• International Journal of Aeronautical and Space Sciences
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• v.5 no.2
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• pp.18-27
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• 2004

An axisymmetric supersonic jet is simulated at a Mach number 2.1 and a Reynolds numberof 70000 to identify the mechanism of Mach wave generation and radiation from the jet. In orderto provide the near-field radiated sound directly and resolve the large-scale vortices highly.high-resolution essentially non-oscillatory(ENO) scheme, which is one of the ComputationalAeroAcoustics(CAA) techniques, is newly employed. Perfectly expanded supersonic jet is selectedas a target to see pure shear layer growth and Mach wave radiation without effect of change injet cross section due to expansion or shock wave generated at nozzle exit. The sound field ishighly directional and dominated by Mach waves generated near the end of potential core. Thenear field sound pressure levels as well as the aerodynamic properties of the jet, such asmean-flow parameters are in fare agreement with experimental data.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.29 no.1 s.232
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• pp.80-86
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• 2005

This study presents numerical solutions of the two-dimensional Navier-Stokes equations for supersonic unsteady flow in a convergent-divergent nozzle with heat addition. The TVD scheme in generalized coordinates is employed in order to calculate the moving shock waves caused by thermal choking. We discuss on transient characteristics, start and unstart phenomena, fluctuations of specific thrust caused by thermal choking and viscous effects. We prove that the control of separation of boundary layer is the most important key problem to prevent the thermal choking.

• 한국전산유체공학회:학술대회논문집
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• 2003.08a
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• pp.126-131
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• 2003

A series of computational simulations have been carried out for supersonic flows in a scram jet engine with and without a cavity. Transverse injection of hydrogen, a simplest form of fuel supply, is considered in the present study with the injection pressure varying from 0.5 to 1.5 MPa. The corresponding equivalence ratios are 0.167 - 0.50. The work features detailed resolution of the flow dynamics in the combustor, which was not typically available in most of the Previous studies. In particular, oscillatory flow characteristics are captured at a scale sufficient to identify the underlying physical mechanisms. Much of the flow unsteadiness is related not only to the cavity, but also to the intrinsic unsteadiness in the flowfield. The interactions between shock waves and shear layer may cause a large excursion of flow oscillation. The role of the cavity and injection pressure are examined systematically.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.31 no.2 s.257
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• pp.109-115
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• 2007

Airframe-integrated Scramjet engines of NASA Langley type consist of a compressor, a combustion chamber and a nozzle. When some disturbances occur in one module of the engine, its influences are propagated to other modules. In this study, it is investigated numerically how shock waves were caused by thermal choking in one module propagate upstream and how they influence adjacent modules. The calculations are carried out in 2-dimensional supersonic viscous flow model using explicit TVD scheme in generalized coordinates. The adverse pressure gradient caused by heat addition brings about separation of the wall boundary layers and formation of the oblique shock wave that proceed to upstream. This moving shock wave formed one module blocks the flow coming into the adjacent modules, which makes the modules unstarted.

• 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 of Propulsion Engineers
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• v.8 no.3
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• pp.53-60
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• 2004

In order to investigate the nozzle - rotor interactions and the effect of partial admission, the flows in supersonic turbine rotor cascades with a nozzle have been computed. Extensive computations of partial admission supersonic turbines provide the shock structures and flow patterns in the nozzle and rotor. The governing equations were discretized with Euler implicit method in time and 2nd-order upwind scheme of FVM in space. The $\kappa$-$\varepsilon$ turbulence model was utilized to describe the turbulent flow field. It is clearly shown that the nozzle flow is highly affected by the shocks or expansion waves propagated from the rotor leading edge. And the rotor flow is also affected by the shocks or wakes originated from the nozzle.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.21 no.10
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• pp.1273-1283
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• 1997

