• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.573-577
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• 2009

Dynamic behavior simulation of supersonic engine was performed and PI control algorithm was studied for the buzz control in the inlet and the thrust control. Firstly, required thrust was tracked according to the fuel flow control and then inlet pressure was regulated through the nozzle throat area control so that the buzz margin has the positive all the time. The control was performed according to the change of flight Mach number, altitude and angle of attack. The proportional gain and the integral gain for regulating the buzz margin was induced and simulated. In the results, it was confirmed and satisfied that control target in the operating area was changed the angle of attack from $0^{\circ}$ to $10^{\circ}$ at the flight Mach number of 2.1~3.0.

• Journal of The Korean Astronomical Society
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• v.45 no.5
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• pp.127-138
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• 2012

We explore how wave-particle interactions affect diffusive shock acceleration (DSA) at astrophysical shocks by performing time-dependent kinetic simulations, in which phenomenological models for magnetic field amplification (MFA), Alfv$\acute{e}$nic drift, thermal leakage injection, Bohm-like diffusion, and a free escape boundary are implemented. If the injection fraction of cosmic-ray (CR) particles is ${\xi}$ > $2{\times}10^{-4}$, for the shock parameters relevant for young supernova remnants, DSA is efficient enough to develop a significant shock precursor due to CR feedback, and magnetic field can be amplified up to a factor of 20 via CR streaming instability in the upstream region. If scattering centers drift with Alfv$\acute{e}$n speed in the amplified magnetic field, the CR energy spectrum can be steepened significantly and the acceleration efficiency is reduced. Nonlinear DSA with self-consistent MFA and Alfv$\acute{e}$nic drift predicts that the postshock CR pressure saturates roughly at ~10 % of the shock ram pressure for strong shocks with a sonic Mach number ranging $20{\leq}M_s{\leq}100$. Since the amplified magnetic field follows the flow modification in the precursor, the low energy end of the particle spectrum is softened much more than the high energy end. As a result, the concave curvature in the energy spectra does not disappear entirely even with the help of Alfv$\acute{e}$nic drift. For shocks with a moderate Alfv$\acute{e}$n Mach number ($M_A$ < 10), the accelerated CR spectrum can become as steep as $E^{-2.1}$ - $E^{-2.3}$, which is more consistent with the observed CR spectrum and gamma-ray photon spectrum of several young supernova remnants.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.3
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• pp.241-251
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• 2017

Computational fluid dynamics analysis was performed for the case 3 of the EFD-CFD workshop. Solvers were used for three commercial CFD codes(Star-CCM+, Fluent and CFX) and an open source CFD code(SU2). The grid were generated four types depending on the total cells using commercial grid generation code(Pointwise). Mach number of 0.4 and 0.8, 2 degree angle of attack and Mach number of 0.9, 1 degree angle of attack were calculated. Similar pressure coefficient curve and normal force coefficient were showed from the coarse grid to fine grid of four codes. But there is a difference in the drag coefficient. The position of the shock wave was predicted forward as the discretization order increased in calculations using Star-CCM+ and Fluent. The computation time to converge, Fluent, Star-CCM +, CFX are in order, and SU2 takes much time to converge.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.2
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• pp.109-117
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• 2020

A structured grid system can be efficiently constructed by applying the patched-grid algorithm that alleviates many constraints of the conventional structured grid system. Three approaches were applied to case 4 of the EFD-CFD workshop: delta wing-cylindrical body shape to solve the existing grid generation problems and verify the results by comparing them with experimental data. Surface pressure distributions slightly differed from the experimental data at high angles of attack. The slope variation of the pitching moment with Mach number is analyzed and the variation can be explained with the tuck under phenomenon. In the supersonic region, the bow shock waves in front of the shape expand the region generating lift up to the rear of the configuration. Also, the tendency of the pitching moment with both Mach number and angle of attack was analyzed by comparing the positions of the center of pressure and the center of gravity.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.1
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• pp.1-7
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• 2016

Most of the work has been done till now focused on flows over wall mounted cavities in a straight wall where the incoming flow is uniform. However, the investigation on such kind of flow over a cavity mounted on the curved walls has been seldom reported in the existing literatures. In the present study, the numerical analysis was performed to investigate the cavity flow mounted on the curved walls. The effects of wall shape, the curvature radius and the flow Mach number, were investigated for high-subsonic flows. The results show that the static pressure of cavity floor increases as the L/R increases. This effect is found to be more significant when the flow Mach number is higher. The cavity drag for the curved walls are higher as compared with that of straight wall.

