• International Journal of Aeronautical and Space Sciences
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• 제10권1호
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• pp.20-29
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• 2009

There are concerns about the influence of the gust wind caused by helicopters affecting the moving vehicles while hovering over the road during rescue activities. For the understanding of such complicated flow. numerical simulation of a rotor hovering above the ground has been carried out, changing the rotor/ground clearances. The rotor thrust is kept constant. and the rotor control is determined by trim adjustments incorporated into the CFD algorithm. Collective pitch angle and the required power decreases with the rotor/ground clearance which agrees with experience. Changes of the flowfield near the rotor with regard to the rotor height are investigated based on the calculated results.

• 항공우주시스템공학회지
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• 제18권4호
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• pp.18-26
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• 2024

본 논문은 Generic 쿼드 틸트 로터 UAM 항공기 비행 동력학 기반 비선형 수학적 모델링 및 실시간 시뮬레이션 소프트웨어 개발에 관한 연구 결과를 기술한다. 본 연구에서는 NASA의 UAM 임무 형상 및 요구도를 참고하여 Generic 쿼트 틸트 로터 eVTOL UAM 항공기 형상을 설계하고, 공력 데이터베이스 기반 공기역학, 추력 데이터베이스 기반 프롭로터역학, 항공기의 지면반력, 대기환경을 운동모델로 모델링하였다. 또한 회전익 모드, 천이 모드 및 고정익 모드 별 조종방법을 정의, 나셀 틸트각 Corridor 설정 후 실시간 비행 시뮬레이션 소프트웨어에 구현 후 수평비행 트림 해석을 수행하였다. 본 실시간 비행 시뮬레이션 소프트웨어를 통해 쿼드 틸트 로터 UAM 항공기의 트림 해석뿐 아니라 조종성(Handling Qualities) 예측, 동특성을 고려한 나셀 틸트각 스케줄러 최적화와 회전익, 고정익 및 천이 모드 별 비행 제어법칙 설계/검증 및 비행 시뮬레이터 탑재를 통한 비행 훈련 등 쿼트 틸트 로터 UAM 분야에서 다양한 목적으로 활용 가능할 것으로 예상한다.

• 한국항공우주학회지
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• 제44권11호
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• pp.989-996
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• 2016

본 연구는 비선형 응답모델과 실수기반의 혼합형 유전자 알고리즘을 적용하여 로터의 트?-발란스(RTB) 기법을 개발하는 데 목적이 있다. 트?-발란스 조절 파라미터의 변화에 따른 트림해석 결과를 이용하여 2차의 근사함수를 이용하는 비선형 응답모델을 개발하였다. 트?편차와 기체의 진동응답을 최소화하기 위해 균형추 무게, 트림 탭(Trim Tab) 및 피치링크 길이를 최적화하기 위한 비선형계획 문제를 정식화하였다. 정식화 결과는 수렴성 향상을 위해 군집최적화 기법을 실변수기반의 유전자 알고리즘에 통합한 혼합형 유전자 기법을 사용함으로써 효율적인 해석이 가능하였다. 비선형 모델을 이용한 본 연구의 방법을 선형모델의 결과와 비교하여 본 연구의 방법을 검증하였으며 비선형모델을 사용하는 경우 선형모델의 결과보다 향상된 응답특성을 계산할 수 있음을 밝혔다.

• International Journal of Aeronautical and Space Sciences
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• 제10권1호
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• pp.104-111
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• 2009

An unstructured overset mesh method has been developed for the simulation of unsteady flow fields around isolated rotors and rotor-fuselage configurations. The flow solver was parallelized for the efficient calculation of complicated flows requiring a large number of cells. A quasi-unsteady mesh adaptation technique was adopted to enhance the spatial accuracy of the solution and to better resolve the rotor wake. The method has been applied to calculate the flow fields around rotor-alone and rotor-fuselage configurations in forward flight. Validations were made by comparing the predicted results with those of measurements. It was demonstrated that the present method is efficient and robust for the prediction of unsteady time-accurate flow fields involving multiple bodies in relative motion.

• International Journal of Aeronautical and Space Sciences
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• 제16권1호
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• pp.110-122
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• 2015

In this paper, the unsteady aerodynamics and blade structural dynamics of an experimental bearingless rotor were analyzed. Due to the multiple load path and nonlinear behavior of a bearingless rotor, sophisticated structural modeling and structural-aerodynamic coupled analysis is required. To predict the internal load and deformation of an experimental bearingless rotor, trim analysis was implemented. The results showed good agreement when compared with those predicted by CAMRAD II the rotorcraft comprehensive analysis. It is possible to extend the present structural-aerodynamic combined analysis to further advanced configurations of the bearingless rotor in the future.

