• Proceedings of the KSME Conference
• /
• 2003.11a
• /
• pp.544-549
• /
• 2003

When a shock wave arrives at an open end of tube, an impulse wave is discharged from the tube exit and causes serious noise and vibration problems. In the current study, the effect of the cross-sectional area of tube on the impulse wave is numerically investigated using a CFD method. The Harten-Yee's total variation diminishing(TVD) scheme is used to solve the axisymmetric, two-dimensional, unsteady, compressible Euler equations. With three different cross-sectional areas of tube, the Mach number of the incident shock wave $M_{s}$ is varied between 1.01 and 1.5. The results obtained show that the directivity and magnitude of impulse wave strongly depend upon the Mach number of incident shock wave and are influenced by the tube area. It is also known that the tube cross-sectional area significantly affects the magnitude of impulse wave at or near the tube axis.

• Journal of KSNVE
• /
• v.10 no.6
• /
• pp.1035-1040
• /
• 2000

The Impingement of a weak shock wave discharged from the open end of a shock tube upon a flat plate was investigated using shock tube experiments and numerical simulations. Harten-Yee Total Variation Diminishing method was used to solve axisymmetric, unsteady, compressible flow governing equations. Computations predicted the experimented results with a good accuracy. The peak pressure on the flat plate was not strongly dependent of the shock wave Mach number in the present range of Mach Number from 1.05 to 1.20. The distance between the plate and shock tube was changed to investigate the effect on the peak pressure. From both the results of experiments and computations we obtained a good empirical equation to predict the peak pressure on the flat plate.

• Proceeding of EDISON Challenge
• /
• 2014.03a
• /
• pp.544-549
• /
• 2014

임계마하수보다 큰 자유흐름 마하수에서는 충격파의 발생으로 인해 급격한 항력증가가 발생하므로, 임계마하수 증가는 고속 공기역학에서 중요한 분야로 다뤄지고 있다. Whitcomb R. T.에 의해 천음속영역에서 순항할 수 있는 초임계익형이 개발되었으나, 충격파 제어 기법들에 대한 실험적인 검증은 형상 제작의 어려움으로 인해 한계를 지닌다. 따라서 본 논문에서는 2D_Comp-2.1_P와 Prandtl-Glauert 압축성 보정식을 이용하여 NACA0012와 RAE2822의 임계마하수를 해석하고, 충격파 제어 장치 중 하나인 Bump를 RAE2822에 설치하여 임계마하수를 향상시키기 위한 연구를 수행하였다. 연구 결과 충격파를 압축파로 분산시켜 충격파의 강도를 약화시키고, 양항비의 4.7% 증가를 확인하였다. 따라서 Bump를 설계한 RAE2822가 기본 익형보다 높은 천음속 조건에서 효율적인 공력특성을 가지는 것을 확인하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.04a
• /
• pp.299-302
• /
• 2007

The aerodynamics of a projectile overtaking a moving shock wave is analyzed using a chimera scheme. The flow field characteristics for various shock wave Mach number and projectile masse are investigated. the unsteady forces acting on the projectile for both supersonic and impossible overtaking conditions are computed in order to analyze the aerodynamic characteristics of the projectile. It is seen that the projectile Mach number significantly affects the flow fields for both supersonic and impossible overtaking. Unsteady drag is influenced by the overtaking conditions. The unsteady drag coefficient is the highest for the impossible overtaking condition.

• International Journal of Fluid Machinery and Systems
• /
• v.4 no.1
• /
• pp.97-103
• /
• 2011

With the aim to increase blade loadings and stable operating range in highly loaded compressors, this article has been conducted to explore, through a numerical parametric study, the potential of passive control using slotted bladings in cascade configurations. The objective of this numerical investigation is to analyze the influence of location, width and slope of the slots and therefore identify the optimal configuration. The approach is based on two dimensional cascade geometry, low speed regime, steady state and turbulent RANS model. The results show the efficiency of this passive technique to delay separation and enhance aerodynamic performances of the compressor cascade. A maximum of 28.3% reduction in loss coefficient have been reached, the flow turning is increased with approximately $5^0$ and high loading over a wide range of angle of attack have been obtained for the optimized control parameter.

