The demand for new propeller designs has increased alongside the development of new technology, such as urban aircraft and large unmanned aerial vehicles. In order to experimentally identify the performance of a propeller, a wind tunnel that provides the operating flow is essential. However, in the case of a meter class or larger propeller, a large wind tunnel is required and the related equipment becomes heavy; therefore, it is difficult to implement in reality. For this reason, propeller studies have been conducted via reduced models. In this case, it is necessary to investigate the different performance outputs between the full- and model-scale propellers due to the size difference. In the current study, a method is proposed to investigate the difference in the aerodynamic performance caused by the difference in propeller scale using VLM and RANS calculations, and the differences are analyzed. The wind tunnel test also verified the propeller performance prediction method. The boundary of aerodynamic performance independent of the Reynolds number could be predicted through the VLM based on the ideal fluid assumption. From the RANS calculations, it was possible to present the difference in the aerodynamic performance when propellers of the same geometry with different ratios were operated using different Reynolds numbers. It was confirmed that each numerical method matched well with the wind tunnel test results in the range of the advance ratio that produced the maximum efficiency, and from the results, it was possible to observe the change in aerodynamic performance that differed according to the scale change.

This paper identifies the phase lag of a coaxial helicopter UAV to remove the effect of the structural vibration and rotor wake by improper phase lag. This vibration affects the sensor measurement of the flight control computer and deteriorates the attitude control performance of the internal controller. The characteristics that depend on the phase lag in the frequency band are found to identify the phase lag. Then, control commands, body velocity, and angular velocity are visualized and analyzed in a three-dimensional vector field normalized according to the phase lag in a specific motor rotation speed. Additionally, the phase lag is analyzed in the frequency domain using a Fast Fourier Transform using the angular velocity obtained from the gyro. The angular velocity, except for the motor rotation frequency, is filtered according to the cutoff frequency through 2D and 3D power spectrograms to improve the inner-loop control performance. Finally, waypoint flight is performed by applying the selected phase lag for control mixing and inner-loop cutoff frequency.

This paper conducts parametric studies on flapping wing design, one of the most important design parameters of insect-mimicking aerial vehicles. Experimental study on wing shape was done through comparison and analysis of thrust, pitching moment, power consumption, and thrust-to-power ratio. A two-axis balance and hall sensor measure force and moment, and flapping frequency, respectively. Wing configuration is biplane configuration which can develop clap and fling effect. A reference wing shape is a simplified dragonfly’s wing and studies on aspect ratio and wing area were implemented. As a result, thrust, pitching moment, and power consumption tend to increase as aspect ratio and area increase. Also, it is found that the flapping mechanism was not normally operated when the main wing has an aspect ratio or area more than each certain value. Finally, the wing shape is determined by comparing thrust-to-power ratio of all wings satisfying the required minimum thrust. However, the stability is not secured due to moment generated by disaccord between thrust line and center of gravity. To cope with this, aerodynamic dampers are used at the top and bottom of the fuselage; then, indoor flight test was attempted for indirect performance verification of the parametric study of the main wing.

The endurance of the multi-copter is one of the important variables that determine the mission performance. Therefore, accurate endurance should be defined as essential for performing effective missions. In this paper, we present the results of the study on the flight performance of the aircraft, especially the hovering of the drone(multi-copter). Unlike conventional aircraft, which consider aerodynamic performance by the fuselage, the multi-copter is mostly determined by the propulsion system. Therefore, the research method classifies the various parts constituting the drone system into functions, analyzes the performance of the unit parts and obtains the experimental data by sorting out the specifications and functions at the component level and mathematical formulation, The results of this study are as follows. In addition, the 5kg class quad copter was used to predict and verify the voltage change with endurance through analysis of in situ flight. By predicting endurance under various conditions, it can help design/build the right Multi-copter for mission.

