Optical and thrust measurement of a pulse detonation combustor with a coaxial rotary valve

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Front shock behavior of stable curved detonation waves in rectangular-cross-section curved channels

The propagation of curved detonation waves of gaseous explosives stabilized in rectangular-cross-section curved channels is investigated. Three types of stoichiometric test gases, C2H4 + 3O2, 2H2 + O2, and 2C2H2 + 5O2 + 7Ar, are evaluated. The ratio of the inner radius of the curved channel (ri) to the normal detonation cell width (λ) is an important factor in stabilizing curved detonation waves. The lower boundary of stabilization is around ri/λ = 23, regardless of the test gas. The stabilized curved detonation waves eventually attain a specific curved shape as they propagate through the curved channels. The specific curved shapes of stabilized curved detonation waves are approximately formulated, and the normal detonation velocity (Dn)−curvature (κ) relations are evaluated. The Dn nondimensionalized by the Chapman–Jouguet (CJ) detonation velocity (DCJ) is a function of the κ nondimensionalized by λ. The Dn/DCJ−λκ relation does not depend on the type of test gas. The propagation behavior of the stabilized curved detonation waves is controlled by the Dn/DCJ−λκ relation. Due to this propagation characteristic, the fully-developed, stabilized curved detonation waves propagate through the curved channels while maintaining a specific curved shape with a constant angular velocity. Self-similarity is seen in the front shock shapes of the stabilized curved detonation waves with the same ri/λ, regardless of the curved channel and test gas

Stable detonation wave propagation in rectangular-cross-section curved channels

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Propulsive Performance and Heating Environment of Rotating Detonation Engine with Various Nozzles

Geometric throats are commonly applied to rocket combustors to 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 to simulate varied backpressure conditions in a range of 1.1–104 kPa. For the throatless case, the detonation channel area was regarded to be equivalent the throat area, and three throat-contraction ratios were tested: 1, 2.5, and 8. Results revealed that combustor pressure was approximately proportional to equivalent throat mass flux for all test cases. Specific impulse was measured for a wide range of pressure ratios, defined as the ratio of the combustor pressure to the backpressure in the vacuum chamber. The rotating detonation engine could achieve almost the same level of optimum specific impulse for each backpressure, whether or not flow was squeezed by a geometric throat. In addition, heat-flux measurements using heat-resistant material are summarized. Temporally and spatially averaged heat flux in the engine were roughly proportional to channel mass flux. Heat-resistant material wall compatibility with two injector shapes of doublet and triplet injection is also discussed

Full Kinetic Analysis of Small-scale Magneto Plasma Sail in Magnetized Solar Wind

Abstract: Magneto Plasma Sail (MPS) is spacecraft propulsion that produces an artificial magnetosphere to block solar wind particles, and thus imparts momentum to accelerate a spacecraft. In the present study, we conducted three-dimensional particle-in-cell simulations on small-scale magnetospheres to investigate thrust characteristics of MPS, in which the magnetosphere is inflated by an additional plasma injection. As a result, we revealed that finite thrust generation and the increase in thrust is obtained in the small-scale magnetosphere even if the electron kinetics is taken into consideration. The thrust of MPS (0.58 mN @ magnetic moment M=1.3 10 7 Wb·m) becomes up to 97 times larger than that of the original magnetic sail (6.0 µN). It was also revealed that the thrust gain of MPS (thrust of MPS / (thrust of magnetic sail + thrust of plasma jet)) is more than unity (up to 5.2). However, the relation of trade-off between specific impulse, thrust-mass ratio and thrustpower ratio is revealed and the optimal design of the spacecraft and missions are required for the realization of MPS. Nomenclatur

Propulsion Performance of Cylindrical Rotating Detonation Engine

This study evaluated the propulsion performance of a nozzleless, cylindrical rotating detonation engine (RDE). Using a C2H4–O2 mixture, the RDE was tested in a low-back-pressure environment at propellant mass flow rates of 8–45  g/s. In high-speed imaging of the self-luminescence within the combustor, rotating luminous regions were observed at mass flow rates above 22  g/s. Measured pressure distributions suggest that burned gas reached sonic velocity at the combustion chamber outlet. This paper proposes the structure of internal flow in the RDE and confirms that calculated pressure distribution based on the structure was close to the experimental distribution. This study also estimated the RDE’s thrust by pressure and momentum exchange and confirmed it by experimental measurement. Moreover, the theoretical thrust calculated under the assumption that exhaust is a sonic flow agreed with the load cell thrusts, suggesting that RDE combustion is perfectly completed inside the chamber. Specific impulses are 80–90% of specific impulses for ideal correct expanded flow for all mass flow rates, and its value was close to that of an annular RDE. In addition, RDE performance will increase by about 20% if the RDE is equipped with a divergent nozzle and the gas is correctly expanded to back pressure.journal articl

Synchronized Initiation of Two Cylindrical Rotating Detonation Engines

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 (40±2  g/s), equivalence ratio (1.0±0.1), 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 (10-mm 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.Published Online:20 Jun 2024journal articl

Torque Around Axial Direction on Rotating Detonation Engines

A rotating detonation engine (RDE) generates a continuous thrust with one or more rotating detonation waves. Because of the velocity on the order of kilometers/second, the reaction zone is relatively small. Therefore, the RDE realizes a short combustion chamber length. However, the detonation waves induce an azimuthal motion of propellant, resulting in torque around the thrust axis. Because the motion does not contribute to the thrust, the torque is important in terms of performance loss. Herein, we conducted combustion tests with a six-axis force sensor to simultaneously measure 0.149±0.009  Nm torque and 48.1±0.9  N thrust. A comparison of detonation waves captured by high-speed camera revealed that the torque followed the direction and was offset when the waves existed in both of two directions simultaneously, which indicates the possibility of controlling the torque. Under a mass flow rate at 87±9  g/s and an equivalence ratio at 1.43±0.28, when the azimuthal component of shear force was 8.8±0.6% of the thrust, 0.77±0.10% of the total kinetic energy of the exit flow was distributed to the azimuthal component of velocity and did not contribute to the thrust. We therefore concluded that the effect of the azimuthal motion on the RDE’s performance was small.journal articl

Experimental study of internal flow structures in cylindrical rotating detonation engines

The internal flow structures of detonation wave were experimentally analyzed in an optically accessible hollow rotating detonation combustor with multiple chamber lengths. The cylindrical RDC has a glass chamber wall, 20 mm in diameter, which allowed us to capture the combustion self-luminescence. A chamber 70 mm in length was first tested using C2H4single bondO2 and H2–O2 as propellants. Images with a strong self-luminescence region near the bottom were obtained, confirming the small extent of the region where most of the heat release occurs as found in our previous research. Based on the visualization experiments, we tested RDCs with shorter chamber walls of 40 and 20 mm. The detonation wave was also observed in the shorter chambers, and its velocity was not affected by the difference in chamber length. Thrust performance was also maintained compared to the longer chamber, and the short cylindrical RDC had the same specific impulse tendency as the cylindrical (hollow) or annular 70-mm chamber RDC. Finally, we calculated the pressure distributions of various chamber lengths, and found they were also consistent with the measured pressure at the bottom and exit. We concluded that the short-chamber cylindrical RDC with equal length and diameter maintained thrust performance similar to the longer annular RDC, further expanding the potential of compact RDCs.Available online 17 August 2020journal articl