Masakatsu Nakano

Miniature water ion thrusters provide CubeSats with orbit transfer capability and enable missions that require high [Formula: see text]. Erosion on the grids is critical for the thruster’s lifetime and propellant utilization efficiency. It was shown numerically that the nonuniformity of the plasma in the discharge chamber causes inhomogeneous erosion of the grid. However, deviation from experiments was still a problem. In this study, we evaluate the performance and the change in the grid erosion of the miniature water ion thruster during 1000 h of operation. The propellant utilization efficiency decreases from 16% to 12% due to grid erosion. The contamination effect of the antenna and magnets on the performance of the thruster was less than that due to the change of grid geometry. In terms of grid erosion, in the center and the periphery of the grid where the ion beam density is low, the direct impingement of the ion beam is dominant, and the diameter of the accelerator grid eroded from 0.31 to 0.52 mm. On the other hand, in the middle part of the grid, where the ion beam is dense, charge exchange ions dominate the erosion, and the diameter was eroded from 0.31 to 0.39–0.45 mm in hexagonal.

Molecular dynamics simulations were performed to estimate the composition of deposited products from alumina ablation plume at the atomic level. The parameters considered were the Al/O ratio of ablation plume and the temperatures of ablation plume and recovery plate surface. Simulation results showed that oxygen-deficient alumina was formed on the recovery plate surface, whose Al/O ratio was in good agreement with the composition of oxygen-deficient alumina obtained in the experiment. The simulation results also showed that the Al/O ratio of the deposited layer could be increased when the Al/O ratio of the alumina ablation plume was larger and the temperature of the recovery plate surface was higher. These results suggest that the Al/O ratio can be increased by repeated process of laser ablation and deposition of ablation plume, and the relationship between the Al/O ratio and the number of repetitions of this process was modeled and evaluated.

<p>In order to improve ion engine's performance, the neutralization performance with two field emission cathodes in an ion engine is investigated. The neutralization performance is evaluated by the potential difference between cathode and ground, using a 100μN class ion thruster developed at Kyushu University and two 50×50mm² field emission cathodes with carbon nanotube emitter. The potential difference between cathode and ground is not only determined by cathode electron supply capacity and the position of the cathodes but also foot print of the neutralizers, it would be due to the space charge limitation. That is, the potential difference between cathode and ground is improved with increase in total emission current, and that with a single field emission cathode at emission current of 6mA is -20V, on the contrary, that with two field emission cathodes is -12V. </p>

<p>The grid design of ion thrusters is currently based on trial and error development and durability tests. These methods make it difficult to rapidly develop optimal thrusters for diverse missions. As a way to address this problem, this paper provides a method of designing grids using numerical simulation with a genetic algorithm. Screen grid thickness, screen grid hole diameter, accel grid thickness, accel grid hole diameter, accel grid potential, and grid gap are optimized for long-life grids. The JIEDI tool is a powerful tool to design grid systems; it is used here as a fitness function to evaluate the lifetime of the grid in an ion thruster. As a result of the optimization, the lifetime of the model developed using the genetic algorithm is about 6.5 times longer than that of the experimental model. The optimization of grid parameters using the Genetic Algorithm and the JIEDI tool shows promise for future design tasks.</p>

<p>  A panel discussion was held to discuss the technical challenges in orbital transfer vehicle (OTV) development using electric propulsion. Electric propulsion is advantageous in the transportation of SSPS if the solar array panel in the payload can be used in the OTV. Argon is the candidate for the propellant, and the technical issues for the development of high-power argon-propellant thruster are discussed. The collaboration with the ground launch vehicle is indispensable to attain the SSPS, in addition to the cross-field collaboration for the optimization of the OTV. The necessity of technology demonstration missions and technical roadmap is reconfirmed to promote the collaboration among the researchers.</p>

<p>The two-grid ion optics of a water propellant miniature ion propulsion system (MIPS) was numerically studied by full-aperture ion-optics simulation considering an inhomogeneous plasma source in the discharge chamber. Three dominant ion species (H₂O⁺, OH⁺, and H⁺) of water vapor plasma were included in the model. Their trajectories were tracked considering charge-exchange and elastic collisions between ions and neutral water molecules. The model was validated by experiments using the MIPS with two grids having 295 apertures each. The calculated accelerator impingement current agreed reasonably well with those obtained from the experiment. To increase its low propellant utilization efficiency, a new grid design method was introduced. This successfully improved neutral confinement by decreasing the diameter of the accelerator grid apertures without causing direct impingement of ions by optimizing the grid thickness distribution using the Particle In Cell-Monte Carlo Collision (PIC-MCC) calculation results.</p>

<p>  A large amount of propellant would be needed to realize SSPS, as it requires large amount of mass transportation. We propose sublimable substances as a new propellant, which is solid under room temperature and normal pressure, considering that Xenon is expensive. By using solid propellant, orbital transfer system can be reasonable because it requires no high pressure tank. In this report, we discuss whether sublimable substances can be alternative propellant by evaluating the performance of ion engine using sublimable substances.</p>

