• 한국전산유체공학회:학술대회논문집
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• 1998.05a
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• pp.117-122
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• 1998

A numerical study is carried out to investigate the transient process of combustion phenomena associated with hypersonic propulsion devices. Reynolds averaged Navier-Stokes equations for reactive flows are used as governing equations with a detailed chemistry mechanism of hydrogen-air mixture and two-equation SST turbulence modeling. The governing equations are discretized by a high order accurate upwind scheme and solved in a fully coupled manner with a fully implicit time accurate method. At first, oscillating shock-induced combustion is analyzed and the comparison with experimental result gives the validity of present computational modeling. Secondly, the model ram accelerator experiment was simulated and the results show the detailed transient combustion mechanisms. Thirdly, the evolution of oblique detonation wave is simulated and the result shows transient and final steady state behavior at off-stability condition. Finally, shock wave/boundary layer interaction in combustible mixture is studied and the criterion of boundary layer flame and oblique detonation wave is identified.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.12
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• pp.1025-1031
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• 2012

The present paper deals with concept design of supersonic inlet based on compressible flow theory and flow control of bleeding in order to guarantee stability of supersonic inlet of ramjet engine in broad range of operating conditions. Shock instability, shock wave-boundary layer interaction and flow separation should be properly controlled to improve performance of the supersonic inlet. Considering shock strength, boundary layer and flow separation, the supersonic inlet is modified from the basic model which is designed under inviscid theory. Consequently, shock is stabilized, and required mass flow rate is obtained. Furthermore, bleeding is applied to the supersonic inlet to maintain performance in off-design conditions. Mass flow condition is adopted for modeling of bleeding effect, and performance of the supersonic inlet is evaluated by changing bleeding locations and numbers.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.12
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• pp.18-24
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• 2006

Passive control of the shock wave/turbulent boundary-layer interaction utilizing slotted plates and a porous plate over a cavity has been carried out. Effect of various slot configurations on the characteristics of the interaction has been observed. Pitot/wall surface pressure distributions and flow visualizations including Schlieren images, kerosene-lampblack tracings and interference fringe patterns over a thin oil-film have been obtained at the downstream of the shock interactions. For the streamwise-slot configuration, a local higher pitot pressure was noticed at the downstream of the interaction as compared with the case of no control, however, not much improvement in pitot pressure was observed for the spanwise-slot configuration.

• Journal of the Korean Society of Visualization
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• v.13 no.3
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• pp.29-34
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• 2015

The interaction between a normal shock wave and a boundary layer along a wall surface in internal compressible flows causes a very complicated flow. This interaction region containing shock train and mixing region is called as pseudo-shock waves. Pseudo-shock waves in the divergent part of a rectangular nozzle have been investigated by using large-eddy simulation (LES). LES studies have been done for the complex flow phenomena of three-dimensional pseudo-shock waves. The LES results have been validated against experimental wall-pressure measurements. The LES results are in good agreement with experimental results. Pseudo-shock length and corner separation have been studied in three-dimensional LES model. Comparison of centerline pressure measurement and 3D visualization measurement has been discussed for the corner separation position. It has been concluded that the pseudo-shock length should be measured by using 3D visualization measurement.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.21 no.1
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• pp.174-184
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• 1997

A shock wave, being an irreversible process, gives rise to entropy increase. A great deal of effort has been made to control shock wave and boundary layer interaction related to energy losses as well as problems of vibration and noise. In the present study, tests are performed on a roof mounted half circular arc in an indraft type supersonic wind tunnel to evaluate the effects of porosity, length and depth of cavity in passive control of shock wave on the attenuation of shock strength by reviewing the measured static pressures at the porous wall and cavity. Also the flow field is visualized by a Schlieren system. The results show that in the present study the porosity of 8% produced the largest reduction of pressure fluctuations and that for the same porosity, the strength of shock wave decreases with the increasings of the depth and length of cavity.

