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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].

As a propellant for hybrid rocket engines using liquid oxidizer and solid fuel and for liquid rocket engines, the use of self-pressurized fluids such as nitrous oxide has become widespread. Since these fluids can be self-pressurized by their high saturated vapor pressure, the propulsion system becomes smaller and simpler. However, this self-pressurization generally forms a gas–liquid two-phase flow by flashing or cavitation. This flow is considered highly unsteady because the temperature and pressure greatly change with the discharge process. In this study, unsteady flow characteristics due to self-pressurization were experimentally obtained by conducting many cold flow tests with carbon dioxide as self-pressurizing fluids. As a result, it was clarified that the fluid temperature dropped about 10–15 K with the pressure drop due to feedline pressure loss during the discharge process. From these experimental results, we estimated the bubble growth and void fraction change that would satisfy the temperature drop. In this paper, the obtained test results and estimated temperature drop are reported.

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., Published online: 12 November 2018

Nitrous oxide (N2O) is a suitable propellant for small rocket engines (mostly kN class), and has been widely used in various countries given its high saturated vapor pressure (i.e., 6 MPa at 300 K), which enables self-pressurization. Because nitrous oxide exists with gas-liquid equilibrium in tanks, cavitation occurs when the pressure in the tanks and feed lines drops slightly, which easily forms a gas-liquid two-phase flow. Since accurately estimating the performance of rocket engines requires ascertaining the characteristics of their propellant flows, flow tests with self-pressurization using nitrous oxide were conducted, as were firing tests of hybrid rocket engines using nitrous oxide as a liquid oxidizer and acrylic as a solid fuel. This paper presets the results of those tests, along with findings obtained regarding the characteristics of the gas-liquid two-phase injector flow with self-pressurization. In the N2O flow test, the injector upstream pressure was approximately 1.5 MPa, while the injector downstream pressure was approximately 0.1 MPa. At 0.070, the ratio of upstream to downstream pressure of the injector was thus extremely large, which suggested that the gas-liquid two-phase flow was choked with the injector. By contrast, in the firing test with a chamber pressure of approximately 1.0 MPa and a thrust of approximately 650 N, the ratio of the injector downstream pressure (i.e., chamber pressure) and its upstream pressure was approximately 0.56. Although that ratio was relatively large, because the injector upstream pressure is relatively low (i.e., approximately 1.8 MPa) and the backpressure fluctuated due to combustion, it remains unclear whether the gas-liquid two-phase flow was choked.

A detonation is a combustion wave that propagates at a supersonic speed (2-3 km/s) in a combustible mixture. There are many fundamental studies of detonation waves and detonation engine systems. The detonation cycle has a higher thermal efficiency than a conventional constant-pressure combustion cycle.

形態: 図版あり, Physical characteristics: Original contains illustrations, 資料番号: PA1810070000

形態: 図版あり, Physical characteristics: Original contains illustrations, Accepted: 2017-09-25, 資料番号: PA1710052000

ASME 2016 International Mechanical Engineering Congress and Exposition (November 11-17, 2016. Phoenix Convention Center), Phoenix, Arizona, USA, 資料番号: PA1710032000

Currently in development at Muroran Institute of Technology, a small gas-generator-cycle air turbo ramjet engine (GG-ATR) for a supersonic unmanned aerial vehicle requiring deep throttling uses liquid oxygen/ethanol as a propellant. However, it is difficult to maintain a liquid-phase flow at a low mass flow rate or at a low pressure due to flashing in the feedline. In response, we investigated the ability of our ground test facility to feed liquid oxygen at a low mass flow rate. The experimental results of the flow tests validated the proposed one-dimensional model, including heat input and the density change in the feed line. The minimum possible flow rate as a subcooled liquid flow was examined with the model and the throttling level of 75 % was possible when assuming 110 K in tank temperature.

