• 한국해양학회지
• /
• v.28 no.2
• /
• pp.86-100
• /
• 1993

Three-dimensional baroclinic hydrodynamic model, BACHOM-3, is developed using ADI finite difference scheme. The model is applied to a uni-nodal standing wave in a rectagular basin. The model results for the surface elevation and velocities coincide with the analytical results. To verify the field applicability of the model and to investigate the flow patterns of the Suyoung Bay in Pusan, Korea, the model is applied to the bay. The numerically predicted velocity predicted velocity fields during spring tide at normal river flow are compared with field measurements, the comparisons show good agreement. A clockwise residual circulations at the first level (depth = 0∼2m) and the second level (depth=2∼5 m) of the central part of the bay occur, and the ebb flow is stronger than the flood flow. Computed velocity fields show that the phase difference of velocities between the surface layer and bottom layer occurs and the phase lag increases with height from the bottom. Then, the model is applied successfully for the computation of flow fields considering flood river flow and wind effects. When the wind is blowing toward the land from the sea, the flow patterns at the surface layer correspond with the wind direction, but the flow patterns at the near solid boundary of the lower layer show opposite currents to the wind direction.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.38 no.8
• /
• pp.747-757
• /
• 2010

The ice accreted on the airfoil is one of the critical drivers that causes the degradation of aerodynamic performance as well as aircraft accidents. Hence, an efficient numerical code to predict the accreted ice shape is crucial for the successful design of de-icing and anti-icing devices. To this end, a numerical code has been developed for the prediction of glaze ice accretion shape on 2D airfoil. Constant Source-Doublet method is used for the purpose of computational efficiency and heat transfer in the icing process is accounted for by Messinger model. The computational results are thoroughly compared against available experiments and other computation codes such as LEWICE and TRAJICE. The direction and thickness of ice horn are shown to yield similar results compared to the experiments and other codes. In addition, the effects of various parameters - temperature, free-stream velocity, liquid water contents, and droplet diameter - on the ice shape are systematically analyzed through parametric studies.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.43 no.3
• /
• pp.204-212
• /
• 2015

The infrared lock-on range of target aircraft plays a critical role in determining the aircraft survivability. In this investigation, the effects of various UAV engine nozzle configurations on the aircraft lock-on range were theoretically analyzed. A virtual subsonic aircraft was proposed first, based on the mission requirement and the engine performance analysis, and convergent-type nozzles were then designed. After determining thermal flow field and nozzle surface temperature distribution with the CFD code, an additional analysis was conducted to predict the IR signature. Also, atmospheric transmissivity for various latitude and seasons was calculated, using the LOWTRAN code. Finally, the lock-on and lethal envelopes were calculated for different nozzle configurations, assuming the sensor threshold of the given IR guided missile. It was shown that the maximum 55.3% reduction in lock-on range is possible for deformed nozzles with the high aspect ratio.

• Solar Energy
• /
• v.12 no.3
• /
• pp.47-59
• /
• 1992

In the concentric tube type recuperator, which is the most typical type of radiation recuperator, installed on the exhaust-secion of industrial furnace, air flows between the adiabatic outer tube and the inner tube in which exhausted gas flows with high temperature. The waste heat of the exhausted gas is transferred to the inner tube, and transferred from the inner tube to the flowing air. The heat transfer by radiation In the concentric tube type recuperator is modeled using spherical harmonics approximation, namely, P-N method and numerically analyzed considering the effect of dynamic flow field. The results are compared with the existing empirical data. In addition, a theoretical method is presented for the analysis of the heat transfer characteristics of a recuperator with a reradiator installed in the inner tube, which causes re-radiant in the inner tube, and the characteristics of the recuperator is analyszed and defined.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.6
• /
• pp.75-81
• /
• 2014

The flame deflector should be constructed to minimize the induced environmental effects on the launch vehicle and to minimize the exhaust impingement effects on the launch complex structures during the lift-off operation. Therefore, it should be designed to avoid recirculation and reverse flow of rocket exhaust plumes. The circumstance around launch complex and characteristics of launch vehicle should be taken into consideration for the flame deflector design. In this paper, we designed the flame deflector reflecting KSLV-II 1st engine characteristics and analyzed the effect of exhaust plumes related to change geometry by means of computational flow analysis.

