Ikkoh Funaki

This paper explores the innovative direction control of rotating detonation waves in rotating detonation engines (RDEs) by adjusting the ignition location and employing helical combustors with a sinusoidal cross section. In our experimental setup, we conducted 25 combustion tests using two distinct combustor geometries, each featuring different helical profile directions. The following conclusion drawn from the results: when the ignition was positioned 30.6–46.0 mm from the inlet, the detonation wave direction was invariably influenced by the helical direction. This correlation was statistically significant, with an occurrence probability (assuming a random direction probability of 0.5) being [Formula: see text], far exceeding the 0.05 significance level. Furthermore, the helical combustors generated a measurable torque due to the pressure differentials created by shock waves within the combustor. This torque, recorded between [Formula: see text] at a mass flow rate of 28.5–29.0 g/s, indicates the potential of power extraction from the combustor. Notably, the torque direction was also controllable via the helical direction. This study presents a significant advancement in propulsion technology, demonstrating a novel method to control detonation wave direction and torque generation in RDEs through helical combustor design, paving the way for more efficient and controllable propulsion systems.

A coupled cylindrical rotating detonation engine (RDE) with two cylindrical RDEs (both combustors had a combustor inner diameter of 23 mm and an axial length of 30 mm) placed next to each other was tested for rocket clustering application. The objective of the experiment was to achieve two-engine synchronized initiation with a single igniter. Experiments were conducted on the inner wall of the combustors with different connecting-hole diameters and wall heights to evaluate the ignition delay time, combustion mode, and propulsion performance. The propellants were gaseous ethylene and oxygen, and experiments were conducted under constant conditions of mass flow rate ([Formula: see text]), equivalence ratio ([Formula: see text]), and backpressure (approximately 10 kPa). When the two combustion chambers were completely separated by a wall, ignition occurred with a time delay of 260 ms in the chamber without an igniter. However, when a large hole ([Formula: see text] diameter) was placed in the wall separating the two combustion chambers, synchronous initiation was successful. Synchronous initiation was also successful when the wall height was lowered (7-mm height). Under both conditions, the same level of specific impulse was achieved as for RDEs operating at the same mass flux.

There are few experimental studies on rotating detonation engines (RDEs) with liquid propellants. This study reveals the static thrust performance of a cylindrical RDE with ethanol and liquid nitrous oxide as propellants under atmospheric pressure. This RDE had an inner diameter of 40 mm, a maximum combustor length of 230 mm, a nozzle contraction ratio of 1.7, and a nozzle expansion ratio of 9.1. Nineteen experiments were conducted at total mass flow rates of [Formula: see text], mixture ratios of 3.6–5.9, and combustion pressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of [Formula: see text] (approximately 80% of the theoretical detonation velocity, [Formula: see text]), maximum thrust at sea level of 294 N, and maximum specific impulse at sea level of 148 s. In addition, the maximum characteristic exhaust velocity, [Formula: see text], was [Formula: see text], which was 99% of the theoretical value. The characteristic length of the combustion chamber at this time was 0.15 m. Since conventional rocket combustion requires 1.57 m to achieve the same [Formula: see text] efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.

Magsail is a space propulsion system using the interactions between the solar wind and the magnetic field generated by the onboard coils. Magnetoplasma sail is a propulsion system that increases thrust by expanding the magnetosphere through plasma injection from the spacecraft. There are two mechanisms on the magnetospheric inflation: method using frozen-in of magnetic field to carry magnetic field lines by high dynamic pressure plasma and method using the diamagnetic current by thermal plasma, which is called the ring current. We investigated the effect of the dynamic pressure and thermal pressure on the MPS thrust performance used electromagnetic hydrodynamic simulation. It was shown that the ring current is enhanced by adding dynamic pressure to the thermal plasma and increases thrust gain. The high thrust gain over 2.25 was obtained at β_th = 0.5 - 2 and β_k = 4 - 8. However, the thrust is reduced because the super magneto acoustic wave region is generated in the magnetosphere, which prevents the propagation of thrust in large β conditions. The wide parameter survey reveals injection plasma parameter regions where thrust reduction is restrained and high thrust gain is obtained.

KOSEN-1, the first artificial satellite developed by 10 National Institutes of Technology (KOSEN), was launched in November 2021. KOSEN-1 is a 2U size (20×10×10cm) CubeSat orbiting in a low orbit at an altitude of about 575km. In order to extending the operating time of KOSEN-1, a propulsion system for increasing altitude is necessary. In the present study, we designed and manufactured a compact Pulsed Plasma Thruster (PPT) that operates with a power of several Watts. We performed an experiment for demonstration of PPT in vacuum chamber. As a result, we showed that the ignition voltage at which the main discharge of 1.4 kV occurs stably repeatedly is around 10 kV and observed the generation of plasma in the PPT. The velocity of PPT plasma of 19 km/s was measured by TOF method. The PPT plasma temperature and density were measured 3.04 eV and 2.81×10²¹ m^-3 by Double Probes respectively. At the main discharge voltage of the PPT of 1.4kV, the impulse bit measured using the target method was 65 µNs.