• 한국전산유체공학회:학술대회논문집
• /
• 1999.05a
• /
• pp.137-143
• /
• 1999

The unsteady transonic viscous flow has been analyzed over an oscillatory airfoil. The CSCM(Conservative Supra Characteristic Method) upwind flux difference splitting method and the iterative time marching scheme having first order accuracy in time and second to third order accuracy in space was applied on dynamic meshes. A steady flow field of Mach number 0.7 has been calculated for the verification of unsteady algorithm. The time-accurate unsteady calculations have been done for NACA 0012 airfoil oscillating around quarter chord about freestream Mach number 0.6 on dynamic meshes. The results have been compared with the AGARD Case 3 experimental data. The periodic characteristics have been compared with the experimental results.

• 한국전산유체공학회:학술대회논문집
• /
• 2002.05a
• /
• pp.29-34
• /
• 2002

TIn this study, the counter-jet flows which designed for improvement of aerodynamic performance of the blunt body vehicle have been analyzed. The variations of the drag force and jet penetration depth due to changes in the stagnation properties of counter jet new such as total pressure, mach number, and total temperature. The counter jet flow, which is injected toward incoming supersonic freestream at stagnation region of blunt cone-cylinder vehicle, have been studied by using upwind flux difference splitting navier-stokes method. The changes in the stagnation pressure and Mach number resulted in large effects on the wall pressure and drag force, on the other hand tile total temperature changes did not.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.649-656
• /
• 2004

Hypersonic wind tunnel test of the rectangular variable geometry intake is performed. For realization of a Precooled turbojet engine, development of a hypersonic ramjet engine is planned. To investigate performance of the intake of the hypersonic ramjet engine, wind tunnel test is done with freestream Mach number of 5.1. The total pressure recovery was 18 % with 12.9 % of ramp bleed. Several reasons for low total pressure recovery are shown. Supersonic internal compression is not enough. Then, the throat Mach number is high (M2.61) and total pressure losses at the terminal shock is large. Supersonic flow at the throat and position of the terminal shock is sensitive to a difference of the second ramp's throat height and the third ramp's throat height. Flow separations at the second ramp's trailing edge and the third ramp's leading edge are seen those could result in the trigger of unstart. The seal mechanism between the ramps and the sidewalls is important.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.784-788
• /
• 2004

A visualization study of shock formation of the supersonic jet nozzle using a Shadowgraph Method (SM) was carried out to investigate the effect of the longitudinal variation of coaxial pipe end tip position inside the supersonic nozzle. The experiment was performed for the Mach number range from 1.1 to 1.2 at nozzle exit. The well known shock cell structure was shown with the pipe end located deep inside the nozzle for the studied Mach number. With the pipe end approaches nozzle exit, it was found that the shock cell structure disappeared and turned into complex formation. In order to understand the mechanism of the shock structural change, computational simulation was carried out using the Navier-Stokes solver, FLUENT. Topological sketch was added with an aid of the visualization and the numerical simulation.

• 한국항공운항학회:학술대회논문집
• /
• 2004.11a
• /
• pp.240-246
• /
• 2004

For the prediction on the onset of oblique shock wave-induced vortex breakdown, computational studies on the Oblique Shock wave/Vortex Interaction (OSVI) are conducted and compared with both experimental results and analytic model. A Shock-stable numerical scheme, the Roe scheme with Mach number-based function (RoeM), and a two-equation eddy viscosity-transport approach are used for three-dimensional turbulent flow computations. The computational configuration is identical to available experiment, and we attempt to ascertain the effect of parameters such as a vertex strength, streamwise velocity deficit, and shock strength at a freestream Mach number of 2.49. Numerical simulations using the ${\kappa}-{\omega}SST$ turbulence model and suitably modeled vortex profiles are able to accurately reproduce many fine features through a direct comparison with experimental observations. The present computational approach to determine the criterion on the onset of oblique shock wave-induced vortex breakdown is found to be in good agreement with both the experimental result and the analytic prediction.