Nobuhiro Yamanishi

LE-X -Japanese Next Liquid Booster Engine-

The LE-X engine is under study for Japan’s next flagship expendable launcher (post H2A) to be operated in the next decade with enhanced reliability and reduced cost. The goal of LE-X development is to meet the requirements from the vehicle for higher reliability, lower production cost and appropriate performance. Technology development itself is also a purpose of this investigation and will be applied to other forthcoming engines to be developed in Japan. The early-stage feasibility study of the LE-X engine was completed in 2005 through primary studies on system design, engine component design, cost reduction, reliability prediction, subscale testing, and computational simulation. In 2006, engine system analysis and fundamental studies on LE-X components by means of element tests were successfully conducted. In 2007, we have optimized the engine baseline configuration from aspect of cost reduction activities. Significant cost reduction will be achieved by drastic simplification of the engine system, and the innovation of the manufacturing process. Technology development will be ongoingly conducted to mitigate development risks, such as precise life prediction analysis of combustion chamber, prediction of combustion instability, and high-fidelity simulation.

Combustion and Heat Transfer Modeling in Regeneratively Cooled Thrust Chambers (Co-axial Injector Flow Analysis)

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.

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

In recent years, methane has attracted attention as a propellant for liquid rocket engines because of its various advantages compared to typical propellants such as hydrogen. When methane is used as a coolant for a regenerative cooling system, its near-critical thermodynamic and transport properties experience large variations because its critical pressure is higher than that of typical propellants; this significantly influences the flowfield and heat transfer characteristics. Therefore, adequate understanding of the flowfield and heat transfer characteristics of methane in regenerative cooling channels is a prerequisite for future engine development. In this study, conjugated coolant and heat transfer simulations were performed to investigate the flowfield and heat transfer characteristics of transcritical methane flows in a sub-scale methane-cooled thrust chamber. The computed results were validated against experimental data measured in hot firing tests. They compared well with the measured pressures and temperatures in cooling channels, and wall temperatures were within the permitted levels. Detailed flow analysis revealed peculiar flow structures in the cooling channel: a strong secondary flow induced in the concave-heated part in the channel throat section and the coexistence of two different gas phases—ideal and real—in a single cross-section in the cylindrical region. A high wall temperature appeared in the cylindrical region of the thrust chamber under the considered conditions; this was due to the heat transfer deterioration induced by an M-shaped velocity profile and a turbulent heat flux reduction.

Next Booster Engine LE-X in Japan

The LE-X is a new cryogenic booster engine with high performance, high reliability and low cost, being designed for the next-generation Japanese launch vehicle. Following an early- stage feasibility study, detailed studies focusing on front-loading design approaches were conducted in 2006 and 2007. An optimum engine system was determined and potential risks were mitigated by various element tests. These activities will lead to the realization of the first booster engine in the world with expander cycle.

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Subscale Thrust Chamber

Flowfield and heat transfer characteristics of supercritical parahydrogen flows in a cooling channl of a sub-scale hydrogen-cooled thrust chamber are investigated using Reynolds-Averaged Navier-Stokes simulation, in which conjugated heat transfer between coolant flow and chamber wall is taken into account directly. The considered system pressure ranges from 3.8 to 4.6 MPa, and temperature from 43 to 324 K at Reynolds number more than 1× 10 5 . The computed results are validated against the experimental data measured in the hot firing testings, which compare well with measured pressures and temperatures in a cooling channel, and wall temperature in a hot firing testings. Detailed flow structure and heat transfer characteristics in a cooling channel are clarified, showing strong thermal and density stratification, secondary flow effect, and particular asymmetric heat transfer characteristics.

Numerical Investigation of Supercritical Coolant Flow in Liquid Rocket Engine

Understanding and predicting the flowfield and heat transfer characteristics in cooling channels are prerequisite to improve design and performance of regeneratively cooled rocket thrust chambers. In order to realize them, a CFD code, able to predict such characteristics, is developed based on a pressure-based solver and the cubic-type equation of state to take into account the real gas effect. As a preliminary study, simulations of transcritical parahydrogen flows in a uniformely heated circular tube are performed in order to validate the developed code against the reference experiment and investigate the flowfield and the heat transfer characteristics under transcritical conditions. The computed results agree well with the experimental data with regard to the wall temperature, the heat transfer coefficient, the bulk pressure and temperature. Also, the peculiar behavior, called “heat transfer deterioration”, under transcritical condition with high heat flux, is successfully predicted. The simulated flowfield reveal the mechanism of it. The parametric studies with different heat flux levels clarify the condition in which the heat transfer deterioration takes places.

Combustion and Heat Transfer Modeling in Regeneratively Cooled Thrust Chambers (Optimal Solution Procedures for Heat Flux Estimation of a Full-Scale Thrust Chamber)

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.

Conjugated Combustion and Heat Transfer Modeling for Full-Scale Regeneratively Cooled Thrust Chambers

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 and Heat Transfer Modeling in Regeneratively Cooled Thrust Chambers (Wall Heat Flux Validation)

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.

Multi Objective Optimization of a Supersonic Axial Turbine Blade Row Shape for a Rocket Engine Turbopump

A rocket engine should be small and low weight, therefore, a turbopump for a rocket engine must be smaller and have higher rotation speed than a conventional pump. However, to achieve high thrust, the required power of the pump should be high enough to achieve high specific impulse and thrust. To attain these requirements, a study of design optimization with generic algorithm was applied to blade row shape of the supersonic axial turbine. In this study, a multi-objective optimization was carried out to obtain a tradeoff tendency between multi-objective functions, turbine performance and turbine structural strength. In the present optimization, unsteady CFD was carried out in each optimization population to estimate turbine efficiency more clearly since shock interaction between stator and rotor is one of the most important points for supersonic turbine performance estimation. The optimized result showed that there is a strong tradeoff between turbine efficiency and diameter. The tradeoff information can be used to improve turbopump performance to satisfy requirements as a component of a rocket engine.