• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.2
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• pp.119-125
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• 2020

A high Reynolds number transonic wind tunnel has been built in 2018 at Agency for Defense Development(ADD). The tunnel has a closed circuit with a 1.5m×1.5m test section and is injection driven from a 140bar air supply system. The Mach number range is 0.3-1.2 with a conventional contracting nozzle and 1.4 with a convergent-divergent contraction. The stagnation pressure range is 100-550kPa at the lowest Mach number. An AGARD-B standard model is tested in the transonic wind tunnel to obtain 6-DOF aerodynamic coefficients. The results are compared with those obtained from ADD trisonic wind tunnel and others. We verify that the transonic wind tunnel become available to develop an aircraft from the testing results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.2
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• pp.11-16
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• 2006

An experimental study of the transonic flows over NACA and double wedge airfoils was conducted with a shock tube. The configuration of test section with a slotted wall and chamber was designed and tested to minimize wall and reflected shock wave effects and use the shock tube as simple and less costly wind tunnel generating the relatively high Reynolds numbers transonic flow. Transonic airfoil flows at hot gas Mach numbers of 0.80~0.84, Reynolds number of about $1.2{\times}10^6$ on airfoil chord length and angles of attack of $0^{\circ}$ and $2^{\circ}$ were visualized with the shadowgraph method. The shock wave profiles on the airfoils were compared with the corresponding results from the conventional transonic wind tunnel tests. The experimental results showed that present shock tube exhibited the proper performance characteristics as transonic wind tunnel for tested Mach number range and airfoils.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.10-10
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• 1998

천음속 익형에서 발생하는 응축충격파와 경계층의 간섭을 피동제어 하여 항력감소에 대한 연구를 2.5$\times$7$\textrm{cm}^2$ 천음속 풍동에서 수행하였다. 익형표면에 설치한 정압공으로 정압을, 익형후방에 설치한 8개의 Pitot probe로 전압을 동시에 측정하여 충격파를 통한 에너지의 손실과 항력의 변화를 계산하였고, 또한 유동장과 충격파의 형상을 가시화하기 위해 슈리렌 가시화 시스템을 사용하였다. 실험은 NACA 0012 익형에서 기공률 변화에 따른 피동제어의 항력감소 초과를 조사한 다음 NACA 64-018 익형에서는 기공률과 공동의 크기의 변화가 미치는 효과를 연구하였다. 피동제어의 개념은 충격파가 발생하는 하부벽을 다공벽으로 만들고 그 아래를 공동으로 만들면 충격파 후방의 상대적으로 높은 압력이 기류의 일부를 공동으로 자연스럽게 유입시키고 다시 공동에서 낮은 압력의 충격파 상류로 유출시키는 것이다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.10
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• pp.894-900
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• 2012

The purpose of this research is to investigate the dynamic roll-damping characteristics of a spin-stabilized projectile in wind-tunnel testing. In the present work, the high-speed wind-tunnel tests for the roll-damping measurements were conducted on a finned spin-stabilized projectile model in the Agency for Defense Development's Trisonic Wind Tunnel at spin rates about 8,000 rpm. The test Mach numbers ranged from 0.6 to 0.9, and the angles of attack ranged from 0 to +15 deg. The evaluation of the bearing friction parameter was also conducted to eliminate the tare damping moment from the aerodynamic damping moment.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.10
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• pp.912-918
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• 2011

The purpose of this research is to determine the dynamic roll-damping data of a spinning projectile in wind-tunnel testing. In the present work, the high-speed wind-tunnel tests for the roll-damping measurements were conducted on a spin-stabilized projectile model in the Agency for Defense Development's Tri-Sonic Wind Tunnel at spin rates about 12,000 rpm. The test Mach numbers ranged from 0.7 to 1.05, and the angles of attack ranged from -4 to +10 deg. The validity of the wind-tunnel measurement techniques was evaluated by comparing them with the previous test results on the same configuration.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.3
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• pp.252-258
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• 2017

This paper describes on introduction of CASE 3 for EFD-CFD comparison workshop which is incharged of aerodynamic subcommittee of The Korean Societry or Aeronautical and Space Science. In addition, the results of candidate for CASE 3 were compared with experimental result and computational result. Currently, for this case 3, there are eight candidates from company, university and research institute. According to comparison of their results, they are in accordance with experimental data and computational data.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.8-15
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• 2001

In the continuing quest for increased turbomachinery efficiency, the part played by blade profile shape remains crucial. The application of a heated thin metallic film with CTA(constant temperature anemometer) to the measurements of the laminar and turbulent boundary layer behavior(shock-boundary layer-interaction) in a transonic wind tunnel. Results of measurements with hot-film sensors on transonic compressor blades are extremely difficult to interpret because of ambiguous probe signals due to the complexity of the local flow pattern. In order to get the explicit information and give the designer to interpret characteristic signals from hot-film probes, a method was developed by comparing the results with other measuring technic results.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.1
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• pp.41-52
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• 2021

During atmospheric ascent of a launch vehicle, airborne acoustic loads act on the vehicle and its effect becomes pronounced at transonic speed. In the present study, acoustic loads acting on a hammerhead launch vehicle at a transonic speed have been analyzed using ��-ω SST based IDDES and the results including mean Cp, Cprms, and PSD are compared to available wind-tunnel test data. Mesh dependency of IDDES results has been investigated and it has been concluded that with an appropriate turbulence scale-resolving computational mesh, the characteristic flow features around a transonic hammerhead launch vehicle such as separated shear flow at fairing shoulder and its reattachment on rear body as well as large pressure fluctuation in the region of separated flow behind the boat-tail can be predicted with reasonable accuracy for engineering purposes.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.46 no.6
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• pp.445-451
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• 2018

This paper describes on introduction of CASE 4(Delta Wing) for EFD-CFD comparison workshop which is in charged of aerodynamic subcommittee of The Korean Society for Aeronautical and Space Science. The wind tunnel test will be performed later, angle of attack is set -5~20deg and mach number is set 0.7, 0.85, 1.2 to solve the transonic flow. The simulation test of grid dependency is conducted to determine the proper grid size of delta wing analysis. The tendency of lift and drag coefficient is determined according to the change of angle of attack based on the selected grid size.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.9
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• pp.27-33
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• 2005

This research presents the influence of the base region modeling on the aerodynamic characteristics of a launch vehicle using CFD. The vicinity of a launch vehicle is divided into four zones, and four computational cases are made using these four zones. The aerodynamic coefficients are predicted for the angle-of-attack of 6 degrees and Mach numbers ranging from 0.4 to 2.86. It was found that modeling of the base region should not be neglected for the prediction of the aerodynamic characteristics of a launch vehicle in subsonic and transonic regions. It was also found that the modeling of the sting support used in the wind tunnel test is necessary to get a better agreement with the experiments.