• Journal of computational fluids engineering
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• v.1 no.1
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• pp.112-122
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• 1996

A finite element scheme using the concept of PISO method has been developed to solve the Navier-Stokes viscous flows in all speed range. This scheme includes development of new pressure equation that retains both the hyperbolic term related with the density variation and the elliptic term reflecting the incompressibility constraint. The present method is applied to the incompressible two-dimensional driven cavity flow problems(Re=100, 400 and 1,000). For compressible flows, the Carter plate problem(M=3 and Re=1,000) is computed. Finally, we have simulated the shock-boundary layer interaction(M=2 and Re=2.96×10/sup 5/), a more difficult problem, and compared its results with the experiment to demonstrate the shock capturing capability of the present solution algorithm.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.75-78
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• 2004

Various jet engines (Turbine engine family and RAM Jet engine) have been developed for high speed aircrafts. but their application to hypersonic flight is restricted by principle problems such as increase of total pressure loss and thermal stress. Therefore, the development of next generation propulsion system for hypersonic aircraft is a very important subject in the aerospace engineering field, SCRAM Jet engine based on a key technology, Supersonic Combustion. is supposed as the best choice for the hypersonic flight. Since Supersonic Combustion requires both rapid ignition and stable flame holding within supersonic air stream, much attention have to be given on the mixing state between air stream and fuel flow. However. the wider diffusion of fuel is expected with less total pressure loss in the supersonic air stream. So. in this study the direction of fuel injection is inclined 30 degree to downstream and the total pressure of jet is controlled for lower penetration height than thickness of boundary layer. Under these flow configuration both streams, fuel and supersonic air stream, would not mix enough. To spread fuel wider into supersonic air an aerodynamic force, baroclinic torque, is adopted. Baroclinic torque is generated by a spatial misalignment between pressure gradient (shock wave plane) and density gradient (mixing layer). A wedge is installed in downstream of injector orifice to induce an oblique shock. The schlieren optical visualization from side transparent wall and the total pressure measurement at exit cross section of combustor estimate how mixing is enhanced by the incidence of shock wave into supersonic boundary layer composed by fuel and air. In this study non-combustionable helium gas is injected with total pressure 0.66㎫ instead of flammable fuel to clarify mixing process. Mach number 1.8. total pressure O.5㎫, total temperature 288K are set up for supersonic air stream.

• Advances in aircraft and spacecraft science
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• v.5 no.5
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• pp.581-594
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• 2018

This paper is the follow-on of a previous paper by the author where it was pointed out that the forthcoming, manned exploration missions to Mars, by means of complex geometry spacecraft, involve the study of phenomena like shock wave-boundary layer interaction and shock wave-shock wave interaction also along the entry path in Mars atmosphere. The present paper focuses the chemical effects both in the shock layer and on the surface of a test body along the Mars orbital entry and compares these effects with those along the Earth orbital re-entry. As well known, the Mars atmosphere is almost made up of Carbon dioxide whose dissociation energy is even lower than that of Oxygen. Therefore, although the Mars entry is less energized than the Earth re-entry, one can expect that the effects of chemistry on aerodynamic quantities, both in the shock layer and on a test body surface, are different from those along the Earth re-entry. The study has been carried out computationally by means of a direct simulation Monte Carlo code, simulating the nose of an aero-space-plane and using, as free stream parameters, those along the Mars entry and Earth re-entry trajectories in the altitude interval 60-90 km. At each altitude, three chemical conditions have been considered: 1) gas non reactive and non-catalytic surface, 2) gas reactive and non-catalytic surface, 3) gas reactive and fully-catalytic surface. The results showed that the number of reactions, both in the flow and on the nose surface, is higher for Earth and, correspondingly, also the effects on the aerodynamic quantities.

• Journal of computational fluids engineering
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• v.20 no.4
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• pp.36-43
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• 2015

Accurate prediction of supersonic transition is required for the heat transfer estimation over supersonic double compression ramp flows. Correlation-based transition models were assessed for a supersonic double ramp problem. Numerical results were compared with experimental data from RWTH Aachen University. A parametric study on the nose bluntness was performed using a selected transition model. As the nose bluntness increases, the boundary layer thickness is increased and the triple point of shock interactions moves downstream. The peak magnitude of the heat transfer is consequently decreased with the nose bluntness.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.5
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• pp.375-384
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• 2012

This paper presents numerical results comparing the performance of the 2008 Wilcox $\mathcal{k}-{\omega}$ turbulence model to the one of the 1988 Wilcox $\mathcal{k}-{\omega}$ model for supersonic flows. A comparison with experimental data is offered for a shock wave/turbulent boundary layer interaction case and two ramp injector mixing cases. Furthermore, a comparison is performed with empirical correlations on the basis of skin friction for flow over a flat plate and shear layer growth for a free shear layer. It is found that the maximum injectant mass fraction of some ramp injector cases is better predicted using the 1988 Wilcox model. On the other hand, the 2008 model performs better in simulating shock-boundary layer cases.