• 한국항공우주학회지
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• 제33권9호
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• pp.27-33
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• 2005

전산유동해석을 이용하여 발사체 공력특성을 예측함에 있어서 발사체 기저부 영역 모델링에 따라 그 결과가 어떻게 달라지는가에 대하여 알아보았다. 기저부 영역 모델링 특성을 보기 위해 발사체 주변을 네 개의 영역으로 구분하고 이를 네 가지로 서로 조합하여 받음각 $6^{\circ}$에 대해 마하수를 0.4부터 2.86까지 변화시켜 가며 공력 계수를 산출하였다. 먼저 발사체 기저부 영역이 계산결과에 미치는 영향을 살펴보았을 때, 아음속 및 천음속 영역 해석 시, 기저부 영역 모델링이 반드시 필요함을 확인했다. 다음으로 풍동시험에 사용한 스팅의 영향을 살펴보았을 때, 스팅 형상을 고려하여 계산/보정한 결과가 전반적으로 풍동시험에 가장 근접했다.

• 한국항공우주학회지
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• 제36권6호
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• pp.532-540
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• 2008

최근 구조변형을 고려한 항공기의 공력특성을 계산하는데 필요한 CFD와 CSD 기법이 연계된 FSI 시스템에 관한 관심이 증대하고 있다. 본 연구에서는 유체유발 구조 변형을 고려한 수렴된 구조형상에 대한 공력특성 예측을 위해 유체-구조 연계 시스템인 FSI(Fluid- Structure Interaction)를 구축하였다. 각 모듈의 연계, 특히 CSD와 CFD의 결합 및 변형된 형상에 대한 공력격자 재생성을 위해 VSI(Volume Spline Interpolation)와 격자 변형 코드를 개발하였으며, 공력과 구조의 해석 모듈로 상용 프로그램인 FLUENT와 NASTRAN을 사용하였다. 구축된 시스템을 DLR-F4 날개에 적용하여 정적 유체-구조 연구를 수행하였으며, 그 결과 마하수 0.75에서 변형된 형상에 대한 양력 및 항력 계수는 약 20.26%, 18.5% 감소하는 것으로 나타났다.

• 한국항공우주학회지
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• 제46권11호
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• pp.883-891
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• 2018

본 연구에서는 항공기 랜딩 기어와 유사한 형태의 탠덤 실린더에 대해 DES (Detached Eddy Simulation)와 FW-H (Ffowcs Williams and Hawkings) 음향상사법을 이용한 전산공력 소음해석을 수행하였다. 실린더 중심 간 거리 3.7D 조건에 대한 전산공력소음해석 결과를 탠덤 실린더를 지나는 외부 유동 특성 및 원거리 소음을 계측한 선행 연구 결과와 비교하였다. 전방 실린더의 상/하부에서 발생한 와류(Vortex)가 특징적인 주파수를 갖고 후방 실린더에 충돌하면서 소음 발생원이 되는 것을 확인하였다. DES 와 FW-H 음향상사법을 이용한 전산공력소음해석 방법의 타당성을 확인한 후, 전방 실린더의 후방에 분할판을 적용한 경우의 전산공력소음해석을 수행하였다. 전방 실린더의 후방에 분할판이 적용되면 전방 실린더의 상/하부에서 발생하는 와류의 특징을 변화시키고 이극자 소음원 크기를 감소시켜 특정 주파수 대역에서 발생하는 소음 크기가 감소하는 것을 확인하였다.

• 항공우주시스템공학회지
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• 제16권4호
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• pp.17-27
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• 2022

본 논문에서는 에어포일의 좌표 데이터에 대해 공력 특성을 예측할 수 있는 합성곱 신경망 기반 네트워크 프레임 워크를 설계하였으며 Xfoil을 이용한 공력 데이터를 적용하여 네트워크의 가능성을 확인하였다. 이 때 에어포일의 두께 변화에 따른 공력 특성 예측을 수행하였다. 부호화 거리 함수를 이용하여 에어포일의 좌표 데이터를 이미지 데이터로 변환하였으며 받음각 정보를 반영하였다. 또한 에어포일의 압력 계수 분포를 축소 모델 기법 중 하나인 적합 직교 분해를 이용하여 축소된 데이터로 표현하였으며 이를 네트워크의 출력 데이터로 사용하였다. 제시하는 네트워크의 내삽과 외삽 성능을 평가하기 위하여 시험 데이터를 구성하였고, 결과적으로 내삽 데이터에 대한 예측 성능이 외삽에 비해 우수함을 확인하였다.

