• Journal of computational fluids engineering
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• v.10 no.4 s.31
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• pp.51-58
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• 2005

The unsteady supersonic flow over two- and three-dimensional cavities has been analyzed by the integration of unsteady Reynolds-Averaged Navier-Stokes(RANS) with the k-$\omega$ turbulence model. The unsteady flow is characterized by the periodicity due to the mutual relation between the shear layer and the internal flow in the cavity. An explicit 4th order Runge-Kutta scheme and an upwind TVD scheme based on the flux vector split with the van Leer limiters are used for time and space discritizations, respectively. The cavity has a L/D ratio of 3 for two-dimensional case, and same L/D and W/D ratio of I for three-dimensional case. The Mach and Reynolds numbers are 1.5 and 450000 respectively. In the three-dimensional flow, the field is observed to oscillate in the 'shear layer mode' with a feedback mechanism that follows Rossiter's formula. In the two-dimensional simulation, the self-sustained oscillating flow has more violent fluctuation inside the cavity. The primary fluctuating frequencies of two- and three- dimensional flow agree very well with the 2nd mode of Rossiter's frequency. In the three-dimensional flow, the 1st mode of frequency could be seen.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1995.11a
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• pp.69-83
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• 1995

A great number of experimental data indicating shock-induced separation(SIS) in internal or external supersonic flows were reviewed to make clear the mechanism of SIS and to present the criterion of turbulent boundary layer separation. The interesting conclusions were obtained for the considerably wide range of flow geometries that the incipient separation is almost independent of the flow geometries, and that it is relatively unaffected by changes in gas specific heat, and boundary layer Reynolds number, Furthermore, the pressure rise necessary to separate boundary layer in external flows was found to be applicable to SIS in overexpanded propulsion nozzles. This is due to the fact that the SIS phenomenon caused by the interaction between shock waves and turbulent boundary layers is processed through a supersonic deceleration. This is, the SIS in almost all of interacting flow fields is governed by the concept of free interaction, and criterion of SIS is only a Function of upstream Mach number.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.10
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• pp.61-70
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• 2006

A Supersonic air intake model was designed for the high performance ramjet and dual-mode scramjet engine to operate at Mach 4 flight condition. The air intake was tested in the blowdown-type wind tunnel of Kyushu University to identify the internal flow characteristics corresponding to the flight parameters such as the back pressure, angle of attack and angle of yaw. Flow visualization was achieved by the Schlieren and oil flow visualization techniques. The intake performance was analyzed quantitatively based on the surface pressure and total Pressure measurements. The experimental results were compared with the computational fluid dynamics results. The present study exhibits the fundamental but rarely found experimental results of the high Mach number supersonic air intake.

• 한국전산유체공학회:학술대회논문집
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• 2005.10a
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• pp.235-238
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• 2005

The unsteady supersonic flow over two- and three-Dimensional cavities has been analyzed by the integration of unsteady Reynolds-Averaged Navier-Stokes(RANS) with the k - w turbulence model. The unsteady flow is characterized by the periodicity due to the mutual relation between the shear layer and the internal flow in cavities. Numerical method is upwind TVD scheme based on the flux vector split with the Van Leer limiters, and time accuracy is used explicit 4th stage Runge-Kutta scheme. Cavity flows are Comparison of two- and three-dimensional. The cavity has a L/D ratio of 3 for two-dimensional case. and same L/D and W/D ratio is 1 for three-dimensional case. The Mach and Reynolds numbers are held constant at 1.5 and 450000 respectively. For the three-dimensional case, the flow field is observed to oscillate in the 'shear layer mode' with a feedback mechanism that follow Rossiter's formula. On the other hand, the self-sustained oscillating flow transitions to a 'wake mode' for the two-dimensional simulation, with more violent fluctuations inside the cavity.

