• International Journal of Aeronautical and Space Sciences
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• 제13권3호
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• pp.317-322
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• 2012

The low speed aerodynamic characteristics for a modified sonic arc airfoil which is designed by using the nose shape function of sonic arc, the shape function of NACA four-digit wing sections, and Maple are experimentally investigated. The small rotor blades of a modified sonic arc and NACA0012 airfoil are precisely fabricated with a commercially available light aluminum(Al 6061-T6) and are spin tested over a low speed range (3000rpm-5000rpm). In a consuming power comparison, the consuming powers of NACA0012 are higher than that of modified sonic arcs at each pitch angle. The measured rotor thrust for each pitch angle is used to estimate the rotor thrust coefficient according to momentum theory in the hover state. The value of thrust coefficients for both two airfoils at each pitch angle show almost constant values over the low Mach number range. However, the rotor thrust coefficient of NACA0012 is higher than that of the modified sonic arc at each pitch angle. In conclusion, the aerodynamic performance of NACA0012 is better than that of modified sonic arcs in the low speed regime. This test model will provide a convenient platform for improving the aerodynamic performance of small scale airfoils and for performing design optimization studies.

• 한국항공우주학회지
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• 제31권3호
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• pp.1-7
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• 2003

본 연구에서는 저 레이놀즈 수 유동장에서 유연 익형의 공탄성적 거동과 공기 역학적 성능이 평가되었다. 유연 익형은 비정상 유동장에서 저 레이놀즈 수 익형으로 흔히 사용되는 CLARK-Y 익형 윗면의 일정부분에 질량이 없는 박막을 장착하여 모델링 하였다. 박막의 거동은 공기역학적 힘과 박막의 평형 방정식에 의해 지배되며 평형 방정식의 무차원화로부터 유동과 박막간의 상호작용을 나타내는 무차원 변수가 도출되며 이 무처원 변수가 박막의 거동에 큰 영향을 미친다. 박막의 분포를 익형 윗면의 지정된 지점에서부터 뒷전까지 분포시키되 지정된 박막 분포의 시작점을 변화시켜가며 각 박막 분포에서 박막의 공탄성적 거동을 지배하는 무차원 변수에 대해 공기역학적 성능의 최적화를 수행하였다. 그 결과 박막 분포의 시작점이 뒷전으로 이동할수록 무차원 변수는 거의 선형적으로 증가해야함을 알 수 있었다.

• Smart Structures and Systems
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• 제16권3호
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• pp.555-577
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• 2015

In the work at hand, the development of a morphing flap, actuated through shape memory alloy load bearing elements, is described. Moving from aerodynamic specifications, prescribing the morphed shape enhancing the aerodynamic efficiency of the flap, a suitable actuation architecture was identified, able to affect the curvature. Each rib of the flap was split into three elastic elements, namely "cells", connected each others in serial way and providing the bending stiffness to the structure. The edges of each cell are linked to SMA elements, whose contraction induces rotation onto the cell itself with an increase of the local curvature of the flap airfoil. The cells are made of two metallic plates crossing each others to form a characteristic "X" configuration; a good flexibility and an acceptable stress concentration level was obtained non connecting the plates onto the crossing zone. After identifying the main design parameters of the structure (i.e. plates relative angle, thickness and depth, SMA length, cross section and connections to the cell) an optimization was performed, with the scope of enhancing the achievable rotation of the cell, its ability in absorbing the external aerodynamic loads and, at the same time, containing the stress level and the weight. The conceptual scheme of the architecture was then reinterpreted in view of a practical realization of the prototype. Implementation issues (SMA - cells connection and cells relative rotation to compensate the impressed inflection assuring the SMA pre-load) were considered. Through a detailed FE model the prototype morphing performance were investigated in presence of the most severe load conditions.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2001년도 춘계학술대회논문집C
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• pp.380-385
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• 2001

The paper presents a new multidisciplinary design optimization architecture using optimal sensitivity and coordination of interdisciplinary design variables. Original design problem is decomposed into a number of sub-problems that represent individual engineering analysis. The coupled effects between sub-problems are computed by interdisciplinary design variables. System level coordination is determined by optimal parameter sensitivity calculated by finite difference method. The proposed. MDO strategy is applied to a simplified model of rotorcraft blade design associated with structures and aerodynamic disciplines.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2008년도 춘계학술대회논문집
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• pp.413-414
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• 2008

Abstract should be in English. The leaning angle optimization of turbine blade using the genetic algorithm was conducted in this paper. The calculation CFD technique was based upon the Diagonalized Alternating Directional Implicit scheme(DADI) with algebraic turbulencemodeling. The leaning angle of VKI turbine blade was represented using B-spline curve. The control points are the design variable. Genetic algorithm was taken into account as an optimization tool. The objective was to minimize the total pressure loss. The optimized final geometry shows the better aerodynamic performance compared with the initial turbine blade.

