• Proceeding of EDISON Challenge
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• 2016.11a
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• pp.19-22
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• 2016

Leading edge thrust is generally caused by passing air flow from lower to upper surface and it is required to have sufficient angle of attack for notable leading edge thrust. To produce leading edge thrust at low angle of attack, utilizing expansion wave accompanying low pressure is able to be a solution. Fore structure changes the direction of flow, and this flow passes the projected edge. As a result, from a perspective of the edge, it is able to have high angle of attack, and artificial expansion wave is generated. This concept shows 9.48% increase of L/D in inviscid flow, at Mach number 1.3 and angle of attack $1^{\circ}$ in maximum, and this model shows the 3.98% of increasement at angle of attack $2^{\circ}$. Although advantage of the artificial expansion wave decreased as angle of attack increase, it shows the possibility of aerodynamical improvement with artificial expansion wave.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.3
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• pp.425-435
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• 2017

In this work, an overview of flow diagnosis in a shock tunnel is made by means of using established techniques that are easy to setup, economical to arrange, and simple to measure. One flow condition was considered having Mach number of 6 at the nozzle-exit, regarded as freestream. Measured aerothermodynamic data such as shock wave speed, wall static and total pressures, surface heat flux, and shock stand-off distance ahead of test model showed good agreement with calculation. This study shows an overall procedure of flow diagnosis in a shock tunnel in a single manuscript. Outcomes are thought to be useful in the field of education and also in a preliminary stage of high-speed vehicle design and tests, that need to be performed within a short time with decent accuracy.

• Journal of the Korea Institute of Military Science and Technology
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• v.2 no.2
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• pp.70-82
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• 1999

During the aircraft conceptual design stage, rapid aerodynamic analyses over various configurations are required. Hence, empirical and analytical methods play important roles in predicting the aero-dynamic characteristics. In this study, total drag estimation method based on empirical and analytical approaches is developed. By comparing with the results of the wind tunnel experiment and existing prediction methods, it is demonstrated that the developed method is accurate and useful in predicting total drag for the whole Mach number range.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.11a
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• pp.31-31
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• 2000

초음속 충돌제트(impinging jet)의 열 및 운동량 전달(heat and momentum transfer)은 로켓의 이ㆍ착륙, 다단 로켓의 분리, 로켓의 방향조절을 위해 배기 노즐에 부착되는 제트 베인(jet vane)이나 스포일러 탭(spoiler tab), 수직/단거리 이착륙기의 발진, 미사일 발사시스템, 전투기 동체, 날개, 후미 부분에서 발사되는 미사일의 배기가스가 주변장치 등에 충돌할 때 발생되는 문제점 등을 사전 예측하여 관련장비의 설계 둥에 유용한 자료로 이용된다. 따라서 이에 대한 기초 연구로서 초음속 유동 실험장치를 이용하여 마하수(Mach Number) 1.0 및 1.8인 경우에 대하여 수직/경사평판에서 팽창 비, 거리, 경사각에 따른 충돌 면에서의 단열 벽면온도를 측정하였다. (중략)

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.341-344
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• 2008

Numerical simulations have been carried out for a supersonic two-dimensional flow over open, rectangular cavities (length-to-depth ratios are L/D = 1.0) in order to investigate the effect of non-equilibrium condensation of moist air on supersonic flows around the cavity for the flow Mach number 1.83 at the cavity entrance. In the present computational investigation, a condensing flow was produced by an expansion of moist air in a Laval nozzle. The results obtained showed that in the case with non-equilibrium condensation for L/D = 1.0, amplitudes of oscillation in the cavity became smaller than those without the non-equilibrium condensation. Furthermore, the occurrence of the non-equilibrium condensation reduced the peaks of power spectrum density and the frequency of the flow field oscillation increased in comparison with the case of $S_0$ = 0.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.345-348
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• 2008

The subcavity passive control technique is used in present study. Cavity-induced pressure oscillation has been investigated numerically for a supersonic three-dimensional flow over rectangular cavities at Mach number 1.83 at the cavity entrance. The three-dimensional, compressible Navier-stokes equations are numerically solved based on a fully implicit finite volume scheme. The results showed that the resultant amount of attenuation of cavity-induced pressure oscillations was dependent on the length and thickness of the flat plate.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.353-356
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• 2008

The effects of design parameters of supersonic ejector system under the assumption of constant pressure mixing; such as mass flow rate ratio, area ratio, Primary mach number on ejector system performance were investigated by theoretical formulations. And for a given design condition and working fluid, Computational Fluid Dynamics was conducted.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.193-195
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• 2011

An abrupt increase of duct cross-section is frequently encountered in pressure reducing devices, valves of internal combustion engines and in gas pipelines. Supersonic flow in a rectangular duct passing an abrupt increase of cross-section is studied numerically. The behavior of base pressure of the dead-air region at sudden enlargement of the duct is clarified. This investigation concerns the determination of the base pressure, which is independent of the size of the enlarged part. Several flow patterns are identified with different enlargements according to the ratio between the downstream ambient pressure and the upstream reservoir pressure. Base pressure and the resulting shock-structure are highly depending on the size of duct enlargement. For a given duct, base pressure tends to minimum for a particular pressure ratio. In addition, the locations of secondary separation and reattachment points of the jet plume are found with respect to different duct enlargements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.132-135
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• 2011

A technology for buzz margin determination is suggested to obtain stable shock structure and high compression efficiency of supersonic intake. By using the shock equilibrium equation of supersonic intake, sensitivity equation of terminal shock position for free stream and back pressure is induced and disturbances are quantified through statistical approach. Numerical results show that the sensitivity of shock position for disturbances is proportional to Mach number and the back pressure is dominant for variance of terminal shock position.

• Structural Engineering and Mechanics
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• v.20 no.4
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• pp.435-450
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• 2005

Here, the dynamic instability characteristics of aero-thermo-mechanically stressed functionally graded plates are investigated using finite element procedure. Temperature field is assumed to be a uniform distribution over the plate surface and varied in thickness direction only. Material properties are assumed to be temperature dependent and graded in the thickness direction according to simple power law distribution. For the numerical illustrations, silicon nitride/stainless steel is considered as functionally graded material. The aerodynamic pressure is evaluated based on first-order high Mach number approximation to the linear potential flow theory. The boundaries of the instability region are obtained using the principle of Bolotin's method and are conveniently represented in the non-dimensional excitation frequency-load amplitude plane. The variation dynamic instability width is highlighted considering various parameters such as gradient index, temperature, aerodynamic and mechanical loads, thickness and aspect ratios, and boundary condition.