Masatoshi Kodera

Applications of unstructured hybrid grid method to high-Reynolds number viscous flows

An unstructured hybrid grid method is discussed for its capability to compute three-dimensional compressible viscous flows of complex geometry. A hybrid of prismatic and tetrahedral grids is used to accurately resolve the wall boundary layers for high-Reynolds number viscous flows. The Navier-Stokes equations for compressible flows are solved by a finite volume, cell-vertex scheme. The LU-SGS implicit time integration method is used to reduce the computational time for very fine grids in boundary layer regions. Two kinds of one-equation turbulence models are evaluated here for their accuracy. The method is applied to computations of transonic flows around the ONERA M5 airplane and ONERA M6 wing, and supersonic shock/boundary layer interacting flows inside a scramjet inlet to validate the accuracy and efficiency of the method

Experimental Study of Strut Injectors in a Supersonic Combustor Using OH-PLIF

The present experimental study, carried out in the framework of ONERA-JAXA cooperation, deals with the ignition and hydrogen flame development in a supersonic air flow at Mach 2.5, with stagnation conditions of 0.6 MPa and 1620 K. Different types of struts for hydrogen injection (equivalence ratio of 0.45) have been investigated. Some (Alternating-Wedge struts) are specially designed to enhance fuel-air mixing by the creation of streamwise vortices. The principal aim of this study is to obtain information on the flow structures in the supersonic combusting mixing layer, and the location of ignition and combustion zones. Instantaneous maps of the OH radical by means of the OH-PLIF technique in two planes downstream of the strut injector and conventional observations of the flame by video camera have been obtained. These visualizations are completed by global measurements of the ignition delays and heat release by means of wall pressure measurements. The efficiency of the different strategies of fuel injections are then compared and discussed.

Influence of Thermal Non-equilibrium on Nozzle Flow Condition of High Enthalpy Shock Tunnel HIEST

Axisymmetric RANS simulations using thermal and chemical non-equilibrium models are applied to a nozzle flow of a free-piston driven shock tunnel HIEST. The static pressure and Pitot pressure profiles at the nozzle exit are compared with the experimental results to examine influence of thermal non-equilibrium on the nozzle flow and validate the CFD code. As the thermal non-equilibrium model, the two-temperature model of Park and a fourtemperature model, for which the conservation of the vibrational-electronic excitation energy for N2, O2 and NO is taken into account individually, are applied to the code. The results show that the HIEST nozzle flow is close to thermal equilibrium with air as the test gas and thermal non-equilibrium with nitrogen. Good agreement between the CFD and the experimental results is obtained at the total enthalpy conditions of 3.81 MJ/kg with air and 7.99 MJ/kg with nitrogen. However, some discrepancy in the Pitot pressure profiles appears at the 10.0 and 19.0 MJ/kg conditions with air. The CFD with the 4T model does not simulate the air nozzle flow at the 3.81 MJ/kg condition. The results indicate that the 4T model underestimates the vibrational relaxation rate.

Mixing and Combustion Control Strategies For Efficient Scramjet Operation in Wide Range of Flight Mach Numbers

In this paper, we present our main results of the firing tests of a newly proposed hydrogen fueled dual-mode scramjet combustor. The present scramjet combustor aims at obtaining a better engine performance, working characteristics and operability in wide range of flight Mach numbers. To realize such a scramjet, we especially focus on the following technical issues 1) use of parallel/low angle fuel injection, 2) control of combustor boundary layer separation, 3) good ignition/flameholding ability as well as efficient fuel/air mixing and combustion in the supersonic core flow by the streamwise vortices, and 4) selective operability of supersonic/subsonic combustion modes and efficient combustor operation in these modes. Involving these technical issues, our basic idea for the combustor of such efficient scramjet performance and operability is a multiple staged combustor characterized by the combination use of the “Alternating-Wedge strut” injector and wall-mounted ramp injectors both of which generate streamwise vortices. Firing test results showed a superior ability of this type of combustor to perform a supersonic combustion in the combustor core flow and operability in supersonic/subsonic combustion modes with high thrust performance in wide range of flight Mach numbers.

Aerodynamic force measurement on a large-scale model in a short duration test facility

A force measurement technique has been developed for large-scale aerodynamic models with a short test time. The technique is based on direct acceleration measurements, with miniature accelerometers mounted on a test model suspended by wires. Measuring acceleration at two different locations, the technique can eliminate oscillations from natural vibration of the model. The technique was used for drag force measurements on a 3m long supersonic combustor model in the HIEST free-piston driven shock tunnel. A time resolution of 350μs is guaranteed during measurements, whose resolution is enough for ms order test time in HIEST. To evaluate measurement reliability and accuracy, measured values were compared with results from a three-dimensional Navier–Stokes numerical simulation. The difference between measured values and numerical simulation values was less than 5%. We conclude that this measurement technique is sufficiently reliable for measuring aerodynamic force within test durations of 1ms.