• 한국유체기계학회 논문집
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• 제19권2호
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• pp.16-20
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• 2016

Design optimization of a transonic compressor rotor(NASA rotor 37) was carried out using response surface method(RSM) which is one of the optimization methods. A numerical simulation was conducted using ANSYS CFX by solving three-dimensional Reynolds-averaged Navier Stokes(RANS) equations. Response surfaces that were based on the results of the design of experiment(DOE) techniques were used to find an optimal shape of blade which has the maximum aerodynamic performance. Two objective functions, viz., the adiabatic efficiency and the loss coefficient were selected with three design configurations to optimize the blade shape. As a result, the efficiency of the optimized blade is found to be increased.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2007년도 제28회 춘계학술대회논문집
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• pp.38-42
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• 2007

본 연구에서는, 다분야 통합 해석을 이용한 순항미사일 형상 최적설계를 수행하였다. FORTRAN을 이용하여 개발한 공력, 중량, 성능 및 임무 해석 모듈을 프레임워크를 통하여 통합하였으며 최적화를 위하여 전역최적화 도구인 다윈 알고리즘을 사용하였다. 최적설계 결과, 다른 설계 구속조건을 만족시키면서 17% 가량 전체 무게를 줄일 수 있었다.

• 한국추진공학회지
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• 제16권1호
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• pp.25-35
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• 2012

본 논문에서는 가스터빈 유로의 열유동 환경을 개선하기 위해 끝벽면 경계층 판을 이용한 방법을 적용하였으며, 이 방법에 의한 효과를 최대화하기 위해 경계층 판의 형상에 대한 최적화를 수행하였다. 터빈 유로를 모사하기 위해 $90^{\circ}$ 곡관을 이용하였다. 터빈 유로에서의 전압력 손실과 유로 끝벽면에서의 열전달 계수, 그리고 끝벽면 면적을 최소화하는 �微蛙� 판 형상 도출을 연구의 목적으로 하였으며, 최적화 과정의 효율성을 위해 근사 최적화 기법을 적용하였다. 연구결과를 통해, 최적화된 경계층 판에 의한 상당한 공력환경 개선을 확인할 수 있었으며, 열환경 개선 정도는 공력환경 개선정도에 비해 작게 나타났다.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2000년도 춘계 학술대회논문집
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• pp.60-65
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• 2000

An aerodynamic design method has been developed by using a three-dimensional unstructured Euler code and an adjoint code with a discrete approach. The resulting adjoint code is applied to a wing design problem of super-sonic transport with a wing-body-nacelle configuration. Hicks-Henne shape functions are adopted far the surface geometry perturbation, and the elliptic equation method is employed fer the interior grid modification during the design process. Interior grid sensitivities are neglected except those for design parameters associated with nacelle translation. The Sequential Quadratic Programming method is used to minimize the drag with constraints on the lift and airfoil thickness. Successful design results confirm validity and efficiency of the present design method.

• International Journal of Aeronautical and Space Sciences
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• 제8권1호
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• pp.21-31
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• 2007

Relaxed Static Stability (RSS) concept has been applied to improve aerodynamic performance of modern version supersonic jet fighter aircraft. The T-50 advanced supersonic trainer employs the RSS concept in order to improve the aerodynamic performance. And the flight control system stabilizes the unstable aircraft and provides adequate handling qualities. The T-50 longitudinal control laws employ a proportional-plus-integral type controller based on a dynamic inversion method. The longitudinal dynamic modes consist of short period with high frequency and phugoid mode with low frequency. The design goal of longitudinal control law is optimization of short period damping ratio and frequency using Lower Order Equivalent System (LOES) complying the requirement of MIL-F-8785C. This paper addresses phugoid mode characteristics such as damping ratio and natural frequency that is affected by the nonlinear control laws such as angle of attack limiter, auto pitch attitude command system and autopilot of pitch attitude hold.