• International Journal of Aeronautical and Space Sciences
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• 제16권2호
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• pp.148-156
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• 2015

A continuous type side jet controller which has four nozzles with thrust control devices was considered. It is deployed to a missile for high maneuverability and fast controllability in the terminal guidance phase. However, it causes more complex aerodynamic jet interactions between the side jet and the supersonic free stream than does the conventional impulse type side jet with a small single thruster. In this paper, a numerical investigation of the jet interference effects for the missile equipped with a continuous type side jet thruster is presented. A three-dimensional flow field was simulated by using a commercial unstructured-based CFD solver. The numerical simulation method was validated through comparison with wind tunnel test results for the single jet. The method of defining jet direction for this type of side jet control to minimize simulation cases was also introduced. Flow fields investigation and jet interaction effects for various flow conditions, jet pressure ratios and defined jet direction conditions were performed. From the numerical simulation for the continuous type side jet, extensive aerodynamic interference data were obtained to construct an aerodynamic coefficients database for precise missile control.

• 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.

• 제어로봇시스템학회논문지
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• 제6권8호
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• pp.727-731
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• 2000

This paper presents a new practical autopilot design approach to acceleration control for tail-controlled STT(Skid-to-Turn) missiles. The approach is novel in that the proposed parametric affine missile model adopts acceleration as th controlled output and considers the couplings between the forces as well as the moments and control fin deflections. The aerodynamic coefficients in the proposed model are expressed in a closed form with fittable parameters over the whole operating range. The parameters are fitted from aerodynamic coefficient look-up tables by the function approximation technique which is based on the combination of local parametric models through curve fitting using the corresponding influence functions. In this paper in order to employ the results of parametric affine modeling in the autopilot controller design we derived a parametric affine missile model and designed a feedback linearizing controller for the obtained model. Stability analysis for the overall closed loop sys-tem is provided considering the uncertainties arising from approximation errors. the validity of the proposed modeling and control approach is demonstrated through simulations for an STT missile.

• 한국군사과학기술학회지
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• 제12권6호
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• pp.808-814
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• 2009

Numerical flow-filed computations have been conducted around a projectile with a triangular ballute at Mach numbers of 0.3 ~ 0.9 and angles of attack of $0^{\circ}\;{\sim}\;30^{\circ}$. The expansion shape of a triangular ballute has been determined from both the manufacturing dimensions and the wind tunnel test results. It has been assumed that the shape does not alter after the ballute has been expanded completely. The computed results showed a good agreement with the wind tunnel test results. A sensitivity analysis on the aerodynamic coefficients has been performed to evaluate the quantitative effects on the flight performance. Trajectory simulation results were also in good agreement with the flight test results.

• Wind and Structures
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• 제26권5호
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• pp.305-315
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• 2018

A numerical study based on a delayed detached eddy simulation (DDES) is conducted to investigate the aerodynamic mechanism behind the suppression of vortex-induced vibrations (VIVs) of twin box girders by central grids, which have an inhibition effect on VIVs, as evidenced by the results of section model wind tunnel tests. The mean aerodynamic force coefficients with different attack angles are compared with experimental results to validate the numerical method. Next, the flow structures around the deck and the aerodynamic forces on the deck are analyzed to enhance the understanding of the occurrence of VIVs and the suppression of VIVs by the application of central grids. The results show that shear layers are separated from the upper railings and lower overhaul track of the upstream girder and induce large-scale vortices in the gap that cause periodical lift forces of large amplitude acting on the downstream girder, resulting in VIVs of the bridge deck. However, the VIVs are apparently suppressed by the central grids because the vortices in the central gap are reduced into smaller vortices and become weaker, causing slightly fluctuating lift forces on the deck. In addition, the mean lift force on the deck is mainly caused by the upstream girder, whereas the fluctuating lift force is mainly caused by the downstream girder.

• International Journal of Aeronautical and Space Sciences
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• 제14권4호
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• pp.310-323
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• 2013

In this paper, the aeroelastic static response of flexible wings with arbitrary cross-section geometry via a coupled CUF-XFLR5 approach is presented. Refined structural one-dimensional (1D) models, with a variable order of expansion for the displacement field, are developed on the basis of the Carrera Unified Formulation (CUF), taking into account cross-sectional deformability. A three-dimensional (3D) Panel Method is employed for the aerodynamic analysis, providing more accuracy with respect to the Vortex Lattice Method (VLM). A straight wing with an airfoil cross-section is modeled as a clamped beam, by means of the finite element method (FEM). Numerical results present the variation of wing aerodynamic parameters, and the equilibrium aeroelastic response is evaluated in terms of displacements and in-plane cross-section deformation. Aeroelastic coupled analyses are based on an iterative procedure, as well as a linear coupling approach for different free stream velocities. A convergent trend of displacements and aerodynamic coefficients is achieved as the structural model accuracy increases. Comparisons with 3D finite element solutions prove that an accurate description of the in-plane cross-section deformation is provided by the proposed 1D CUF model, through a significant reduction in computational cost.