• Transactions of the Korean Society of Mechanical Engineers B
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• v.40 no.2
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• pp.93-102
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• 2016

An experimental study has been conducted to verify the startup characteristic of a Center Body Diffuser (CBD) for simulating a low pressure environment when at high altitudes. Vacuum chamber pressure and startup characteristics of the CBD were investigated according to various geometries of the center body structure by a cold gas flow test. The test results show that the startup pressure is lowest when the center body contraction angle is approximately $15^{\circ}$. The startup characteristic of the CBDs significantly improves when the diffuser inlet length ($L_d/D_d$) is decreasing and the divergence length ($L_s$) is increasing. Additionally, it is possible to simulate various high altitude, low pressure conditions for various rocket engines that have different nozzle expansion ratios by adjusting the center body's position inside the diffuser.

• The Journal of the Acoustical Society of Korea
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• v.37 no.4
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• pp.163-173
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• 2018

This paper performed the prediction of the acoustic loads applied to the surface of the flight vehicle during flight. Acoustic loads during flight arise from the pressure fluctuations on the surface of body. The conventional method of predicting the acoustic loads in flight uses semi-empirical method derived from theoretical and experimental results. However, there is a limit in obtaining the flow characteristics and the boundary layer parameters of the flight vehicle which are used as the input values of the empirical equation through experiments. Therefore, in this paper, we use the hybrid method which combines the results of CFD (Computational Fluid Dynamics) with semi-empirical methods to predict the acoustic loads acting on flight vehicle during flight. For the flight vehicle with cone-cylinder-flare shape, acoustic loads were estimated for the subsonic, transonic, supersonic, and Max-q (Maximum dynamic pressure) condition flight. For the hybrid method, two kind of boundary layer edge estimation methods based on CFD results are compared and the acoustic loads prediction results were compared according to empirical equations presented by various researchers.

• Journal of computational fluids engineering
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• v.19 no.2
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• pp.86-92
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• 2014

In this paper, prediction of separation trajectory for Two-stage-To-Orbit space launch vehicle has been numerically simulated by using an aerodynamic database based on steady state analysis. Aerodynamic database were obtained for matrix of longitudinal and vertical positions. The steady flow simulations around the launch vehicle have been made by using a 3-D RANS flow solver based on unstructured meshes. For this purpose, a vertex-centered finite-volume method was adopted to discretize inviscid and viscous fluxes. Roe's finite difference splitting was utilized to discretize the inviscid fluxes, and the viscous fluxes were computed based on central differencing. To validate this flow solver, calculations were made for the wind-tunnel experiment model of the LGBB TSTO vehicle configuration on steady state conditions. Aerodynamic database was constructed by using flow simulations based on test matrix from the wind-tunnel experiment. ANN(Artificial Neural Network) was applied to construct interpolation function among aerodynamic variables. Separation trajectory for TSTO launch vehicle was predicted from 6-DOF equation of motion based on the interpolated function. The result of present separation trajectory calculation was compared with the trajectory using experimental database. The predicted results for the separation trajectory shows fair agreement with reference[4] solution.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.9
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• pp.27-34
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• 2004

A computational study on the supersonic flow around the lateral jet controlled missile has been performed. For this purpose a three dimensional Navier-Stokes computer code(AADL3D) has been developed and case studies have been performed by comparing the normal force coefficient and the moment coefficient of a missile body for several parameters such as angles of attack, circumferential jet positions, and spouting jet angles. Missile surface is divided into four regions with respect to the center of gravity, and the normal force and moment distribution at each region are compared. The results show different behavior of the normal force and moment variation according to each parameter. Furthermore, it is shown that the pitching moment can be minimized through proper combination of each parameter.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.3
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• pp.187-194
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• 2015

Course correction munition is a smart projectile which improves its accuracy by the control mechanism equipped in the fuze section with canard. In this paper, various aerodynamic configurations of the fuze section were analysed by utilizing a semi-empirical method and a CFD method. A final canard configuration showing the least drag was then determined. During the CFD simulation, it was found that the k-${\omega}$ SST turbulence model combined with O-type grid base is suitable for the prediction of the base drag. Finally, the aerodynamic characteristics of the smart munition and the change of drag due to the canard installation were analysed.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.4
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• pp.50-56
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• 2012