This paper discussed a noise reduction effect of airfoil and small-scale model rotor by using attached serration trailing edge in the wind tunnel test condition. In order to analyze the changes in the performance due to the inclusion of a serrated trailing edge designed to reduce noise, a 10 kW wind turbine rotor was equipped with a thin serrated trailing edge. The restrictive condition for the serrated trailing edge equipped with the using of a 2D airfoil was examined through the using of a wind tunnel experiment after studying existing restrictive condition and analyzing prior research on serrated trailing edges. The aerodynamic performance and noise reduction effect of a small-scale model were investigated with the using of a serrated trailing edge. Moreover, the noise levels from the experiment were considered that the noise prediction method could be used for a full-scale rotor. It is confirmed that noise reduction effect is compared with wind tunnel test data at the 2D airfoil and model rotor condition.

This article describes a wind tunnel experiment that was undertaken to investigate the changes in the aerodynamic performance of a wind turbine due to the inclusion of a 2D airfoil with a serrated trailing edge designed to reduce the noise caused by a wind turbine rotor blade. The restrictive condition for the serrated trailing edge equipped with the use of a 2D airfoil was examined through the use of a wind tunnel experiment after studying existing restrictive condition and analyzing prior research on serrated trailing edges. The study was conducted according to Howe's theory, which is a cornerstone of the study of noise reduction effects produced by a serrated trailing edge. For the serrated trailing edge equipped on a 2D airfoil, the wake distribution and the relation to noise were analyzed in order to determine the restrictive condition in accordance with Howe's theory. The results indicated that an empirical formula or a theoretical approach should consider changes in the boundary layer thickness of a 2D airfoil, so an empirical noise prediction formula is suggested for the serrated trailing edge. Also, a comparison and an analysis of the prediction and the experimental results for the noise produced by the NACA0012 or the baseline airfoil equipped with a serrated trailing edge suggested a novel formula for a 2D airfoil. Finally, it is confirmed that 2D airfoil noise data are compared with wind tunnel test data by using an empirical formula estimationmethod. (C) 2017 Institute of Noise Control Engineering.

The development of a low-noise wind turbine rotor and propeller is often cost-effective and is in fact a race against time to those who wish to build and test a small-scale rotor instead of an expensive full-scale rotor. The issue of this approach has to do with the interpretation of wind tunnel model test data in terms of both the frequency band and sound pressure level information for the noise scaling effect. This paper discusses a prediction method for the estimation of the noise generated from a full-scale wind turbine rotor using wind tunnel test data measured with both a small-scale rotor and a 2D section of the blade. The 2D airfoil self-noise and the scaled rotor noise were investigated with a series of wind tunnel experiments. Wind tunnel data post-processing considered four aspects: removal of the test condition effect, scaling to full scale, consideration of the wind turbine rotor operating conditions, and the most important terms of full-scale rotor noise as adjustments to address the differences between the wind tunnel test conditions and the full-scale operating conditions. A full-scale rotor noise prediction results comparison was performed by initially dividing the test conditions into the condition of a 2D section noise test and the condition of a small-scale rotor noise test. Based on an airfoil section, the rotor was selected from a blade section at r / R = 0.75. The small-scale rotor was scaled down by a factor of 5.71 for the wind tunnel test. Finally, the full-scale rotor noise data was compared with the wind tunnel test data using a scaling estimation method.

This paper discusses the procedure of a blockage effect correction method involving small-scale wind turbine rotor experimental data. To simulate the aerodynamic performance of full-scale rotors in the field, however, measured data from scaled model experiments need to be analyzed appropriately. One of the most important elements of such an analysis is a procedure to remove the blockage effect of the wind tunnel wall from the measured power data. In this paper, a correction algorithm proposed as part of Glauert's blockage effect correction method is used to process the data from a wind turbine rotor tested with three different wind tunnel sizes. Also, this study considered the modified blockage effect correction method, which has been used to process the rotor thrust data in closed-circuit wind tunnels and open-circuit wind tunnels. A small-scale rotor was tested under the same operating conditions, i.e., the same advance ratio, rotating speed, rotor torque and speed of the wind tunnel. The small-scale wind turbine rotor has a diameter of 1.408 m and a rotating speed according to the tip speed ratio. In each case, the effect of the blockage ratio and aerodynamic characteristics are determined using wind tunnel test results and with a simple analytical correction method. The results of the modified correction method show that the aerodynamic performance levels during a wind tunnel test are cleared by the blockage effect.(C) 2014 Elsevier Ltd. All rights reserved.

