A5022066530- Combustion and Detonation Processes
- Energetic Materials and Combustion
- Fire dynamics and safety research
- Risk and Safety Analysis
- Computational Fluid Dynamics and Aerodynamics
- Combustion and flame dynamics
- Rocket and propulsion systems research
- Structural Response to Dynamic Loads
- Gas Dynamics and Kinetic Theory
- Electromagnetic Launch and Propulsion Technology
- Plasma and Flow Control in Aerodynamics
- Laser-Plasma Interactions and Diagnostics
- Advanced Combustion Engine Technologies
- High-Velocity Impact and Material Behavior
- Radiative Heat Transfer Studies
- Particle Dynamics in Fluid Flows
- Aerodynamics and Acoustics in Jet Flows
- Fluid Dynamics Simulations and Interactions
- Fluid Dynamics and Turbulent Flows
- Nuclear Issues and Defense
- Fluid Dynamics and Heat Transfer
- Guidance and Control Systems
- Earthquake Detection and Analysis
- Refrigeration and Air Conditioning Technologies
- Granular flow and fluidized beds
Keio University
2016-2025
Kyoto first Red Cross hospital
2023
Tama University
2023
The University of Tokyo
2023
Tokai Gakuen University
2020
Kawasaki Heavy Industries (Japan)
2008
Keiai University
2004
Yamaguchi University
1999-2000
Kyoto Prefectural University of Medicine
1999
Institute of Space and Astronautical Science
1994-1995
A Lagrangian approach was proposed to analyze induction and reaction times in the cellular gaseous detonation. Two-dimensional simulations an argon-diluted non-diluted hydrogen-based mixtures were performed with detailed chemistry, along particle trajectories. The distribution of inside cell significantly different between Eulerian perspectives, latter showing non-monotonic behavior. Preferential thermodynamic paths laid Rankine–Hugoniot curve behind transverse waves (TW). All particles...
Gas dynamics in a simplified pulse detonation engine (PDE) was theoretically analyzed. A PDE as straight tube with fixed cross section. One end of the closed, namely, this thrust wall, and other open. wave initiated at closed simultaneously started to propagate toward open end. When broke out from end, rarefaction This reflected by By considering be self-similar analysis interference between its reflection we analytically formulated decay portion pressure history (thrust wall) without any...
Geometric throats are commonly applied to rocket combustors increase pressure and specific impulse. This paper presents the results from thrust measurements of an ethylene/gas-oxygen rotating detonation engine with various throat geometries in a vacuum chamber simulate varied backpressure conditions range 1.1–104 kPa. For throatless case, channel area was regarded be equivalent area, three throat-contraction ratios were tested: 1, 2.5, 8. Results revealed that combustor approximately...
The rotating detonation engine is a propulsion system that obtains thrust using continuously existing waves. A combustor usually has an annular shape allows waves to propagate in the circumferential direction. In this study, we used disk-shaped with combustion chamber flat-plane glass walls observe structure of phenomena. Self-luminescence, shadowgraphs, and schlieren visualization experiments were performed compared. Results revealed propagating mixture layer three gases, fuel, oxidizer,...
To create a new flyable detonation propulsion system, engine system (DES) that can be stowed in sounding rocket S-520-31 has been developed. This paper focused on the first flight demonstration space environment of DES-integrated rotating (RDE) using S-520-31. The result was compared with ground-test data to validate its performance. In experiment, stable combustion annulus RDE plug-shaped inner nozzle observed by onboard digital and analog cameras. With time-averaged mass flow [Formula: see...
This study evaluated the propulsion performance of a nozzleless, cylindrical rotating detonation engine (RDE). Using mixture, RDE was tested in low-back-pressure environment at propellant mass flow rates . In high-speed imaging self-luminescence within combustor, luminous regions were observed above Measured pressure distributions suggest that burned gas reached sonic velocity combustion chamber outlet. paper proposes structure internal and confirms calculated distribution based on close to...
Rotating detonation engines (RDEs) have been actively researched around the world for application to next-generation aerospace propulsion systems because combustion has theoretically higher thermal efficiency than conventional combustion. Moreover, cylindrical RDEs simpler combustors, further miniaturization of combustors is expected. Therefore, in this study, with aim applying space systems, a RDE converging–diverging nozzle was manufactured; combustor length [Formula: see text] changed 0,...
Performance analyses of pulse detonation rocket engines (PDREs) were numerically studied, focusing on partialfill effects at ground tests. The initial detonable mixture, inert gas, fuel-fill fraction, equivalence ratio, and temperature gas changed as governing parameters. simulation results compared against those previous studies agreed well with them. indicated that the mass fraction mixture to total was predominant factor for specific impulse partially filled PDREs. Based numerical...
The shock-induced combustion with periodic unsteadiness around a projectile fired into hypersonic flows is numerically studied. mechanism of clarified using an x-t diagram the flow variable on stagnation streamline. frequencies obtained quantitatively agree experimental observations. key parameters which are responsible in triggering instability and determining frequency discussed by time integration species equations. result indicates that induction time, heat release, concentration...
In the present research, we experimentally verified partial-fill effect in a multi-cycle pulse detonation rocket engine (PDRE).Intermittent thrust of PDRE was measured by using spring-damper mechanism that smoothes this intermittent time direction.The mass flow rates were assessed gas cylinder pressure or difference measurement.The maximum specific impulse 305±9 seconds at an ethylene and oxygen propellant fill fraction 0.130±0.004.When greater than 0.130, increased as partial decreased.When...