• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 1998.10a
• /
• pp.12-12
• /
• 1998

무기체계의 추진체로서 고체 로켓트 추진기관이 제작의 용이성, 구조의 간단성, 이에 따른 저렴한 제작비, 그리고 고 신뢰도 확보가능 등의 여러 장점으로 대부분의 현존 전술 유도무기에 채택되어 사용되어 오고 있으나 대응, 방어 무기체계의 빠른 발전으로 이에 따른 새로운 추진기관의 유도무기체제가 요구되고 있다. 램제트 기관은 공기흡입추진기관으로 상대적으로 높은 비추력(1000-2000s)과 추력 중량비(∼20)을 가지며, 이로 인해 기존의 로켓 엔진에 비해 4-5배의 성능을 낼 수 있으며, 초음속 장거리 비행에 적합하다며, 또한 높은 속도영역까지 운용가능하고 구조가 비교적 간단하다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.393-396
• /
• 2006

This paper presents the characteristics of the performance parameters of dual mode(ramjet-scramjet combined) engine, inlet and combustor for wide range of flight Mach number, resulted from the analysis of its operating limitations. The transitional-critical flight Mach number from ramjet to scramjet and the performance of two types of combustors, such as constant pressure- and constant area- combustor, are conceptually evaluated.

• Publications of The Korean Astronomical Society
• /
• v.9 no.1
• /
• pp.21-38
• /
• 1994

We have developed a cylindrical mixing layer model of a stellar jet including cooling effect in order to understand an optical emission mechanism along collimated high velocity stellar jets associated with young stellar objects. The cylindrical results have been calculated to be the same as the 2D ones presented by Canto & Raga(1991) because the entrainment efficiency in our cylindrical model has been obtained to be the same value as the 2D model has given. We have discussed the morphological and physical characteristics of the mixing layers by the cooling effect. As the jet Mach number increases, the initial temperature of the mixing layer goes high because the kinetic energy of the jet partly converts to the thermal energy of the mixing layer. The initial cooling of the mixing layer is very severe, changing its outer boundary radius. A subsequent change becomes adiabatic. The number of the Mach disks in the stellar jet and the total radiative luminosity of the mixing layer, based on our cylindrical calculation, have quite agreed with the observations.

• 한국전산유체공학회:학술대회논문집
• /
• 2003.10a
• /
• pp.39-41
• /
• 2003

n this paper, direct numerical simulation of decaying compressible turbulence with passive scalar is performed by using 7th order upwind difference scheme or 8th order group velocity control scheme. The start Reynolds number (defined by Taylor scale) is 72 and turbulent Mach numbers are 0.2-0.9. The Schmidt numbers of passive scalar are 2-10. The Batchelor k-1 range are found in scalar spectra, and the high wavenumber spectra decays faster with increasing turbulent Mach number. The extend self-similarity (ESS) is found in the passive scalar in compressible turbulence.

• 한국전산유체공학회:학술대회논문집
• /
• 2008.03a
• /
• pp.337-339
• /
• 2008

As a shock impinges into a vortex of variable strength, complex shock diffraction can occur. Since a vortex has a fixed rotating direction, the shock wave travelling in one direction creates strong asymmetry in the vortex flow field. The process is that first the shock is divided into two parts by the vortex. One part is moving in the adverse direction opposite to the vortex flow which is captured by the vortex center. The other part is moving in the favorable direction, namely, in the direction same as the vortex flow; it is swung around the vortex, accelerating the vortex flow. In this paper we have investigated numerically using ENO scheme how and why the shock-vortex interaction patterns appear so different for different parametric values. Conclusion is that there are three different types of shock-vortex interaction depending on two related parameters: shock Mach number and vortex Mach number. We present a parameter map by which we can discern what type of interaction pattern appears as a shock impinges into a vortex.

• 한국전산유체공학회:학술대회논문집
• /
• 2008.03a
• /
• pp.251-255
• /
• 2008

Flow field around helicopter involves incompressible flow near the blade root and compressible flow at the blade tip. A problem occurs for low Mach number flow due to the stiffness of the governing equations. Time-derivative preconditioning techniques have been incorporated to reduce the stiffness that occurs at low speed region. The preconditioned form of the compressible Navier-Stokes and Euler equations is used. Computations are performed for the Caradonna-Tung's hovering and non-lifting forward flight case. Computational results are in good agreement with the experimental data.

• 한국전산유체공학회:학술대회논문집
• /
• 2004.10a
• /
• pp.7-14
• /
• 2004

It is well known that preconditioning methods are efficient for convergence acceleration at compressible low Mach number flows. In this study, the original Euler equations and three preconditioners nondimensionalized differently are implemented in two dimensional inviscid bump flows using the 3rd order MUSCL and DADI schemes as flux discretization and time integration respectively. The multigrid and local time stepping methods are also used to accelerate the convergence. The test case indicates that a properly modified local preconditioning technique involving concepts of a global preconditioning one produces Mach number independent convergence. Besides, an asymptotic analysis for properties of preconditioning methods is added.

• 한국전산유체공학회:학술대회논문집
• /
• 2004.10a
• /
• pp.43-46
• /
• 2004

Numerical analysis for pressure drag on boattail afterbodies have been studied by Mach number, boattail angle and length ratio of body diameter and base diameter using CFD-FASTRAN that the commercial external flow CFD code. The numerical results have been compared with the experimental data that have been shown pressure drag reduction and supersonic turbulent flow characteristics for boattail afterbodies. And the prediction equation tot boattail base drag has been made by the numerical results about Mach number and boattail configuration parameters.

• Journal of computational fluids engineering
• /
• v.11 no.4 s.35
• /
• pp.48-55
• /
• 2006

A computational study was carried out in order to investigate aerodynamic characteristics on leading edge sweepback angles of Flying-Wing configurations. The viscous-compressible Navire-Stokes equation and Spalart-Allmaras turbulence model of the commercial CFD code were adopted for this computation analysis. This investigation examined aerodynamic characteristics of three different types of leading edge sweepback angles: $30^{\circ}C,\;35^{\circ}C\;and\;40^{\circ}C$. The freestream Mach number was M=0.80 and the angle of attack ranged from ${\alpha}=0^{\circ}C\;to\;{\alpha}=20^{\circ}C$. The results show that the increases in sweepback angle of the Flying-Wing configuration creates more efficient aerodynamic performance.

• 한국전산유체공학회:학술대회논문집
• /
• 2006.10a
• /
• pp.110-119
• /
• 2006

An accurate and cost efficient method PSE is used for the stability analysis of 2D or 3D compressible boundary layers. A highly accurate finite difference PSE code has been developed at a general curvilinear coordinate system using an implicit marching procedure to deal with a broad range of transition predictions problems. Evolution of disturbances in compressible flat plate boundary layers are studied for free-stream Mach numbers ranging from 0 to 1.5. The effect of mean-flow nonparallelism is found to be weak on two dimensional waves and strong on three dimensional waves. The maximum amplification rate increases monotonically with Mach number. The present PSE solutions are compared with previous numerical investigations and experimental results and are found to be in good agreement.