This paper discussed a procedure for noise certification of Aircraft and predicting the full scale over-flight noise of propeller from acoustic wind tunnel measurement of small scale propeller.Noise Certification Procedures is established from International Civil Aviation Organization(ICAO).The data manipulations are then discussed in extrapolation to simulation flight distance and flight simulation.One of the most important point of flight simulation is adjustments for differences between wind tunnel test conditions and flight test conditions.To simulated the noise level estimation procedure for noise data post-process, simulate procedures from data of the wind tunnel noise measurement and the flight noise measurement by using a 7kg degree UAV.This study confirmed an effectively noise estimation procedures by wind tunnel noise test and flight noise test.초 록 본 논문에서는 풍동 실험을 통해 취득한 소음측정 결과에서 실제 비행에 기체에서 발생 되는 소음을 예측하는 과정에 대하여 다루었으며 정지된 소음원을 비행하는 것과 같은 상 태로 시뮬레이션 할 수 있는 방법에 대하여 소개하고 실험을 통하여 검증하였다.또한 국 제민간항공기구(이하 ICAO)에서 규정한 절차에 따라 틸트로터 항공기 및 무인항공기의 환 경소음을 평가할 수 있는 방안을 제시하였다.제시한 평가 방안을 검증하기 위해서 풍동 실험과 비행시험에 대한 소음실험 시뮬레이션 프로그램을 구성하였으며 7kg급 무인항공기 를 통한 풍동 실험 및 비행 실험을 수행하였으며 풍동 실험 결과로부터 비행 상태의 소음 예측이 가능하다는 것을 확인 할 수 있었다.

To reduce airfoil self-noise from a 10 kW wind turbine, we modified the airfoil shape and planform of a wind turbine blade. To obtain the optimal blade design, we used optimization techniques based on genetic algorithms. The optimized airfoil was first determined based on a section of the rotor blade, and then the optimized blade was designed with this airfoil. The airfoil self-noise from the rotor blades was predicted by using a semi-empirical model. The numerical analysis indicates that the level of the airfoil self-noise from the optimized blade is 2.3 dB lower than that from the baseline blade at the rated wind speed. A wind tunnel experiment was also performed to validate the design optimization. The baseline and optimized rotors were scaled down by a factor of 5.71 for the wind tunnel test. The experimental results showed that airfoil self-noise is reduced by up to 2.6 dB.

본 논문에서는 헬리콥터 전진비행 조건에서 플래핑모션에 대하여 축소 로터 실험을 통하여 실험결과와 이론적인 예측결과를 비교하였다. 축소로터 성능 실험은 충남대학교 아음속 풍동에서 수행하였으며 1.8 x 1.8m의 개방형 시험부를 사용하였다. 전진비행조건에서의 실험 결과에 의하면 축소로터의 전진비행조건에서의 추력 결과를 고정한 조건에서 동력계수는 차이가 있는 것을 확인 할 수 있었다. 또한, 공력 성능 측정 결과와 이론적인 예측결과의 비교를 통하여 헬리콥터의 플랩핑 각도의 범위에 대하여 비교하여 보았다. Coning 각도, 횡방향과 종방향에 대한 플래핑 각도에 대해서는 실험결과와 예측결과의 유사함을 확인하였다. In this paper, wind tunnel test and analytical prediction are compared for result of flapping motion in helicopter forward flight condition. Tests were performed at low speed wind tunnel at Chungnam National University, test section of wind tunnel has 1.8 by 1.8 meter open-jet test section area. According to the results of measured data for aerodynamic performance of model rotor in forward flight. It has to observed the difference of analytical and measured results of power coefficient for fixed thrust coefficient. And calculated and measured data of helicopter rotor flapping angles in forward flight are compared for a model rotor in a wind tunnel. A test was conducted to verify the measured data of coning and lateral/longitudinal flapping angle with predicted values.