• Proceedings of the KSME Conference
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• 2005.11a
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• pp.84.2-84.2
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• 2005

• Transactions of the Korean Society of Mechanical Engineers
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• v.18 no.4
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• pp.933-945
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• 1994

A shock-wave in a supersonic flow can be theoretically determined by a given pressure ratio at upstream and downstream flowfields, and then the obtained shock-wave is stable in its position. Under the practical situation in which the shock-wave interacts with the boundary layer along a solid wall, it cannot, however, be stable even for the given pressure ratio being independent of time and oscillates around a time-mean position. In the present study, oscillations of a weak normal shock-wave in a supersonic diffuser were measured by a Line Image Sensor(LIS), and they were compared with the data of the wall pressure fluctuations at the foot of the shock-wave interacting with the wall boundary layer. LIS was incorporated into a conventional schlieren optical system and its signal, instantaneous displacement of the interacting shock-wave, was analyzed by a statistical method. The results show that the displacement of an oscillating shock-wave increase with the upstream Mach number and the dominant frequency components of the oscillating shock-wave are below 200 Hz. Measurements indicated that shock-wave oscillations may not entirely be caused by the boundary layer separation. The statistical properties of oscillations appeared, however, to be significantly affected by shock-induced separation of turbulent boundary layer.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.3
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• pp.1-13
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• 2021

Effects of the nozzle wall angles on the supersonic flow field in a thrust optimized nozzle were numerically investigated. The combustor and operating condition of 30-tonf rocket engine was selected to study the optimum nozzle shape. The nozzle flow of combustion products was realized by the shifting equilibrium calculation for the propellant of kerosene-LOx. The change of nozzle wall angles induced different developing patterns of the internal and secondary shock wave. The optimum nozzle was obtained when the internal shock was in a specific position at the nozzle outlet. The nozzle wall angles of the optimum nozzle were very similar to those of the optimum nozzle which does not consider the shock wave.

• Journal of computational fluids engineering
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• v.4 no.2
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• pp.57-66
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• 1999

Numerical Analysis has been done for the supersonic off-design jet flow due to the pressure difference between the jet and the ambient fluid. The difference of pressure generates an oblique shock or an expansion wave at the nozzle exit. The waves reflect repeatedly on the center axis and the sonic surface in the shear layer. The pressure difference is resolved across these reflected waves. In this paper, the axi-symmetric Navier-Stokes equation has been used with the κ-ε turbulence model. The second order TVD scheme with flux limiters, based on the flux vector split with the smooth eigenvalue split, has been used to capture internal shocks and other discontinuities. Numerical calculations have been done to analyze the off-design jet flow due to the pressure difference. The variation of pressure along the flow axis is compared with an experimental result and other numerical result. The characteristics of the interaction between the shock cell and the turbulence mixing layer have been analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.105-110
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• 2010

Supersonic ejectors are simple mechanical components, which generally perform mixing and recompression of two fluid streams. Ejectors have found many applications in engineering. In aerospace engineering, they are used for high altitude testing (HAT) of a propulsion system by reducing the pressure of a test chamber. It is composed of three major sections: a vacuum test chamber, a propulsive nozzle, and a supersonic exhaust diffuser (SED). This paper aims at the improvement in HAT facility by focusing attention on the vertical firing rocket test stand with shock generators. Shock generators are mounted inside the SED for improving the pressure recovery. The results clearly showed that the performance of the ejector-diffuser system was improved with the addition of shock generators. The improvement comes in the form of reduction of the starting pressure ratio and the vertical height of test stand. It is also shown that shock generators are useful in reducing the total pressure loss in the SED.

• 한국전산유체공학회:학술대회논문집
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• 1999.05a
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• pp.85-89
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• 1999

Numerical Analysis has been done for the supersonic off-design jet flow due to the pressure difference between the jet and the ambient fluid. The difference of pressure generates an oblique shock or an expansion wave at the nozzle exit, The waves reflect repeatedly at the center axis and on the sonic surface in the shear layer, and the pressure difference is resolved across these waves interacted with the turbulence mixing layer. In this paper, the axi-symmetric Navier-Stokes equation has been used with two equation $k-{\varepsilon}$ turbulence closure model. The second order TVD scheme with flux limiters, based on the flux vector split by the smooth eigenvalue split, has been used to capture internal shocks and other discontinuities. The correction term for the compressible flow and the damping function are used in the turbulence model. Numerical calculations have been done to analyze the off-design jet flow due to the pressure difference. The variation of pressure along the flow axis is compared with an experimental result and other numerical result. The characteristics of the interaction between the shock cell and the turbulence mixing layer have been analyzed.