TY - GEN
T1 - Experimental Study on Truncated Conical Rotating Detonation Engines with Diverging Flows
AU - Nakata, Kotaro
AU - Ota, Kosei
AU - Ito, Shiro
AU - Ishihara, Kazuki
AU - Goto, Keisuke
AU - Noboru, Itouyama
AU - Watanabe, Hiroaki
AU - Kawasaki, Akira
AU - Matsuoka, Ken
AU - Kasahara, Jiro
AU - Matsuo, Akiko
AU - Funaki, Ikkoh
N1 - Funding Information:
This study was subsidized by a “Study on Innovative Detonation Propulsion Mechanism”, Research and Development Grant Program (Engineering) from the Institute of Space and Astronautical Science, the Aerospace Exploration Agency, by a Grant-in-Aid for Specially Promoted Research (No.19H05464), and by a “Research and Development of an Ultra-High-Thermal-Efficiency Rotating Detonation Engine with Self-Compression Mechanism,” Advanced Research Program for Energy and Environmental Technologies, the New Energy and Industrial Technology Development Organization.
Publisher Copyright:
© 2021, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.
PY - 2021
Y1 - 2021
N2 - This study focused on acceleration of propellant to supersonic speed with a compact engine. Increase in the exhaust velocity of a rocket engine leads to improvement of the thrust performance. Several combustion tests were conducted for a small rotating detonation engine (RDE) whose channel is truncated conical shaped (diverging angle (formula presented)), under low back pressure conditions. In these tests, gaseous (formula presented) were used as the propellant, and the mass flow rate were ranged from 56 to 123 g/s. The pressure ratio between the maximum value in the engine and at the exit was approximately 0.16, which was confirmed to be significantly below the critical value for a sonic flow. Namely, the exhaust flow was supersonic, even though there was no convent section in the engine. In a range of propellant mass flow rate, specific impulses were approximately 110% compared to those in a cylindrical RDE with a uniform cross-section combustor.
AB - This study focused on acceleration of propellant to supersonic speed with a compact engine. Increase in the exhaust velocity of a rocket engine leads to improvement of the thrust performance. Several combustion tests were conducted for a small rotating detonation engine (RDE) whose channel is truncated conical shaped (diverging angle (formula presented)), under low back pressure conditions. In these tests, gaseous (formula presented) were used as the propellant, and the mass flow rate were ranged from 56 to 123 g/s. The pressure ratio between the maximum value in the engine and at the exit was approximately 0.16, which was confirmed to be significantly below the critical value for a sonic flow. Namely, the exhaust flow was supersonic, even though there was no convent section in the engine. In a range of propellant mass flow rate, specific impulses were approximately 110% compared to those in a cylindrical RDE with a uniform cross-section combustor.
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U2 - 10.2514/6.2021-3657
DO - 10.2514/6.2021-3657
M3 - Conference contribution
AN - SCOPUS:85126743049
SN - 9781624106118
T3 - AIAA Propulsion and Energy Forum, 2021
BT - AIAA Propulsion and Energy Forum, 2021
PB - American Institute of Aeronautics and Astronautics Inc, AIAA
T2 - AIAA Propulsion and Energy Forum, 2021
Y2 - 9 August 2021 through 11 August 2021
ER -