Experiments of shock wave/turbulent boundary layer interaction were conducted by using a supersonic wind tunnel. Nominal Mach number was varied in the range of 1.6 to 3.0 by means of different nozzles. The objective of the present study is to investigate the effects of boundary layer flow bleed on the interaction flow field in a straight tube. Two-dimensional slits were installed on the tube walls to bleed the turbulent boundary layer flows. The bleed flows were measured by an orifice. The ratio of the bleed mass flow to main mass flow was controlled within the range of 11 per cent. The wall pressures were measured by the flush mounted transducers and Schlieren optical observations were made for almost all of the experiments. The results show that the boundary layer flow bleed reduces the multiple shock waves to a strong normal shock wave. For the design Mach number of 1.6, it was found that the normal shock wave at the position of the silt was resulted from the main flow choking due to the suction of the boundary layer flow.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.64-64
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• 2002

Two dimensional velocity distributions outside a Mach 2.0 supersonic nozzle have been investigated using digital particle image velocimetry (PIV). Mean velocities, turbulence intensities, vorticity field and volume dilatation field are obtained from 300 instantaneous PIV images using 0.33 $\mu\textrm{m}$ $TiO_2$ particle. The seeding particle of larger size, 1.4 $\mu\textrm{m}$ $TiO_2$, is also used for the experimental measurements of velocity lag downstream of shock waves according to particle sizes. The results have been compared and analyzed with schlieren photographs and computational fluid dynamics (CFD) results for the velocity distribution, the locations of shock waves and over-expanded shock structure. It was shown that the locations of normal shock and shock waves can be resolved by the axial or radial velocities, and the velocity lag is more significantly increased due to particle inertia as a particle size increases. And it was also found that over-expanded shock structures call be predicted by volume dilatation fields, and streamwise turbulence intensities are influenced significantly by normal shock waves.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.1
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• pp.132-142
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• 2000

The objective of this study is to experimentally investigate the noise propagating characteristics, the noise reduction mechanism and the performance of a slotted tube attached at the exit plane of a circular convergent nozzle. The experiment is performed through the systematic change of the jet pressure ratio and the slot length under the condition of two kinds of open area ratios, 25% and 51%. The open area ratio calculated by the tube length equivalent for the slot length is defined as the ratio of the total slot area to the surface area of a slotted tube. The experimental results for the near and far field sound, the visualization of jet structures and the static pressure distributions in the jet passing through a slotted tube are presented and explained in comparison with those for a simple tube. The propagating characteristics of supersonic jet noises from the slotted tube is closely connected with the slot length rather than the open area ratio, and its propagating pattern is similar to the simple tube. It is shown that the slotted tube has a good performance to suppress the shock-associated noise as well as the turbulent mixing noise in the range of a limited jet pressure and slot dimension. The considerable suppression of the shock‘associated noise is mainly due to the pressure relief caused by the high-speed jets passing through the slots on the tube. Both the strength of shock waves and the interval between them in a jet plume are decreased by the pressure relief. Moreover, the pressure relief is divided into the gradual and the sudden relief depending upon the open area ratio of the slotted tube. Consequently, the shock waves in a jet plume are also changed by the type of pressure relief. The gradual pressure relief caused by the slotted tube with the open area ratio 25% generates the weak oblique shock waves. On the contrary, the weak normal shock waves appear due to the sudden pressure relief caused by the slotted tube with the open area ratio 51%.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.67-74
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• 1999

Problems created by supersonic jet impinging on solid objects or ground arise in a variety of situations. For example multi-stage rocket separation, deep-space docking, V/STOL aircraft, jet-engine exhaust, gas-turbine blade, terrestrial rocket launch, and so on. These impinging jet flows generally contain a complex structures. (mixed subsonic and supersonic regions, interacting shocks and expansion waves, regions of turbulent shear layer) This paper describes experimental works on the phenomena (surface pressure distribution, flow visualization) when underexpanded supersonic jets impinge on the perpendicular, inclined plate using a supersonic cold-(low system. The used supersonic nozzle is convergent-divergent type, exit Mach number 2, The maximum on the plate when it was inclined was much larger than perpendicular plate, owing to high pressure recoveries through multiple shocks. Surface pressure distribution as to underexpanded ratio showed similar patterns together.