본 논문에서는 정지비행 조건에서의 무향실 내 축소 로터 실험을 이용해 Lowson의 하중 소음식과 FW-H의 음향상사식으로 예측한 이산 주파수 소음(Discrete frequency noise)을 검증하였다. 소음 예측 기법의 방향성(Directivity) 검증은 전반적으로 실험결과와 유사하게 예측되었으며, 거리에 대한 검증의 경우 근거리(Near-field)에서는 FW-H식의 예측결과가, 원거리(Far-field)에서는 Lowson식의 예측결과가 실험결과와 더 유사한 것을 확인하였다. 피치 각(Collective pitch angle)에 대한 검증의 경우 낮은 피치각에서는 FW-H식의 예측결과가, 높은 피치각에서는 Lowson식의 예측결과가 실험결과와 더 유사한 것을 확인하였다. In this paper, a series of experiment is performed for a scaled hovering rotor in a semi-anechoic chamber and the results are compared to the noise spectra predicted by using Lowson's loading noise equation and FW-H equation. It was founded that the sound directivity pattern for both experiments and predictions are similar in their trend. Meanwhile the FW-H equation showed better agreement with experiments in the near-field noise spectra, but at the far-field the Lowson's equation performed better. The discrete noise are known to be proportional to the loading on the blades, which can be controlled by collective pitch angle of the blades. It was founded that the predicted spectra with FW-H equation come close to the measured noise spectra in low collective pitch, but in high collective pitch angles the Lowson's equation be more reliable.

In this paper, a selection method of efficient electric propulsion system for small size fixed wing UAV using wind tunnel experimental data base. It will be discussed in three parts: measurements of propeller performance in wind tunnel conditions, measurements of electric motor characteristics using dynamometer, and the developments of propulsion performance prediction method using experimental performance data of its components i. e., propeller and motor. Performances of propellers were measured at low speed wind tunnel at Chungnam National University. Test section of wind tunnel has 1. 25 by 1. 25 meter cross section area and maximum operational speed is 70m/s. 16 RC model propellers which has 9 to 14 inch diameter were tested, at a different operating condition of advance ratio, rotating speed, torque of propeller and speed of wind tunnel were acquired. As a validation of proposed method, predicted matching points were compared with experimental results. And propulsion performance of a 10kg class close-range mini-UAV was compared for the different propeller. At a cruise condition of that UAV an optimum propeller can be selected by the developed analysis procedure, and the enhancement of endurance can be expected from propeller change.

본 연구에서는 축소형 로터 실험 장치에서 로터 회전면에 수직으로 작용하는 추력(Fz)과 평행으로 작용하는 2축 모멘트(Mx, My)를 측정할 수 있는 3분력 Rotating Balance를 설계, 제작하고 성능실험을 통하여 검증하였다. Rotating Balance에 장착된 스트레인게이지의 미세신호는 로터와 함께 회전하는 신호증폭기를 거쳐 증폭되어 slip ring을 통하여 데이터 취득장치로 전달되어진다. 정적 보정법을 통하여 얻어진 보정행렬식의 적합성 여부를 판단하기 위하여 기존에 개발된 reaction type balance를 사용하여, 로터가 회전시 발생하는 추력을 비교함으로서 측정 정밀도를 확인하였다. 또한, Rotating Balance의 확장 불확도는 A형 불확도와 B형 불확도를 계산하였고, 계산 결과로 부터 $2.82\times10^{-1}$의 값을 얻음으로서 신뢰성을 확인하였다. In this study a 3-component rotating balance, which is designed to measure the thrust (Fz) and two moment components (Mx, My) simultaneously for a rotor test jig, is designed and its performance is validated experimentally. The low voltage signal from the strain gages mounted on the balance is amplified with a rotating amplifier, which is then fed through a slip-ring unit into the data acquisition system. In order to validate the accuracy of the calibration matrix obtained from a static calibration test, an additional reaction type balance is used to measure the thrust from a model rotor simultaneously, and shows very good result. Finally, the expanded uncertainty value, which is obtained from ISO method is estimated to be $2.82\times10^{-1}$, and the balance turns out to be reliable.