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形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2021-12-24, PA2210037000

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination,...

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2021-07-18, 資料番号: PA2210010000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-06-24, 資料番号: PA2110026000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-02-23, 資料番号: PA2010025000

Accepted: 2019-12-13, 資料番号: SA1190199000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-04-20, 資料番号: PA2110006000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2019-02-08, 資料番号: PA1910045000

Accepted: 2018-07-23, 資料番号: SA1180173000

51st AIAA/SAE/ASEE Joint Propulsion Conference (July 27-29, 2015.), Orlando, Florida, USA, 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2017-02-27, 資料番号: PA1710043000

In the present research, we experimentally verified the partial-fill effect in a multicycle pulse detonation rocket engine. The intermittent thrust of a pulse detonation rocket engine was measured by using a spring-damper mechanism that smoothed this intermittent thrust in the time direction. The intermittent mass flow rates were assessed by gas cylinder pressure or mass difference measurement. The maximum specific impulse was 305 ± 9 s at an ethylene and oxygen propellant fill fraction of 0.130 ± 0.004. When the fill fraction was greater than 0.130, the specific impulse was increased as the partial-fill fraction was decreased. When the fill fraction was less than 0.130, the specific impulse was sharply decreased as the partial-fill fraction was decreased. This decrease was due to diffusion between propellant and purge gases and the short length of the transition from deflagration to detonation. The multicycle pulse detonation rocket engine had a partial-fill effect that may have been mainly due to the suctioned air and was consistent with the single-cycle partial-fill model of Endo et al. [Endo, T., Yatsufusa, T., Taki, S., Matsuo, A., Inaba, K., and Kasahara, J., "Homogeneous-Dilution Model of Partially-Fueled Simplified Pulse Detonation Engines," Journal of Propulsion and Power, Vol. 235, 2007, pp. 1033-1041.].

ISBN: 978-1-56700-260-7, International Journal of Energetic Materials and Chemical Propulsion (ISSN: 2150-766X) Vol.8 (6), pp.531-539, 資料番号: SA1001011000

The vaporization and burning of the N2H4 and NTO droplets were simulated with the interface-tracking method to accurately explore the auto-ignition processes and the flame structures. The N2H4 vapor plume developed behind the N2H4 droplet and reacted with the ambient NO2 gas through the hydrogen abstraction reactions. Thus, the N2H4 vapor and NO2 gas mixtures behind the droplet were preheated and reached the auto-ignition at a few ms. The auto-ignition occurred in the multiple points almost at the same time. After the ignition, the premixed flame developed around the droplet. Thus, the vaporization of the liquid N2H4 near the surface became significant. Then, the double flame structures which comprise the inner decomposition flame and oxidation flame appeared around the droplet. The NTO droplet was not auto-ignited in the computational time of the present study because little N2O4 vapor near a saturated temperature decomposed to NO2 gas which is necessary for the hydrogen abstraction reactions. When the ignition was forced, the double flames developed. The outer decomposition flame propagated to the boundaries of the computational domain, while the inner oxidation flame appeared near the droplet. Except for the propagation of the decomposition flame, the NTO droplet combustion was similar to that of the industrial fuels.

In Japan, a feasibility study of the new “LE-X” booster engine has been underway since 2005. One of the key technologies of the LE-X engine development is a regenerative cooling design to produce enough power to drive the turbopumps. The LE-X engine employs an elongated chamber design to pick up enough heat energy in the regenerative cooling channels. In the current study, a fully conjugated combustion and heat transfer simulation was performed to investigate the flow field and heat transfer characteristics for the regeneratively cooled combustion chamber of the LE-X engine. A three-dimensional Reynolds-averaged Navier–Stokes simulation was used to consider the injection and combustion processes of propellants on the hot-gas side, heat conduction in the chamber wall, and cooling channel flows. Details on the three-dimensional flow and thermal characteristics of the combustion chamber were clarified in the hot-gas and coolant side domains. In particular, significant thermal stratification was found to form in the radial direction of the cooling channel in the elongated cylindrical part of the chamber, which degraded the heat transfer and resulted in the highest wall temperature there.

Acoustic losses inside a simulated liquid rocket chamber are investigated by numerical simulation coupled with theoretical calculation. In this study, the losses by injector, resonator and chamber are considered and compared. The oscillation amplitude is assumed to be small (within linear range). For injector and chamber, the loss mechanisms, such as the radiation & convection from the inlet or outlet, and viscous & thermal loss at the wall are considered. For a resonator, the viscous & thermal loss would be the major loss factor. A simulated liquid rocket chamber configuration with an injector installed off-center is investigated especially for tangential oscillation modes. It is found that with well-tuned resonator the resonant frequencies and modes would change from those without the resonator. Therefore, the coupled simulation is indispensable for resonator design. Also an order estimation of each acoustic loss factors is conducted. It is found that the viscous & thermal loss of the chamber and resonator dominate the total acoustic loss in the present configuration. However, if several hundreds of injectors is equipped as in actual rocket chamber, the loss related to injectors would become comparable to that related to resonator on the total loss.