• Journal of Advanced Marine Engineering and Technology
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• v.35 no.2
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• pp.175-181
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• 2011

The purpose of this study is to secure the design data for the optimization of the radial turbine and heat cycle system, by using the CFD analysis technique and the design of 100kW class radial turbine applicable to waste heat recovery generation system for ship. Radial turbine was comprised of scroll casing, vane nozzle with 18 blades and rotor with 13 blades, and analysis grid was used to about 2.3 million. Mass flow rate and rotational speed was 0.5kg/s, 75,0000rpm, respectively. Eight kinds of inlet pressure was set between 195 and 620kPa. As the flow accelerated through the nozzle passage to the throat, the pressure level at the pressure and suction sides becomed similar to about Mach number of 0.35. When the inlet temperature and pressure was $250^{\circ}C$, 352kPa respectively, the isentropic efficiency and mechanical power showed the analysis results of 74% and 108kW.

• The KSFM Journal of Fluid Machinery
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• v.14 no.6
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• pp.96-101
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• 2011

Korea Aerospace Research Institute is performing 3 stage transonic axial compressor development program. This paper introduces design step of the compressor, the performance test results and its analysis. In the fore part of the paper, aerodynamic process of the 3 stage axial compressor is presented. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3, while the maximum relative Mach number in the rotor stage is between 1.3~1.4 to increase the compressor flow coefficient. The final design has been confirmed by iterating three dimensional CFD calculations to verify design target and some design intentions. In the latter part of the paper, its performance test processes and results are presented. The performance test result shows that the overall compressor performance targets; pressure ratio and efficiency are well achieved. The stator static pressure distributions show that the blade loading is gradually increasing from the downstream of the compressor.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.12
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• pp.1702-1710
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• 2001

Supersonic coaxial jets are investigated numerically by using the axisymmetric, Wavier-Stokes equations which are solved using a fully implicit finite volume method. Three different kinds of coaxial nozzles are employed to understand the flow physics involved in the supersonic coaxial jets. Two convergent-divergent supersonic nozzles are designed to have the same Mach number 2.0, and used to compare the coaxial jet flows with those discharging from one constant-area nozzle. The impingement angle of the annular jets are varied. The primary pressure ratio is changed in the range from 2.0 to 10.0 and the assistant jet ratio from 1.0 to 3.0. The results obtained show that the fluctuations of the total pressure and Mach number along the jet axis are much higher in the constant-area nozzle than those in the convergent-divergent nozzles, and the constant-area nozzle lead to higher total pressure losses, compared with the convergent-divergent nozzles. The assistant jets from the annular nozzle affect the coaxial jet flows within the distance less than about ten times the nozzle throat diameter, but beyond it the coaxial jet is conical with self-similar velocity profiles. Increasing both the primary jet pressure ratio and the assistant jet pressure ratio produces a longer coaxial jet core.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.1
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• pp.18-23
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• 2006

Flutter analysis of NACA 0012 with Gurney flap was conducted in time domain. Flutter analysis was performed with a conjunction of two governing equations; one is 2D Navier-Stokes equation and, the other is Lagrange equation of two dimensional plunge & pitch model. Both governing equations were coupled by loose-coupling method. From the computed results, the effect of Gurney flap was concluded to move the flutter boundary of NACA 0012 downward, which means flutter occurs at lower speed than that of NACA 0012. Although flutter boundary of gurney flap was above the safety margin when mach number was lower than 0.85, there might be a possibility of crossing the safety margin when mach number was between 0.85 and 0.9. For safety, the effect of gurney flap needs to be investigated carefully before using it.

• Korean Journal of Optics and Photonics
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• v.6 no.4
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• pp.288-295
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• 1995

We have studied about spatial wavefront analysis of a pulsed laser output beam by means of an imaging unstable ring resonator with an infinite Fresnel number. We have constructed a Newtonian telescope type NBUR (Negative Branch Unstable Ring) Nd:YAG laser with 4 plane mirrors. Provided that the NBUR resonator is under self imaging condition, the resonator could be reduced diffraction effects which were occuring to the beam transmitting through the circular aperture of scraper mirror in the resonator. We have observed enhancement in the spatial coherence of wavefronts of laser beam due to the iterative round trips of the self imaging beam inside the ring resonator. The information on the spatial wavefronts was determined by the fringe patterns from Mach-Zehnder interferometer and the fringe analysis by means of the Fourier transform method showed the distortion of wavefronts of less than 0.2), In comparison with a standing wave type resonator, we have confirmed that the spatial coherence of the NBUR Nd:YAG laser output beam was enhanced as much as 75%. s 75%.