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形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-11-21, 資料番号: PA2110018000

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: 2017-11-03, 資料番号: PA1810069000

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2017-07-06, 資料番号: PA1810059000

Hydrazine (N2H4)/nitrogen tetroxide (NTO) co-flowing plane jets were simulated to explore the hypergolic ignition processes and flame structures in N2H4/NTO bipropellant thrusters. The Navier–Stokes equations with a detailed chemical kinetics mechanism were solved in a manner of direct numerical simulation to reveal the influence of the distinct chemical reaction, i.e., hydrogen abstraction by nitrogen dioxide (NO2) and the thermal decomposition of N2H4. In the ignition processes, the hydrogen abstraction sequence played a significant role in preheating the mixture gases. Further, the ignition eventually occurred in the region where both N2H4 and NTO were sufficiently supplied for preheating. Hence, the ignition position and delay strongly depended on the fluid-mixing conditions. After the flames reached a steady state, the combustion flames uniquely comprised two types of flames, the outer diffusion flame and the inner decomposition flame. The outer diffusion flame came from the oxidization by NTO. The inner decomposition flame was caused and maintained by the heat transfer from the outer diffusion flame and a high rate of heat release from the thermal decomposition of N2H4. Because of the decomposition flame, the decomposition products such as NH3 and H2 were the major constituents of the downstream combustion gases., 資料番号: PA1510062000

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.

5th European Conference for Aeronautics and Space Sciences, EUCASS 2013 (July 1-5, 2013.), Munich, Germany, The characteristics of film cooling in a CH4/O2 sub-scale chamber with multi-injector elements and two kinds of film cooling slot dimensions are investigated using a calorimeter chamber in experiments and simulations, in which the finite rate chemistry with a reduced CH4/O2 reaction mechanism is taken into account. The computed wall heat flux and pressure distributions are compared to the experimental results, which overall show good agreement. A large slot dimension is shown to induce mixing with a core flow. This mixing causes a low heat flux distribution near a face plate along with high combustion efficiency., 資料番号: PA1510092000

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.

For LOX/LH2 shear coaxial injectors, it is well-known that high-frequency combustion instabilities may occur when the injection temperature of hydrogen decreases below a certain value, but the mechanism of the initiation of combustion instability with a coaxial injector is still not clear. In the present study, firing tests were conducted with five types of single shear coaxial injector elements by using LOX and LH2 as propellants to further investigate the mechanism of the initiation of combustion instability during temperature-ramping changes during hydrogen injection. Results showed that unstable combustion was initiated when the hydrogen injection temperature decreased to less than a certain cryogenic temperature. The combustion instabilities observed in the present firing tests are discussed and classified into three different types.

The physical phenomena in the liquid rocket combustion chamber are very complicated such as turbulence, reaction, and real-gas effect. The prediction of heat flux on liquid rocket chamber wall is a challenging problem due to these complex physics This work is aimed particularly at the heat flux validation for the turbulent boundary layer in three physical situations; turbulent boundary layer on heated flat plate, turbulent boundary layer with pressure gradient, and recirculation zone of heated expansion tube. Each physical situation corresponds to the straight part of the combustion chamber, the rocket nozzle, and the exit of injectors. For the Reynolds-Averaged Navier-Stokes simulations, an adequate turbulent model should be selected to suit the flow feature. In this paper, three turbulent models have been tested for the above three validation problems. There is no perfect turbulent model for the prediction among the three turbulent models for all validation cases. The two-layer turbulent model works reasonably well for GH2/GO2 single injector conducted at Penn State University among the three turbulent models. Furthermore, turbulent models affect on not only the heat flux but also the turbulent intensity in the contraction tube of the GO2 injector and the flame shape.

The performance of liquid rocket of expander bleed cycle engine system strongly depends on a regenerative cooling performance to provide the required heat in the Main Combustion Chamber (MCC). The prediction of the MCC wall heat transfer characteristics and wall temperature distributions is very important for designing a new engine. Developing a tool for a combined analysis among the combustion gas region, chamber structure, and the cooling channel is our ultimate goal. As a part of the development, we carried out numerical simulations of combustion flow fields in an injector, in order to understand the flow characteristics, which may influence the heat transfer and temperature on the MCC. This report shows the numerical simulation results of GH2/GO2 coaxial flow. This problem is adequate for validation of mixing, diffusion, and chemical reaction. Numerical simulations are carried out using the commercial code FLUENT, Advance/FrontFlow/red, and CRUNCH CFD. Computed velocity and mole fraction are compared with the firing tests data conducted at the Pennsylvania State University, using Eddy Dissipation model, Laminar Finite Rate model, and Eddy Dissipation Concept (EDC) model. The results of steady state simulations have a tendency to estimate a longer flame than the test data, since the vortex mixing around the GO2 post was underestimated and moreover the flow field is basically unsteady. The time-averaged values of unsteady simulations of FLUENT with EDC model simulate turned out to be good agreement with the test data relative to the steady state solution, indicating the importance of capturing the mixing process accurately. Influence of several important numerical factors such as grid resolution, combustion model, and turbulence model on the unsteady combustion flow field will be discussed in feature.

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.

Combustion flowfields in GH2/LOX sub-scale calorimeter chambers with multi-injector elements and full-scale thrust chamber are investigated using Reynolds-Averaged NavierStokes simulation, in which the finite rate chemistry with the H2/O2 detailed reaction mechanism is taken into account. The computed wall heat flux distributions are compared to that of the simplified cases to reduce a computational cost. The considered simplifications are a presence of reaction and a number of injector rows. At first, these simplifications are validated in the simulation of sub-scale chambers. The reaction is essential for the prediction of heat flux because it makes change the species distribution in a thermal boundary layer on a thrust chamber wall. A heat flux using a combustion simulation with only outermost injectors shows a good agreement with that with an original configuration near a face plate. On the other hand, it overestimates the heat flux around nozzle and throat parts. It is clarified that this overestimate comes from the shortage of unburned hydrogen near a chamber wall in the simplified method. Next, the simplification of the number of injector rows are applied to the simulation of full-scale thrust chambers. The effectiveness of this simplification for the prediction of wall heat flux is revealed. The optimal solution by using of the simplification is proven to be effective for the prediction of heat flux in a full-scale thrust chamber.