To build a framework of a prediction tool for injection-coupled combustion instability with coaxial-type injectors, Hutt and Rocker’s methodology including Crocco’s n-τ model was applied. This linear stability analysis considers a coupling among LOX/fuel flow-path acoustics, chamber responses, and Crocco’s combustion characteristics. To validate the tool, LOX/methane subscale firing tests, which were performed and reported by NASA, were analyzed. The injection-coupled oscillating combustion, which was occurred at 5 kHz, was selected as unstable case. A stable combustion case was also selected for comparison purposes. Injection and combustion characteristics, which amplify the oscillation, were found to be dominated by the LOX-post acoustic characteristics. Chamber responses, which decay the oscillation, were estimated with two approaches: (a) onedimensional acoustic analysis with Natanzon’s methodology with a short-nozzle approximation, and (b) threedimensional acoustic analysis with a commercial software, ACTRAN. The stability was evaluated with AFC (amplitude-frequency characteristics) diagram, in which the injection and combustion characteristics and chamber responses are compared with regard to amplitudes. As a result, significant differences were not seen in the AFC diagrams between the unstable and stable cases (both analyses showed “unstable”). Further investigations for the chamber responses are needed to evaluate the potential instabilities of the system, correctly. In addition, we need to introduce a phase relationship into the tool to understand the underlying physical phenomena.

§A numerical framework is established to predict the erosion amount of nozzle wall in solid rocket motor. Understanding of the flow field and erosion mechanism is very important for the performance of solid rocket motor and the success of the launch. A coupled analysis of fluid dynamics and surface recession simulates total ablation amount. The analysis consists of two-dimensional axisymmetric fluid simulation and estimation of ablation amount using a one-dimensional heat conduction equation with thermal decomposition. The model includes the effect of the mechanical erosion. Three kinds of simulations are carried out to reveal the features of ablation and evaluate the erosion model. Features of nozzle shape can explain heat flux distributions and surface recession amount for each nozzle. The simulation results of total surface recession amount qualitatively recreate the experimental results. The numerical simulations estimate the erosion rate on the safe-side for all nozzles.

Regenerative cooling is still one of key technologies to develop high performance liquid rocket engines. To achieve high efficiency and reliability, understanding and accurate prediction of flowfield and heat transfer characteristics in regeneratively cooled thrust chambers are prerequisite. In the current study, a fully conjugated combustion and heat transfer simulation for full-scale regeneratively cooled thrust chambers was proposed and demonstrated for the LE-5B thrust chamber. In the proposed strategy, the injection and combustion processes in the hot-gas side, heat conduction in the chamber wall, and cooling channel flows are taken into account based on three-dimensional Reynolds-Averaged Navier-Stokes simulation. The computed results were validated against measured data from a hot firing test, showing reasonable agreement except for chamber outer wall temperatures. Detailed three-dimensional flow and thermal characteristics in the thrust chamber were clarified in the hot-gas side and the coolant side domains. Although the proposed numerical approach needs to be further improved quantitatively, it was confirmed that the present methodology is promising to understand and precisely predict flowfield and heat transfer characteristics in regeneratively cooled thrust chambers.

A detailed chemical kinetic mechanism for hypergolic ignition of N2H4/NO2 gas mixture at low temperatures has been constructed. In this mechanism, the hypergolic ignition is caused by following sequential reactions of H atom abstraction from N2Hm by NO2. N2H4 + NO2 = N2H3 + HONO or HNO2 (R1) N2H3 + NO2 = N2H2 + HONO (R2) N2H2 + NO2 = N2H + HONO (R3) N2H + NO2 = N2 + HONO (R4) These reactions are exothermic, especially heat release by the reaction (R4) is large because of N2 production. Temperature rise caused by the heat release accelerates the initiation reaction (R1). This 'thermal feedback' is responsible to the hypergolic ignition at ambient temperatures. Since no experimental and theoretical information was available on these reactions, rate coefficients were evaluated on the basis of transition state theory, unimolecular rate theory, and master equation analysis with quantum chemical calculations of potential energy curves. Results of simulations by using the present mechanism including reactions (R1)-(R4) reasonably agree with existing experimental data.