A detonation is a combustion wave that propagates at supersonic speed (2∼3 km/s) in a combustible mixture. There are many fundamental studies of detonation waves and detonation engine systems. The detonation cycle has a higher thermal efficiency than a conventional constant-pressure combustion cycle. Therefore, it is expected that a high-efficiency propulsion system can be realized using detonation waves.A rotating detonation engine (RDE) uses continuous detonation propagating at a bottom in an annular combustor. As detonation waves propagate at a supersonic speed only in the bottom region of the RDEs, the combustor can be shortened. However, the combustor needs cooling system due to high heat flux to the combustor wall. In this experimental study, we performed combustion tests of RDE system using gaseous ethylene and oxygen as the propellant. This RDE system performance will also be demonstrated in space environment by the sounding rocket. We measured the combustor pressure, temperatures, heat flus, mass flow rate and thrust. The RDE system used in this study is shown in Figure 1. We performed the long-duration rotating detonation engine combustion tests for at sea level condition. The stable trust histories were obtained.

Fairness measures are means to guess if a curve is suited for an aerodynamic application if no knowledge of the flow field is available. Fair curves are proposed as a means to reduce the number of parameters for generating a supersonic rotor blade. A specific new measure of fairness is proposed with the goal of being numerically less demanding to evaluate than previously known measures of fairness. Gradient search for an optimum of this fairness value is nevertheless computationally expensive. Therefore, a neural network was trained using Adamax optimization that can produce curves of acceptable quality instantly for applications where real time response is needed.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2016-02-11, 資料番号: PA1710030000

Rotor-dynamic fluid force (RD fluid force) of turbomachinery is one of the causes of the shaft vibration problem. Bulk flow theory is the method for analyzing this RD fluid force, and it has been widely used in the design stage of machine. The conventional bulk flow theory has been carried out under the assumption of concentric circular shaft's orbit with a small amplitude. However, actual rotating machinery's operating condition often does not hold this assumption, for example, existence of static load on the machinery causes static eccentricity. In particular, when such a static eccentricity is significant, the nonlinearity of RD fluid force may increase and become non-negligible. Therefore, conventional bulk flow theory is not applicable for the analysis of the RD fluid force in such a situation. In this paper, the RD fluid force of the annular plain seal in the case of circular whirling orbit with static eccentricity is investigated. The case with both the significant static eccentricity and the moderate whirling amplitude is considered, and the perturbation analysis of the bulk-flow theory is extended to investigate the RD fluid force in such cases. In this analysis, the assumption of the perturbation solution is extended to both static terms and whirling terms up to the third order. Then, the additional terms are caused by the coupling of these terms through nonlinearity, and these three kinds of terms are considered in the extended perturbation analysis of the bulk flow theory. As a result, a set of nonlinear analytical equations of the extended perturbation analysis of the bulk flow theory, for the case with both the significant static eccentricity and the moderate whirling amplitude, is deduced. The RD fluid force for such cases is analyzed, and the occurrence of constant component, backward synchronous component, and super-harmonic components in the RD fluid force is observed in addition to the forward synchronous component. The representation of RD fluid force coefficients (RD coefficients) are modified for the case with significant static eccentricity, and the variation of RD fluid force coefficients for the magnitude of static eccentricity is analyzed. These analytical results of RD fluid force and its RD coefficients are compared with the numerical results using finite difference analysis and experimental results. As a result, the validity of the extended perturbation analysis of the bulk-flow theory for the case with both the significant static eccentricity and the moderate whirling amplitude is confirmed.

In recent years, along with demands for higher rotational speed and higher efficiency in the rotating machinery, shaft vibration has been a serious problem. One of the causes of this shaft vibration problem is the rotor-dynamic fluid force (RD fluid force) generated by working fluid at turbo machinery parts. It is important in the design stage of rotating machines to estimate the RD fluid force and predict the stability of the rotor system accurately. Therefore, many researches have been conducted to clarify the characteristics of RD fluid force. One of the traditional methods for analyzing the RD fluid force is the bulk flow theory. However, in conventional bulk flow analysis, it is assumed that the amplitude of the shaft displacement is sufficiently smaller than the clearance. Therefore, influence of nonlinearity in the large amplitude whirl may not be included in this analysis. Accordingly, this paper focuses on constructing a coupled analysis of the fluid force and the shaft vibration that describes each behavior of fluid and shaft at the same time. By using this coupled analysis, the interaction between fluid and shaft systems can be taken into consideration more accurately. Regarding the fluid region, finite difference method is used for bulk flow continuity and momentum transport equations. Incompressible fluid is assumed, and the pressure field is calculated by solving the Poisson equation of pressure. In solving the Poisson equation of pressure, a specific problem for this coupled analysis relating unknown shaft acceleration arises. In this paper, this problem is solved by obtaining the approximated acceleration based on Newmark-beta technique. This coupled analysis is conducted for a simple flexible rotor system with annular plain seal, and the frequency response is obtained. First, the case with isotropic support stiffness and with no gravitational force is considered. Then, the case with the constant load and the case with anisotropic support stiffness are analyzed. These analytical results show that both the constant load and structural anisotropy may affect the stability of the rotor system. As a result, the usefulness of the proposed coupled analysis procedure of the fluid force and the shaft vibration is validated.

