• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.331-335
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• 2006

Theoretically, stable operations of an inlet are achieved at the design condition. However, at off-design conditions supersonic inlets often encounter the problem of aerodynamic instability, called inlet buzz. During inlet buzz, supersonic inlets exhibit considerable oscillation of the shock system in front of the inlet and corresponding large pressure fluctuations downstream. This phenomenon results in decrease of engine performance. An experimental and numerical study was conducted to investigate the phenomenon of supersonic inlet buzz on a generic, axisymmetric, external-compression inlet with a single-surface center-body. This study suggest that intermittent buzz exist and the frequency become to be large as increasing the back pressure.

• International Journal of Aeronautical and Space Sciences
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• v.4 no.1
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• pp.53-62
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• 2003

In this study, nonlinear aeroelastic characteristics of an supersonic missile wing with strong shock interferences are investigated. The missile wing model has a freeplay structural nonlinearity at its pitch axis. To practically consider the effects of freeplay structural nonlinearity, the fictitious mass method is applied to structural vibration analysis based on finite element method. Nonlinear aerodynamic flows with unsteady shock waves are also considered in supersonic flow regions. To solve the nonlinear aeroelastic governing equations including the freeplay effect, a modal-based coupled time-marching technique based on the fictitious mass method is used in the time-domain. Various aeroelastic computations have been performed for the nonlinear wing structure model. Linear and nonlinear aeroelastic analyses have been conducted and compared with each other in supersonic flow regions. Typical nonlinear limit cycle oscillations and phase plots are presented to show the complex vibration phenomena with simultaneous fluid-structure nonlinearities.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.29 no.12 s.243
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• pp.1285-1290
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• 2005

A theoretical analysis of an annular injection supersonic ejector equipped with a second-throat was developed under the assumption that the secondary flow is choked aerodynamically by interaction with primary flow in the mixing chamber. The predicted secondary flow pressure agrees reasonably well with the measurements. Using the analysis, the compression ratio, the secondary flow Mach number, and the location of the choking point were presented in terms of entrainment ratio.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2000.06a
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• pp.1198-1203
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• 2000

In this study, a nonlinear aeroelastic analysis system for the fighter wing with tip-store has been developed additionally in the transonic and supersonic flow region. The unsteady CFD code based on the transonic small disturbance theory has been incorporated to consider the numerical capability for the aerodynamic nonlinear effects. The coupled time-integration method is used to observe the detailed nonlinear aeroelastic responses for elastic wings in their flight. condition. A conservative wing-box model of a fighter wing with tip-store is modeled by MSC/PATRAN and the corresponding free vibration analysis has been performed by MSC/NASTRAN. The results of flutter analyses are presented in the subsonic, transonic and supersonic flow regime.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.1
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• pp.8-15
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• 2004

Supersonic jets impinging on $60^{\cire}$ wedge were investigated to obtain fundamental design data for jet deflector It was of interest to study flow phenomena such as shock interaction and separation induced by shear layer. Experiments using supersonic cold flow system were conducted for Schlieren flow visualization and measurement of surface pressure. Numerical results were compared with the experimental results. The major parameters are underexpansion ratio, distance from nozzle to apex and design Mach number. Flow conditions were obtained for the wedge shock to attach on or detach from the wedge. The dominant feature of flow-field is shock pattern induced by the Interaction between the wedge shock and the barrel shock.

• Journal of the Korea Institute of Military Science and Technology
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• v.10 no.4
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• pp.12-19
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• 2007

In this study, nonlinear flow-induced vibration(flutter) analyses of a 2-DOF launch vehicle airfoil have been conducted in supersonic and hypersonic flow regimes. Advanced aeroelastic analysis system based on computational fluid dynamics and computational structural dynamics is successfully developed and applied to the present analyses. Nonlinear unsteady aerodynamic analyses considering strong shock wave motions are conducted using inviscid Euler equations. Aeroelastic governing equations for the 2-DOF airfoil system is solved by the coupled integration method with interactive CFD and CSD computation procedures. Typical wedge type airfoil shapes with initial angle-of-attacks are considered to investigate the nonlinear flutter characteristics in supersonic(15). Also, the comparison of detailed aeroelastic responses are practically presented as numerical results.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.777-780
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• 2011

The supersonic wind tunnel is a common facility to studies the aerodynamic phenomenon around the high speed vehicle or weapon system whose operating speed is greater than sonic speed. In this study, a design procedure and selecting the components of a new supersonic wind tunnel whose nozzle exit is $125mm{\times}100mm$ is considered. An operating test of this wind tunnel is being conducted to compare the result with the design values, mach number, etc.

• Journal of computational fluids engineering
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• v.11 no.4 s.35
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• pp.62-66
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• 2006

High-speed flight vehicle have various cavities. The supersonic cavity flow is complicated due to vortices, flow separation, reattachment, shock waves and expansion waves. The general cavity flow phenomena includes the formation and dissipation of vortices, which induce oscillation and noise. The oscillation and noise greatly affect flow control, chemical reaction, and heat transfer processes. The supersonic cavity flow with high Reynolds number is characterized by the pressure oscillation due to turbulent shear layer, cavity geometry, and resonance phenomenon based on external flow conditions. The resonance phenomena can damage the structures around the cavity and negatively affect aerodynamic performance and stability. In the present study, we performed numerical analysis of cavities by applying the unsteady, compressible three dimensional Reynolds-Averaged Navier-Stokes(RANS) equations with the ${\kappa}-{\omega}$ turbulence model. The cavity model used for numerical calculation had a depth(D) of 15mm cavity aspect ratio (L/D) of 3, width to spanwise ratio(W/D) of 1.0 to 5.0. Based on the PSD(Power Spectral Density) and CSD(Cross Spectral Density) analysis of the pressure variation, the dominant frequency was analyzed and compared with the results of Rossiter's Eq.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.1
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• pp.23-29
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• 2006

In an annular injection supersonic ejector, the supersonic primary flow is injected along the side wall, therefore a funnel-shaped shock wave is generated by the contraction angle of the mixing chamber. In the present study, we developed a simple funnel shock wave model using 2-D wedge and conical shock wave relations. In result, the secondary flow pressure can be predicted more accurately than using a simple 2-D wedge shock wave model. Through the same analysis, the compression ratio and the adiabatic efficiency according to the entrainment ratio were calculated.