• 항공우주시스템공학회지
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• 제15권5호
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• pp.60-71
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• 2021

상용 프로펠러의 추력과 토크를 측정하기 위한 성능시험 장치를 고안하였으며, 30인치급 3가지 프로펠러에 대한 성능시험을 수행하였다. 프로펠러 추력과 토크 측정을 위해 모터, 프로펠러와 연결된 6분력 발란스를 적용하였으며, 풍동 저울 교정 장비를 이용하여 측정 시스템의 확인을 수행하였으며, QTP 프로펠러를 적용하여 구축한 성능시험 장치의 검증 시험도 수행하였다. 제자리 비행 조건에서 상용 프로펠러의 제작사에서 제공하는 사양과 시험 결과를 비교하였으며, 추력 및 토크에서 차이가 있음을 확인하였다. 받음각, 프로펠러 형상, 풍속을 변경시켜가며 프로펠러 성능을 측정하였으며, 각 프로펠러에 대해 RPM에 따른 프로펠러 추력 계수로 나타내었다. 저 받음각과 고 받음각에서 경향이 다르게 나타남을 확인하였으며, 차후 공중용 무인이동체 공력 설계에 활용 가능한 공력 데이터베이스를 확보하였다.

• 대한조선학회논문집
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• 제48권2호
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• pp.113-120
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• 2011

Fins are widely used for roll stabilization of passenger ferries and high performance naval ships, among others. The Coanda effect is noticeable when a jet stream is applied tangentially to a curved wing surface since the jet can augment the lift by increasing the circulation. The Coanda effect has been found useful in various fields of aerodynamics and speculated to have practical applicability in marine hydrodynamics where various control surfaces are used to control motions of ships and the other offshore structures. In the present study, numerical computations have been performed to find proper jet momentum coefficients $C_j$ and trailing edge shapes suitable for the application of the Coanda effect to a stabilizer fin. The results show that the lift coefficient of the modified Coanda fin at the zero angle of attack ${\alpha}$ identically coincides with that of the original fin at ${\alpha}\;=\;25^{\circ}$ when Coanda jet is supplied at the rate of $C_j$ = 0.1. It is also shown that a fixed type fin stabilizer utilizing the Coanda effect can be implemented without changing the fin angle to actively control the motions of ships and the other offshore structures.

• Advances in aircraft and spacecraft science
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• 제2권1호
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• pp.45-56
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• 2015

The attitude aerodynamic control is an important subject in the design of an aerospace plane. Usually, at high altitudes, this control is fulfilled by thrusters so that the implementation of an aerodynamic control of the vehicle has the advantage of reducing the amount of thrusters fuel to be loaded on board. In the present paper, the efficiency of a wing-flap has been evaluated considering a NACA 0010 airfoil with a trailing edge flap of length equal to 35% of the chord. Computational tests have been carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km, in the range of angle of attack 0-40 deg. and with flap deflection equal to 0, 15 and 30 deg.. Effects of the flap deflection have been quantified by the variations of the aerodynamic force and of the longitudinal moment. The shock wave-boundary layer interaction and the shock wave-shock wave interaction have been also considered. A possible interaction of the leading edge shock wave and of the shock wave arising from the vertex of the convex corner, produced on the lower surface of the airfoil when the flap is deflected, generates a shock wave whose intensity is stronger than those of the two interacting shock waves. This produces a consistent increment of pressure and heat flux on the lower surface of the flap, where a thermal protection system is required.

• Advances in aircraft and spacecraft science
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• 제4권5호
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• pp.503-514
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• 2017

The attitude control of an aircraft is usually fulfilled by means of thrusters at high altitudes. Therefore, the possibility of using also aerodynamic surfaces would produce the advantage of reducing the amount of fuel for the thrusters to be loaded on board. For this purpose, Zuppardi already considered some aerodynamic problems linked to the use of a wing flap in a previous paper. A NACA 0010 airfoil with a trailing edge flap of 35% of the chord, in the range of angle of attack 0-40 deg and flap deflections up to 30 deg was investigated. Computer tests were carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km of Earth Atmosphere. The present work continues this subject, considering the same airfoil and free stream conditions but two flap extensions of 45% and 25% of the chord and two flap deflections of 15 and 30 deg. The main purpose is to compare the influence of the flap dimension with that of the flap deflection. The present analysis is carried out in terms of: 1) percentage variation of the global aerodynamic coefficients with respect to the no-flap configuration, 2) increment of pressure and heat flux on the airfoil lower surface due to the Shock Wave-Shock Wave Interaction (SWSWI) with respect to the same quantities with no SWSWI or in no-flap configuration, 3) flap hinge moment. Issues 2) and 3) are important for the design of the mechanical and thermal protection system and of the flap actuator, respectively. Under the above mentioned test and geometrical conditions, the flap deflection is aerodynamically more effective than the flap extension, because it involves higher variation of the aerodynamic coefficients. However, tests verify that a smaller deflection angle involves the advantage of a smaller increment of pressure and heat flux on the airfoil lower surface, due to SWSWI, as well as a smaller hinge moment.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2004년도 춘계학술대회
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• pp.1672-1677
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• 2004