Numerical Simulation of a Scramjet Engine for JAXA's Flight Experiment Using Hyshot

In this study, a numerical simulation was made on the scramjet engine examined in the JAXA’s flight experiment program using the HyShot planed for 2005 under Mach 8 flight conditions. Two different fuel injectors were applied to the engine, namely Hyper Mixer injector, which generates streamwise vortices to enhance supersonic mixing and combustion, and Back Step injector, which generates no streamwise vortices. The numerical simulation method well predicted the ground experiment data obtained by the High Enthalpy Shock Tunnel (HIEST) before the flight experiment. The computed results revealed the details of the engine characteristics and the influences of changing the incoming free stream temperature and the wall temperature, which are probably different between the flight experiment and the ground experiment.

Multi-Objective Design and Trajectory Optimization of Space Transport Systems with RBCC Propulsion via Evolutionary Algorithms and Pseudospectral Methods

In this study, a multi-objective design optimization coupling evolutionary algorithms and trajectory optimization via pseudo-spectral methods has been conducted for the first stage of two-stage to orbit (TSTO) system with a rocket-based combined cycle (RBCC) engine which combines rockets and ramjets by blending two kinds of vehicle configurations with different aerodynamic characteristics. The design criteria include the minimization of fuel consumption and the maximization of the final Mach number up to a separation of the TSTO system at the maximum altitude under certain ranges of acceleration and dynamic pressure. The optimization results reveal a counteractive trend between the final Mach number and fuel mass ratio and the major impact of effective specific impulse on those two objectives, which is mainly controlled by thrust throttling parameter within the trajectory optimization. In addition, the RBCC-powered vehicle tends to fly at lower altitude to attain the minimum fuel mass ratio, in contrast to the case for maximum final Mach number, which is attributed to the hybrid aerodynamic performance of the two configurations. The insight gained here can be usefully applied to the design of high-performance RBCCpowered vehicles.

Numerical study of mixing and combustion process of a scramjet engine model

In this study, a numerical analysis of combusting flows was made on the scramjet engine which has been tested at the Ramjet engine Test Facility (RJTF) of the Kakuda Research Center. Our main purpose is to numerically investigate the change of engine performance at Mach 8 flight condition such as mixing and combustion within the combustor and resulting engine thrust with increasing the fuel equivalence ratio as well as the transition mechanism from start to unstart condition. CFD results showed good agreement with experiments qualitatively and quantitatively. We made some considerations on the features of the fuel/air mixing and combustion process in the engine.

Scramjet inlet flow computations by hybrid grid method

Computations of internal viscous flowfields of scramjet models were conducted at inflow Mach number of 5.4. An unstructured hybrid grid method was used to compute complex geometries such as scramjet models with a short strut. The numerical method to solve the Navier-Stokes equations on the hybrid grid was developed using a finite volume cell vertex scheme and the LU-SGS implicit time integration algorithm. The computational results using one-equation turbulence models showed good agreement with the experimental data. The flow features and the changes of flowfields due to the short strut located in the upper passage were discussed. It was revealed that a thick subsonic region did not exist in the combustor near the top wall at Mach number 5.4. It was favorable features to avoid the engine unstart. With the strut, relatively low velocity regions became larger and the down wash flow toward the cowl behind the step became strong. The overconcentration of the fuel toward the top wall during the weak combustion was found in the experiment. From the computational results, the reason of this overconcentration was realized that the airflow near the injector was turned to the top wall due to the small influence by the combustion in the experiment. The computational time and the accuracy of the present method were the same level as the conventional structured grid methods. Thus the present method seemed to be engineeringly very useful for analysis and design of the high-speed propulsion engines.

Ignition Transient Phenomena in a Scramjet Engine at Mach 12 Flight Condition

Ignition transient phenomena in a hydrogen-fueled scramjet engine with hyper-mixer injectors at Mach 12 flight condition were numerically examined by using an unsteady Reynolds-Averaged Navier-Stokes Simulation (RANS) approach in terms of ignition, flame spreading and flow evolution. The calculation results showed the high potential of streamwise vortices generated by the injector for ignition due to the ingestion of a free shear layer with high temperatures. In addition, the results revealed that two notable events for flame stabilization happened during the ignition process. One is the generation of a radical pool within a boundary layer on the engine side walls. Another is the anchoring and extension of flame along the side of fuel flows. However, the flame front traveling along with the fuel flow couldn’t stay in the combustor because of high-speed main flows. It was observed that the reflection points of shock waves on the side walls moved upstream one after another during the flame establishment.