme heat loss and the resulting low flame temperatures, gaseous oxides rapidly condense into Condensed Combustion Products (CCP), forming large residues on the fuel surface that hinder continuous combustion. To mitigate this, we propose a mechanical vibration-induced powderization method to fragment the CCP and improve the combustion rate. Furthermore, we detail system-level engineering advancements for orbital demonstration: theoretical estimations yielding a specific impulse, a tapered combustion chamber designed to stabilize the flame position for automated wire feeding, and a regenerative cooling mechanism for thermal protection and spontaneous oxidizer vaporization. Ultimately, this review highlights the forefront of combustion dynamics and system integration essential for realizing a high-thrust, next-generation micropropulsion system.

The combustion characteristics of a hybrid thruster for small satellite propulsion were investigated through long-duration firing tests and data analysis. The thruster was designed to operate under low oxidizer-to-fuel ratio conditions to suppress nozzle erosion and enable long-duration firing for orbital maneuvering missions. Single-port firing tests were conducted to determine fuel regression characteristics, and the obtained experimental constants were applied to the design of a combustion chamber. Time histories of chamber pressure, oxidizer mass flow rate, and thrust demonstrated stable combustion over a 200 s firing duration. A simplified reconstruction technique was developed to estimate oxidizer-to-fuel ratio and fuel mass flow rate from measurable combustion data without iterative convergence calculations, significantly reducing computational cost while maintaining sufficient accuracy. Analysis of reconstructed data showed that long-duration firing introduces additional effects such as ablator consumption and post-combustion fuel vaporization, affecting the apparent oxidizer-to-fuel ratio. Residual thrust caused by post-firing vaporization under vacuum conditions was also identified through chamber pressure measurements. These results demonstrate stable long-duration hybrid combustion and highlight the importance of thermal history and post-combustion phenomena in performance evaluation and system design.

OMOTENASHI was designed to be the world's smallest Moon lander, launched by the first mission of the Space Launch System. It was equipped with two types of propulsion systems: a solid rocket motor and a pair of cold gas jet thrusters. The development of a laser-ignited 3U-sized solid rocket motor was successful. The imported thrusters, which passed various functional tests, unfortunately experienced a malfunction in orbit and resulted in mission loss. This article reviews the development processes and results, and in-orbit operational results of these propulsion systems.

The rapid increase in small satellite launches has intensified the demand for high-thrust chemical micropropulsion systems. Conventional propellants like hydrazine pose significant safety and cost challenges for small-scale missions. This study proposes and evaluates a green propellant system utilizing water and micrometer-sized aluminum powder, which offers high energy density, safety, and potential for in-situ resource utilization (ISRU). Experiments were conducted using a 0.79-liter cylindrical combustion chamber to investigate the combustion characteristics of water vapor and aluminum particles. Aluminum was dispersed using nitrogen gas and ignited via spark discharge. The combustion performance was evaluated through pressure history and reacted aluminum fraction measurements. The results demonstrated a strong correlation between the maximum combustion pressure and the reacted aluminum fraction. Successful flame propagation and significant pressure rises were primarily observed under fuel-rich conditions, specifically within an oxidant-to-fuel (O/F) mass ratio range of 0.5 to 0.8. In contrast, higher O/F ratios often resulted in local ignition followed by immediate quenching, leading to negligible pressure increases. The maximum experimental pressure reached approximately 40 % of the theoretical maximum calculated by NASA CEA. Furthermore, qualitative design considerations for key system components—including the propellant supply systems, vaporizer, combustion chamber, and nozzle—are discussed based on experimental observations. Crucial factors such as humidity control in powder tanks to prevent agglomeration, the necessity of heating vapor lines to avoid condensation, and the trade-off between nozzle throat diameter and propellant residence time are highlighted. These insights provide a technical foundation for the development of integrated chemical micropropulsion systems using water and aluminum.

Using ammonia as a carbon-free fuel and hydrogen carrier has been proposed to achieve a carbon-neutral society. However, ammonia is known for its corrosive nature, which can deteriorate the components used for ammonia combustion and related systems, such as solid oxide fuel cells. In this article, we review our recent understanding of the materials degradation behavior in an ammonia environment at high temperatures. We present a basic understanding of the degradation mechanism of nickel-base superalloys. Surface cracking was more severe at middle-range temperatures around 500 ˚C, suggesting a stronger surface reaction with ammonia than thermal decomposition. Next, we present the degradation behavior of the CoNiCrAlY bond coat and the yttria-stabilized zirconia top coat used in the thermal barrier coating system. Bond coat degradation can be suppressed by introduction of heat treatment. We observed bond coat degradation under the top coat. We discuss the diffusion kinetics and degradation mechanisms behind these observations.

In order to examine wall chemical effects caused by radical adsorption and recombination on wall surfaces under flame-wall interaction, H-atom recombination on different walls are examined by using a methane (CH₄) -air premixed flame formed in a stagnation flow near a heated wall. Quartz, alumina, magnesia and platinum are chosen as the wall materials. To ensure the identical thermal boundary condition with the different wall materials, 300 nm-thick alumina, magnesia and platinum thin films are deposited on the same quartz substrates. Gas-chromatography analysis and numerical simulation with detailed gas/surface chemistry are carried out to evaluate the near-wall H₂ concentration, which could be increased by the H₂ desorption following the H recombination on the wall. Firstly, the wall stagnation flame near the platinum, of which surface reaction mechanism is well known, are examined to verify the numerical scheme. The simulation results of the flame position and the near-wall H₂ concentrations are in accordance with the measurement data. Then, the measured H₂ concentrations near the quartz, alumina and magnesia walls with different wall temperatures T_w are compared to the simulation results with various H-atom recombination rate constants k_r. It is found that, with increase in T_w, k_r on the quartz is increased and the H-atom recombination of the quartz shifts from adsorption-recombination competition regime to adsorption-limited regime. The H₂ concentrations near the alumina are in good agreement with the numerical results without the surface reaction, indicating that the alumina is a chemically-inert material. In contrast, as for the magnesia, the near-wall H₂ concentration is decreased with increase in T_w and becomes lower than that near the inert wall, implying that H₂ decomposition rather than H recombination occurs at high T_w.