Abstract This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship. Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

Magsail is a space propulsion system using the interactions between the solar wind and the magnetic field generated by the onboard coils. Magnetoplasma sail is a propulsion system that increases thrust by expanding the magnetosphere through plasma injection from the spacecraft. There are two mechanisms on the magnetospheric inflation: method using frozen-in of magnetic field to carry magnetic field lines by high dynamic pressure plasma and method using the diamagnetic current by thermal plasma, which is called the ring current. We investigated the effect of the dynamic pressure and thermal pressure on the MPS thrust performance used electromagnetic hydrodynamic simulation. It was shown that the ring current is enhanced by adding dynamic pressure to the thermal plasma and increases thrust gain. The high thrust gain over 2.25 was obtained at β_th = 0.5 - 2 and β_k = 4 - 8. However, the thrust is reduced because the super magneto acoustic wave region is generated in the magnetosphere, which prevents the propagation of thrust in large β conditions. The wide parameter survey reveals injection plasma parameter regions where thrust reduction is restrained and high thrust gain is obtained.

Rotating detonation engines (RDEs) have been actively researched around the world for application to next-generation aerospace propulsion systems because detonation combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, because cylindrical RDEs have simpler combustors, further miniaturization of conventional combustors is expected. Therefore, in this study, with the aim of applying RDEs to space propulsion systems, a cylindrical RDE with a converging–diverging nozzle was manufactured; the combustor length [Formula: see text] was changed to 0, 10, 30, 50, and 200 mm; and the thrust performance and combustion mode with the different combustor lengths were compared. As a result, four combustion modes were confirmed. Detonation combustion occurred with a combustor length of [Formula: see text]: that is, a converging rotating detonation engine. The thrust performance of this engine was 94 to 100% of the theoretical rocket thrust performance, which is equivalent to the thrust performance of conventional rocket combustion generated at [Formula: see text]. This study shows that detonation combustion can significantly reduce engine weight while maintaining thrust performance.

This paper reports on our study to develop a magnetic circuit with a discrete outer coil and test its performance in a Hall thruster. The results were compared to those of a Hall thruster on board the Engineering Test Satellite-9 (ETS-9), which used a cylindrical outer coil. At the normalized magnetic field strength of 0.58 for the 6 kW operation using a six poles discrete outer coil, the thrust, discharge current, Isp, thrust-to-power ratio, thrust efficiency, and discharge oscillation were 386.7 mN, 19.0 A, 1890 s, 66.9%, 62.0%, and 13.5% respectively. The Hall thruster with the new magnetic circuit achieved equivalent performance (including discharge stability) to the Hall thruster on board the ETS-9. Further, the study investigated the effect of the number of poles on the performance of the Hall thruster: the performance with the three pole coil was lower than that with the six poles coil. This paper also reports on the effect of pole numbers on utilization efficiency. Mass and current utilization efficiency were more affected by the magnetic field generated by the three poles outer coil.

To create a new flyable detonation propulsion system, a detonation engine system (DES) that can be stowed in sounding rocket S-520-31 has been developed. This paper focused on the first flight demonstration in the space environment of a DES-integrated rotating detonation engine (RDE) using S-520-31. The flight result was compared with ground-test data to validate its performance. In the flight experiment, the stable combustion of the annulus RDE with a plug-shaped inner nozzle was observed by onboard digital and analog cameras. With a time-averaged mass flow of [Formula: see text] and an equivalence ratio of [Formula: see text], the RDE generated a time-averaged thrust of 518 N and a specific impulse of [Formula: see text], which is almost identical to the ideal value of constant pressure combustion. Due to the RDE combustion, the angular velocity increased by [Formula: see text] in total, and the time-averaged torque from the rotational component of the exhaust during 6 s of operation was [Formula: see text]. The high-frequency sampling data identified the detonation frequency during the recorded time as 20 kHz in the flight, which was confirmed by the DES ground test through high-frequency sampling data analysis and high-speed video imaging.

Abstract This paper presents a control algorithm for ensuring the stable operation of a 6 kW Hall thruster. The Japan Aerospace Exploration Agency is developing a “wide-range” power processing unit (PPU) to power a 6 kW Hall thruster system, a candidate for all-electric satellites. The PPU provides discharge oscillation, power, and discharge current control. The increments in the control loop are fixed, so the PPU digital controller does not need to calculate them, and the algorithm’s computational complexity is minimal. Results of integration tests on the 6 kW Hall thruster and the PPU breadboard model showed that the three control functions run correctly. The total controlled PPU power was adjusted in the range of 3.45–3.55 kW under constant power control with a target of 3.5 kW. Oscillations of ± 0.05 kW around the target power are acceptable. 70% of the peaks in the discharge current acquired for 100 ms exceeded the limit of 1.5A. Discharge oscillation control varied the coil magnetic field and reduced them to 0%. The PPU successfully controlled the coil current, reducing the discharge current from 13.7 A to 13 A. The propulsion efficiency increased from 59% to 60.5%.

<p>The ion energy angle distribution and its relationship to plasma parameters for spot and plume modes are elucidated for a LaB₆ hollow cathode with a radiative heater. Measurements were conducted using a retarding potential analyzer (RPA) and a single Langmuir probe. The ion energy distribution function (IEDF) characteristics showed different tendencies in the current density and mass flow-rate dependence under different plasma modes. The IEDF peak potential for the spot mode varied from 16 to 23 V with increasing current density, and the IEDF peak potential for the plume mode varied from 16 to 32 V with decreasing mass flow rate. Considering angle dependency of ion energy, when the observation angle was changed from the radial direction to the axial direction, the IEDF peak potential increased from 29 to 40 V for the plume mode (10 A, 10 sccm) and increased slightly from 16 to 18 V for the spot mode (20 A, 30 sccm). The probe measurement analysis revealed that the IEDF peak energies are the same as, or exceed, the plasma potential and have a qualitative correlation with the electron temperature spatial distribution.</p>

<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>  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>  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>