Leading edge extension(LEX) in a highly swept shape applied to a delta wing features the modem air-fighters. The LEX vortices generated upon the upper surface of the wing at high angle of attack enhance the lift force of the delta wing by way of increased negative suction pressure over the surfaces. The present 3-D stereo PIV includes the Identification of 2-D cross-correlation equation, stereo matching of 2-D velocity vectors of two cameras, accurate calculation of 3-D velocity vectors by homogeneous coordinate system, removal of error vectors by a statistical method followed by a continuity equation criterion and so on. A delta wing model with or without LEX was immersed in a circulating water channel. Two high-resolution, high-speed digital cameras($1280pixel{\times}1024pixel$) were used to allow the time-resolved animation work. The present dynamic stereo PIV represents the complicated vortex behavior, especially, in terms of time-dependent characteristics of the vortices at given measuring sections. Quantities such as three velocity vector components, vorticity and other flow information can be easily visualized via the 3D time-resolved post-processing to make the easy understanding of the LEX effect or vortex emerging and collapse which are important phenomena occurring in the field of delta wing aerodynamics.

• 한국가시화정보학회:학술대회논문집
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• 한국가시화정보학회 2003년도 추계학술대회 논문집
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• pp.39-42
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• 2003

Leading edge extension(LEX) in a highly swept shape applied to a delta wing features the modern air-fighters. The LEX vortices generated upon the upper surface of the wing at high angle of attack enhance the lift force of the delta wing by way of increased negative suction pressure over the surfaces. The present 3-D stereo PIV includes the Identification of 2-D cross-correlation equation, stereo matching of 2-D velocity vectors of two cameras, accurate calculation of 3-D velocity vectors by homogeneous coordinate system, removal of error vectors by a statistical method followed by a continuity equation criterion and so on. A delta wing model with or without LEX was immersed in a circulating water channel. Two high-resolution, high-speed digital cameras$(1280pixel\times1024pixel)$ were used to allow the time-resolved animation work. The present dynamic stereo PIV represents the complicated vortex behavior, especially, in terms of time-dependent characteristics of the vortices at given measuring sections. Quantities such as three velocity vector components, vorticity and other flow information can be easily visualized via the 3D time-resolved post-processing to make the easy understanding of the LEX effect or vortex emerging and collapse which are important phenomena occurring in the field of delta wing aerodynamics.

• Wind and Structures
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• 제35권3호
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• pp.177-191
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• 2022

Trains emerging on a streamlined bridge-girder may have salient interference effects on the aerodynamic properties of the bridge. The present paper aims at investigating these interferences by wind tunnel measurements, covering surface pressure distributions, near wake profiles, and flow visualizations. Experimental results show that the above interferences can be categorized into two primary effects, i.e., an additional angle of attack (AoA) and an enhancement in flow separation. The additional AoA effect is demonstrated by the upward-moved stagnation point of the oncoming flow, the up-shifted global symmetrical axis of flow around the bridge-girder, and the clockwise-deflected orientation of flow approaching the bridge-girder. Due to this additional AoA effect, the two critical AoAs, where flow around the bridge-girder transits from trailing-edge vortex shedding (TEVS) to impinging leading-edge vortices (ILEV) and from ILEV to leading-edge vortex shedding (LEVS) of the bridge-girder are increased by 4° with respect to the same bridge-girder without trains. On the other hand, the underlying flow physics of the enhancement in flow separation is the large-scale vortices shedding from trains instead of TEVS, ILEV, and LEVS governed the upper half bridge-girder without trains in different ranges of AoA. Because of this enhancement, the mean lift and moment force coefficients, all the three fluctuating force coefficients (drag, lift, and moment), and the aerodynamic span-wise correlation of the bridge-girder are more significant than those without trains.

• Wind and Structures
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• 제34권1호
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• pp.127-136
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• 2022

In this work a multi-fidelity non-intrusive polynomial chaos (MF-NIPC) has been applied to a structural wind engineering problem in architectural design for the first time. In architectural design it is important to design structures that are safe in a range of wind directions and speeds. For this reason, the computational models used to design buildings and bridges must account for the uncertainties associated with the interaction between the structure and wind. In order to use the numerical simulations for the design, the numerical models must be validated by experi-mental data, and uncertainties contained in the experiments should also be taken into account. Uncertainty Quantifi-cation has been increasingly used for CFD simulations to consider such uncertainties. Typically, CFD simulations are computationally expensive, motivating the increased interest in multi-fidelity methods due to their ability to lev-erage limited data sets of high-fidelity data with evaluations of more computationally inexpensive models. Previous-ly, the multi-fidelity framework has been applied to CFD simulations for the purposes of optimization, rather than for the statistical assessment of candidate design. In this paper MF-NIPC method is applied to flow around a rectan-gular 5:1 cylinder, which has been thoroughly investigated for architectural design. The purpose of UQ is validation of numerical simulation results with experimental data, therefore the radius of curvature of the rectangular cylinder corners and the angle of attack are considered to be random variables, which are known to contain uncertainties when wind tunnel tests are carried out. Computational Fluid Dynamics (CFD) simulations are solved by a solver that employs the Finite Element Method (FEM) for two turbulence modeling approaches of the incompressible Navier-Stokes equations: Unsteady Reynolds Averaged Navier Stokes (URANS) and the Large Eddy simulation (LES). The results of the uncertainty analysis with CFD are compared to experimental data in terms of time-averaged pressure coefficients and bulk parameters. In addition, the accuracy and efficiency of the multi-fidelity framework is demonstrated through a comparison with the results of the high-fidelity model.