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Ammonium dinitramide (ADN)-based energetic ionic liquid propellants (ADN-EILPs) present considerable potential as monopropellants due to high energy density, high thermal/chemical stability, and low toxicity. Consequently, a chemical thruster system based on ADN-EILPs can be employed in the development of high-performance ultrasmall/small satellites. However, the characteristics of the ADN-EILPs present various problems in their ignition during the thruster development. To solve this problem, the authors have previously achieved the continuous-wave (CW) laser ignition of ADN-EILPs using laser absorbers composed of carbon wools. However, the conventional propellant injection of ADN-EILPs with carbon wool using high-pressure gas faces several limitations. Therefore, a propellant feed system suited for the ignition method is proposed here, as well as a conceptual model of a 0.5 U/0.5 N-class thruster operated by the CW laser ignition of ADN-EILPs ([Formula: see text]). Additionally, an attempt is made to manufacture a laboratory model (LM) thruster, and its fundamental operation properties are determined. The future research implications of this study include the further observation of the combustion of the decomposed gas before its emission from the throat of the LM thruster, along with the further development of the propellant feeding system proposed in this study and the hot-fire tests performed for the feeding systems.

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.

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.

A detonation engine system is successfully demonstrated for the first time in space using sounding rocket S-520-31 of the Japan Aerospace Exploration Agency/Institute of Space Astronautical Science. Detailed flight results of an S-shaped pulse detonation engine (PDE) installed in the rocket are presented herein. The flight is conducted to confirm that the PDE and its system operate at scheduled sequences in space, confirm the reproducibility of the PDE cycle, and despin the rocket around its axis. It is confirmed that the PDE operated successfully for 14 cycles in space. The experimental plateau pressure of [Formula: see text] is [Formula: see text] of the calculated plateau pressure, which suggests that detonation occurred in 14 cycles. The pressure profiles of the cycles are similar, and the pressure integrals are [Formula: see text], confirming the excellent reproducibility of the PDE cycle. A probability statistical approach assuming a Gaussian distribution is applied to determine the average angular acceleration difference between processes of the PDE operation, mixture supply, and oxygen supply. The results suggested that the PDE despun the rocket via the thrust produced via detonation combustion, which is consistent with a quasi-steady-state model with an accuracy of [Formula: see text].

Converging–diverging nozzles are common in rocket engine systems to increase the exhaust velocity and improve thrust performance. In this study, we focused on the acceleration of subsonic burned gas without a structural throat via detonation to realize a simple and compact engine. We developed and tested a rotating detonation engine (RDE) without a throat and with a diverging channel (constant diverging angle α=5 deg). Gaseous C2H4 and O2 were used as the propellants, and the mass flow rate ranged from 62 to 134 g/s in the combustion tests under low back-pressure conditions. We measured pressure and thrust, as well as high-speed imaging of self-luminescence of the combustion and imaging of the exhaust plume. The pressure at the exit was less than one-fifth of the maximum pressure in the RDE, significantly below the value for a sonic flow. The results suggested that the exhaust flow was supersonic, with values up to Mach 1.7, without the need of a converging section within the engine. In addition to the estimated Mach number from the measured pressure, the exhaust plume images coherently indicated the existence of supersonic exhaust.

A cylindrical rotating detonation engine (24 mm diameter) with multiple injector holes on the combustor side wall for both propellant injection and cooling by injector gas flow was designed and te...

資料番号: SA1190092000

A reflective shuttling detention combustor (RSDC) is a two-dimensional combustor with two reflection walls. Unlike a rotating detonation combustor (RDC), an RSDC can visualize the entire area of the combustor via an optical technique. In addition, the RSDC can reproduce the counter-rotating detonation mode in RDC. In present study, CH* luminescence and schlieren image were observed to investigate the dynamics of the RSDC. The reflective wall distance was set to 45 mm (Type L) and 25 mm (Type S). As the results, the wave speed of 1226 ± 64 m / s in the single wave mode in Type S was 76% of that in Type L. It was found that the detonation propagation speed decreased with a reduction in the reflection wall distance. Superimposing the luminescence images on the schlieren images revealed that the mixture was mainly burned by deflagration behind shock wave. In addition, the experimental mixture fill height was in good agreement with model in which the mixture filling process was temporarily stopped by detonation and refilled at a constant speed. Using the maximum fill height obtained by the model, it was found that the detonation mode was in the region of 3 ± 0.6 of the dimensionless quantity (the reflection wall distance divided by the wave number and maximum value of the mixture fill height).

Accepted: 2017-10-12, 資料番号: SA1180308000

Particulate matter 2.5 (PM2.5), collected from ambient air in Fukuoka City, was analyzed by gas chromatography combined with multiphoton ionization mass spectrometry using an ultraviolet femtosecond laser (267 nm) as the ionization source. Numerous parent polycyclic aromatic hydrocarbons (PPAHs) were observed in a sample extracted from PM2.5, and their concentrations were determined to be in the range from 30 to 190 pg/m(3) for heavy PPAHs. Standard samples of nitrated polycyclic aromatic hydrocarbons (NPAHs) were examined, and the limits of detection were determined to be in the picogram range. The concentration of NPAH adsorbed on PM2.5 in the air was less than 900-1300 pg/m(3). Graphical Abstract ᅟ.