The bulk-flow theory for the rotordynamic (RD) fluid force has been investigated for many years. These conventional bulk-flow analyses were performed under the assumption and restriction that the whirl amplitude was very small compared to the seal clearance while actual turbomachinery often causes the large amplitude vibration, and these conventional analyses may not estimate its RD fluid force accurately. In this paper, the perturbation analysis of the bulk-flow theory is extended to investigate the RD fluid force in the case of concentric circular whirl with relatively large amplitude. A set of perturbation solutions through third-order perturbations are derived explicitly. It relaxes the restriction of conventional bulk flow analysis, and it enables to investigate the RD fluid force for the whirl amplitude up to about a half of the clearance. Using derived equations, the nonlinear analytical solutions of the flow rates and pressure are deduced, and the characteristics of the RD fluid force are investigated in both radial and tangential directions. The influence of the whirl amplitude on the RD fluid force is explained and validated by comparing with computational fluid dynamics (CFD) analysis. These results are useful for the analysis and prediction of frequency response of the vibration of the rotating shaft system considering the RD fluid forces.

A method to analyze pump dynamic characteristics by cavitation compliance, mass-flow gain factor and dynamic gain was evolved to analyze rotating cavitation, using one-dimensional multi-domain system analysis software. This analysis method was used to analyze one-dimensional phenomena such as surge phenomenon in the past. For example, POGO, which is an auto-oscillation where the thrust oscillation of rocket propulsion system including turbopumps is coupled with the natural frequency of the rocket structure, has been analyzed using this method. In the present study, cavitation compliance and mass-flow gain factor were divided and distributed in each flow path of an inducer and tried to evaluate the possibility of analyzing rotating phenomena such as rotating cavitation by this one-dimensional analysis method. Whirl motion of the rotor assembly was accounted for by variable flow area orifices which connect the next flow path. The areas of the orifices are changed harmonically and the oscillation phase of each orifice is given according to the number of the blades. The variation of leakage flow rate through the tip clearances of each blade is simulated by the motion of the variable flow area orifices. With the evolved analysis model, effects of various parameters on the eigenvalues of the system were investigated. Analysis results showed that there existed rotating phenomena and agreed with inducer test results qualitatively., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: PA1610062000

In this study, numerical analysis is carried out around the cyclic flat-plate cascade with symmetric and asymmetric slit, so as to examine the suppressing or controlling effect of the slit on cavitation instabilities such as cavitation oscillation (CO) which resembles cavitation surge, and rotating cavitation. These instabilities cause various problems for the turbomachinery, for example, rotating cavitation causes an asynchronous shaft vibration, and CO causes an oscillation of column of working fluid as a result of the resonance phenomenon of the system. In liquid propellant rocket engine, suppression device for these instabilities bring increase in cost of the launch. Therefore, it is thought that to develop effective suppression technique is important for turbopumps. Especially, in this paper, two types of the flat-plate three blades cascade which have symmetric slit on each blade and three types of the cascade which have asymmetric slit were analyzed, and the results are compared with those of cascade without slit. As a result, the CO is perfectly suppressed in both of two types cascade with asymmetric slit. Also, other examined cascades have suppression effect of CO These results indicate the possibility of suppressing cavitation instabilities in actual inducers or controlling the type of the cavitation instabilities by the arrangement of the slit. Moreover, the head performance is equal or slightly increased by arranging slit.