• Transactions of the Korean Society of Mechanical Engineers B
• /
• v.35 no.8
• /
• pp.763-771
• /
• 2011

An optimization procedure for the design of a ventilation axial-flow fan is presented in this paper. Flow analyses of the preliminary fan are performed by solving three-dimensional Reynolds-averaged Navier-Stokes equations via a finite-volume solver with the shear-stress transport turbulence model as a turbulence closure. Three variables, the hub-to-tip ratio and the stagger angles at the mid and tip spans, are selected for the optimization. The Latin-hypercube sampling method as a design-of-experiments technique is used to generate twenty-five design points within the design space. and the weighted average surrogate models, WTA1, WTA2, and WTA3, are applied for find optimal designs. The results show that the efficiency is considerably enhanced.

• Journal of Aerospace System Engineering
• /
• v.13 no.5
• /
• pp.32-38
• /
• 2019

The flying wing configuration with high sweep angles and rounded leading edge represent a complex flow of structures by the leading edge vortex. For control of the tailless flying wing configuration with unstable directional stability, flaperon is used. In this study, we conducted numerical simulations for a non-slender flying wing configuration with a rounded leading edge and analyzed the effect of the sideslip angle and flaperon. Through aerodynamic coefficient analysis, it was found that the effect of AoS on lift and drag coefficient was minimal and the side force and moment coefficient were markedly influenced by AoS. As the sideslip angle increased, the pitch break, which is related to the pitching moment coefficient, was delayed. Through stability analysis, the directional and lateral static stability of the flying wing configuration were increased by flaperon. Also, the structure and behavior of the leading edge vortex were analyzed by observing the contour of the pressure coefficient and the skin friction line.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.19 no.3
• /
• pp.39-46
• /
• 2015

A numerical study has been conducted on CO after burning and NO generation of the rocket plume as the cooling water injected to the rocket plume. The present study shows that the cooling water has a role of increasing the degree of CO after burning and reducing NO generation. However the effect varies as the injection configuration of the cooling water. When the cooling water is injected at the side of the plume, NO generation is dramatically reduced while the degree of CO after burning is relatively low. When the cooling water is injected at the side and the center of the plume, CO after burning is highly increased and NO generation is also dramatically reduced.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.5
• /
• pp.62-69
• /
• 2014

Infrared signature of rocket plume plays an important role for detection, recognition, tracking and minimzing for low observability. Infrared signatures of rocket plume with reduced smoke propellant and smokeless propellant are measured. In order to estimate the infrared signature of rocket plume, CFD analysis for flow structure of plume is performed, and layered integration method for estimating of infrared signature is used. Numerical and experimental results were in good agreement. Both propellants had similar infrared signature. Strong peak at $4.3{\mu}m$ region in the experimental results is appeared due to experimental error arising from the calibration procedure.

• Journal of Aerospace System Engineering
• /
• v.13 no.1
• /
• pp.18-28
• /
• 2019

High speed vehicles are subjected to high thermal loadings due to aerodynamic heating during ascent and reentry. Since a thermal protection system panel is mechanically constrained, it may cause thermal buckling under excessive thermal loadings. The thermal buckling could disturb the field of flow and make aerodynamic characteristics unstable. It is thus necessary to design the thermal protection system panel to prevent thermal buckling. This study defines the analytical model of temperature distribution using the finite difference method for the thermal protection system panel with large temperature differences inside and outside. This paper proposes the approximate model of the thermal buckling characteristics for the thermal protection system panel through the use of the Ritz method. The validity of the present method was verified by comparing the results of the finite element analysis. Furthermore, this research performs the parametric analysis of the thermal buckling characteristics for the thermal protection system panel by using the approximate model.