Vent mixer can provide main flow directly into a recirculation region downstream of the mixer to enhance fuel-air mixing efficiency. Based on experimental results of three-dimensional velocity, vorticity and turbulent kinetic energy obtained by a stereoscopic PIV method, the performance of the vent mixer was compared with that of the step mixer which was used as a basic model. Thick shear layers of the vent mixer induced the increase of the penetration height. The turbulent kinetic energy mainly distributed along a boundary layer between the main flow and the jet plume. This turbulent field activates mass transfer in a mixing region, leading to the mixing enhancement.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.94-104
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• 2013

A numerical analysis has been conducted to determine the pitch thrust vectoring characteristics of a two-dimensional convergent divergent exhaust nozzle for supersonic aircraft application. The numerical analysis was done by using Fluent and verified by the experimental test results. Analysis was performed with pitch angle of $0^{\circ}$ and $20^{\circ}$ each at the ambient temperature condition. To see the effect of a ratio of pitch flap length and pitch flap height, the ratio was varied from 0.5 to 2.5. The numerical analysis shows that pitch thrust is changed greatly with pitch flap length. The big difference of the pitch thrust with pitch flap length is due to the shock interactions in the nozzle.

• Korean Journal of Metals and Materials
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• v.49 no.3
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• pp.221-230
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• 2011

The cold spray process is an emerging technology that utilizes high velocity metallic particles for surface coating. Metallic powder particles are injected into a converging-diverging de Laval nozzle and accelerated to a high velocity by a supersonic gas flow. The cold spray process normally uses a circular nozzle that has a rather narrow spraying range. To overcome this fault, a slit nozzle was considered in this study. The slit nozzle is anticipated to reduce the coating process time because it has a wider coating width than the circular nozzle. However, the slit nozzle can reduce the coating efficiency because it does not allow as much gas and particle velocity as the circular nozzle. To improve the coating efficiency of a slit nozzle, the shape of the slit nozzle was modified. And the results of gas flow and particle behaviour according to the nozzlers shape were compared by the a numerical analysis. As a results, as Expansion Ratio(ER) of 7.5 was found to be the most optimal condition for enhancing the spraying efficiency when the ER was changed by the variation of nozzle neck and exit size.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.4
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• pp.316-329
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• 2012

In this paper, optimum design process of ICP (Inductively Coupled Plasma) torch, which has been used widely in aerospace application, such as supersonic plasma wind tunnel, is presented. For this purpose, the behaviors of equivalent circuit parameters (equivalent resistance and inductance, coupling efficiency) were investigated according to the variations of torch design parameters (frequency, $f$, confinement tube radius, $R$ and coil turn numbers, $N$) in the basis of analytical and numerical MHD (Magneto Hydro-Dynamics) models combined with electrical circuit theory. From the results, it is found that equivalent resistance is increased with the increase of $f$ values but vice versa for equivalent inductance. For elevated values of $R$ and $N$, however, both parameters tend to increase. Based on these observations, ICP torch with a power level of 10 kW can be optimized at the design ranges of $f$=4~6 MHz, $R$=17~25 mm and $N$=3~4 to maximize the electrical coupling efficiency, which is the ratio of equivalent resistance to equivalent inductance.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.38 no.3
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• pp.219-225
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• 2014

In this study, to find the characteristics of the oscillation of a terminating shock wave in a transonic airfoil flow with non-equilibrium condensation, a NACA00-12,14,15 airfoil flow with non-equilibrium condensation is investigated through numerical analysis of TVD scheme. Transonic free stream Mach number of 0.81-0.90 with the variation of stagnation relative humidity and airfoil thickness is tested. For the free stream Mach number 0.87 and attack angle of ${\alpha}=0^{\circ}$, the increase in stagnation relative humidity attenuates the strength of the terminating shock wave and inactivates the oscillation of the terminating shock wave. For the case of $M_{\infty}=0.87$ and ${\phi}_0=60%$, the decreasing rate in the frequency of the shock oscillation caused by non-equilibrium condensation to that of ${\phi}_0=30%$ amounts to 5%. Also, as the stagnation relative humidity gets larger, the maximum coefficient of drag and the difference between the maximum and minimum in $C_D$ become smaller. On the other hand, as the thickness of the airfoil gets larger, the supersonic bubble size becomes bigger and the oscillation of the shock wave becomes higher.