<p>  The economic feasibility of SSPS depends strongly on its transportation cost to Geostationary Earth Orbit. This study discusses the transportation cost of SSPS by using orbital transfer vehicles driven by high power electric propulsion. Cost analysis showed that the transportation cost of 10000 ton SSPS will be around 1 trillion yen based on the current-state-of-art technology and the future cost reduction predicted by the experience curve of the aerospace industry. The transportation cost can be reduced to around 500 billion yen when delivering 10 SSPSs to GEO. The required capability of electric propulsion is an output power of 100 kW using argon as a propellant.</p>

A micro-thruster using stacked solid propellant pellets for nano-satellites has the high versatility. This thruster can generate arbitrary total impulse by changing the number of installed pellets. In order to install the thruster in a nano-satellite, it is necessary to clarify conditions of stable operation. We changed materials and manufacturing methods of the thruster, and measured the combustion chamber pressure and the burning rate. A certain condition of the thruster achieved stable operation. A difference between the burning rates of the propellant pins caused unexpected combustion chamber pressure rise. A silicone rubber contributed stable operation because it made the burning rates of the propellant pins even.

<p>  The selling price of electricity depends strongly on the transportation cost of SSPS from the ground to Geostationary Earth Orbit, which greatly influences the economic feasibility of the SSPS. This study discusses the transportation cost of the SSPS by using reusable orbital transfer vehicles (OTVs) driven by high power electric propulsion. Cost analysis showed that the transportation cost of 10000 ton SSPS is around 1.27 trillion yen, which is 44% lower than that of using ground launch vehicles (GLVs) alone. The optimum specific impulse of electric propulsion is around 2660 s with the use of argon as a propellant. A sensitivity analysis showed that the innovation of reducing the initial construction costs of the GLVs and OTVs by 75% is necessary to meet the target transportation cost within 0.3 trillion yen.</p>

The burning rates of small sized Boron Potassium Nitrate (BKNO₃) pellets (3.2-mm-dia., B:KNO₃: binder = 28:70:2) under low pressures were experimentally measured; they increased with pressure (r = 71.1p^0.589[mm], p[MPa]) at an ambient pressure higher than 5 kPa and exhibited little dependence on pressure (r = 3.0±0.5mm/s) at an ambient pressure lower than 5 kPa. The measured pressure increase after reaction becomes smaller under 1 kPa and the photographs taken by a high-speed camera showed unreacted particle emissions at 0.2kPa. These results suggest that the reaction of BKNO₃ at low pressures takes place in condensed phase. From thrust stand measurement, the pressure working on the combustion surface is estimated to be 1.2kPa, which implies that the dependence of burning rates on ambient pressure vanishes under around 1kPa.

<p>A micro-solid rocket as the propulsion system for 1&ndash;10 kg-class micro-spacecraft is proposed here. The micro-solid rocket uses a boron/potassium nitrate pellet as propellant and its total impulse is about 1.5 Ns. Higher total impulse is needed for a propulsion system on small spacecraft to perform advanced space missions such as sample return, formation flight, and active debris removal. To increase the total impulse, it is necessary to increase the propellant mass. However, there is a difficulty in producing new sizes of solid propellant. The author designed a 20&ndash;50 Ns-class micro-solid rocket which uses a stack of existing multiple B/KNO₃ pellets. The side of the propellant pellets was sealed with epoxy resin to prevent an abnormal combustion chamber pressure rise. As a result, all the propellant was burned without an abnormal pressure rise in all combustion tests.</p>

Numerical studies were conducted on a small ion engine system called μ1, which was developed for micro-spacecraft propulsion. For its miniaturized application, the μ1 ion engine is very small in size and its beam current and propellant utilization efficiency are lower than those of standard-sized ion engines. As it operates in a low perveance range with a molybdenum two-grid system, the acceleration grid experiences very severe erosion, and the extraction ion beam performance degrades substantially because of the scattering of highly diverted ions and neutrals and the impact of backstreaming charge exchange ions. In this study, the decrease in the propellant utilization efficiency and the lifetime of the μ1 ion engine were firstly assessed using the JIEDI tool, considering the effect of the change in extraction ion beam performance. Lifetime analysis showed that the μ1 grid system can operate over its required lifetime of 10000 h, however, the propellant utilization efficiency decreased significantly from 0.42 at the beginning of life to 0.30 at the grid structural failure of 27000 h. The agreement of grid erosion patterns after the first 150-h of operation between the simulation and the experiment supports the validity of the analysis. In addition to the lifetime assessment, the grid parameters were parametrically searched to achieve both high propulsion performance and long life. One of the strategies is to double the diameters of the screen and accelerator grid holes as well as the accelerator grid thickness, keeping the separation distance between the screen and accelerator grids unchanged.

Cost evaluation of a solar power satellite (SPS) in-space transportation using Hall thruster propulsion systems is conducted to obtain specific characteristics of Hall thrusters that yield significant transportation cost reduction. Therefore, the transportation scenario is optimized in the first place: the choice of reusing or disposing orbit transfer vehicles (OTVs), and the power of the propulsion system. The result shows that the case of reusing OTVs is superior to the case of disposing them, because the OTVs manufacturing costs are predominant in the total cost. In addition, the power has only limited influence on the cost. Further, specific characteristics of Hall thrusters are discussed, which are necessary to achieve a target cost of $3.75 billion (300 billion yen) for the sum of two costs: the in-space transportation cost from a low Earth orbit (LEO) to a geostationary Earth orbit (GEO) and the propellant launch cost from Earth to LEO.