• 소음진동
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• 제7권5호
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• pp.819-826
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• 1997

The finite element analyses of a composite hingeless rotor blade in forward flight have been performed to investigate the influence of blade design parameters on the blade stability. The blade structure is represented by a single cell composite box-beam and its nonclassical effects such as transverse shear and torsion-related warping are considered. The nonlinear periodic differential equations of motion are obtained by moderate deflection beam theory and finite element method based on Hamilton principle. Aerodynamic forces are calculated using the quasi-steady strip theiry with compressibility and reverse flow effects. The coupling effects between the rotor blade and the fuselage are included in a free flight propulsive trim analysis. Damping values are calculated by using the Floquet transition matrix theory from the linearized equations perturbed at equilibrium position of the blade. The aeroelastic results were compared with an alternative analytic approch, and they showed good correlation with each other. Some parametric investigations for the helicopter design variables, such as pretwist and precone angles are carried out to know the aeroelastic behavior of the rotor.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2003년도 추계 학술대회논문집
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• pp.95-101
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• 2003

A three dimensional inviscid parallel flow solver has been developed for the simulation of rotor blades in forward flight. The computational domain is divided into stationary and rotating zones for the more efficient mesh adaptation. The conservative mesh treatment algorithm is used for the convection of flow variables and fluxes across the sliding boundary. A deforming mesh algorithm using modified spring analogy is used for the blade motion. In the present paper, detail descriptions of numerical analysis for forward flight are introduced. Some results are presented for a two bladed AH-1G rotor and compared with experimental data.

• 한국항공우주학회지
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• 제50권7호
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• pp.497-505
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• 2022

본 연구에서는 회전익기 통합 해석 코드인 CAMRAD II를 이용하여 풍동 시험용 Lift-offset 동축 반전 로터의 성능 및 허브 진동 하중 해석을 수행하였다. 본 논문에서는 전체 로터의 추력을 트림 목표 값으로 사용하는 트림 조건 및 상/하 로터 각각의 추력을 트림 목표 값으로 이용하는 트림조건을 사용하였다. 두 종류의 트림 기법에 대하여 다양한 전진비 및 Lift-offset 값을 고려하여 해석을 수행하였고, 계산된 결과를 선행연구의 풍동 시험 및 해석 결과와 상호 비교하였다. 두 종류의 트림 기법을 이용한 연구 결과에서 모두 다양한 전진비 및 Lift-offset 값에 대하여 상/하 로터 각각의 양력, 토크 및 전체 로터의 로터 유효 양항비가 선행연구의 풍동 시험 결과와 유사함을 확인하였다. 또한 로터의 2P 허브 진동 하중 해석 결과를 기반으로 Lift-offset 동축 반전 로터 허브의 진동 하중 특성을 확인하였다. 본 연구를 통하여 CAMRAD II를 이용한 풍동 시험용 Lift-offset 동축 반전 로터의 모델링 및 해석 기법을 적절히 구축하였으며, 성공적으로 검증하였다.

• 한국소음진동공학회논문집
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• 제27권2호
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• pp.204-212
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• 2017

The rotor dynamics and balancing technic for rotating equipment during engineering and manufacture stage are to be carefully considered in order to minimize the operation troubles regarding vibration during commissioning stage. In this paper, the test rig, which includes the disks as balancing plane, is designed and manufactured, so that the characteristic of rotor dynamics can be analyzed such as critical speed and mode shape. The critical speed predicted through natural frequency analysis is verified by the actual measurement on bearing housing vibration during start-up condition of test rig. The low speed balancing and the operating speed balancing test are performed respectively with consideration of first critical speed, and the residual unbalance amounts are estimated in accordance with the relevant method described in API standard. In addition, the single and dual plane balancing are carried out on main disk and trim disk depended on phase information at each balancing step.

• International Journal of Aeronautical and Space Sciences
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• 제6권2호
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• pp.84-96
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• 2005

Helical tip vortex is known as stable vortex structure, however the specific frequency component of far wake perturbation induces the vortex pairing in hover and axial flight. It is expected that the tip vortex pairing phenomena may happen in transition flight and very low advance ratio flight so that inflow may be most nonuniform in the low advance ratio flight. The objectives of this paper are that a tip-vortex instability during the transition from hover into very low advance ratio forward flight is numerically predicted to understand a physics by using a time-marching free-wake method. To achieve the objectives, numerical method is firstly validated in typical axial and forward flights cases. Present scheme with trim routine can predict airloads and inflow distribution of forward flight with good accuracy. Then, the transition flight condition is calculated. The rotor used in this wake calculation is a small-scale AH-1G model. By using a tip-vortex trajectory tracking method, the tip-vortex pairing process are clearly observed in transient flight($\mu$=0.03) and disappears at a slightly higher advance ratio($\mu$=0.05). According to the steady flight simulation at $\mu$=0.03, it is confirmed the tip-vortex pairing process is continued in the rear part of rotor disk and not occurs in the front part. Time averaged inflow in this case is predicted as smooth distribution.