Collective Pitch Angle Research Articles

Experimental Study of an Extremely Flexible Rotor for Microhelicopters

This paper describes the development and testing of a rotor with extremely flexible blades, focusing on application to a microhelicopter. The goal is to develop rotor blades that are so flexible that they can be rolled into a compact volume and stowed inside the rotor hub. The design of an 18-in-diameter two-bladed rotor with flexible blades is presented. Stiffening and stabilization of the rotor blades are enabled by centrifugal forces acting on a tip mass. The relationship between blade-design parameters, aeroelastic stability boundaries, and hover performance is explored by a systematic set of experiments. It is observed that rotational speed and collective pitch angles have a significant effect on rotor stability. In addition, a large negative spanwise twist is induced in the flexible blades during stable operation, resulting in a decreased efficiency compared with rigid rotor blades having an identical planform and airfoil section. Different approaches to passively tailor the induced twist distribution are investigated. The highest efficiency was measured on a configuration making use of the propeller moment acting on a tip mass located at an index angle to the blade chord. A maximum figure of merit of 0.51 was measured at a blade loading of 0.14, which is comparable to the efficiency of rigid rotor blades.

Performance Measurements of a 1/6-Scale Model of the 1907 Cornu Rotor

Measurements of the performance of an approximate 1/6-scale Cornu 1907 rotor system are discussed. The rotor was tested in hover, both in and out of ground effect, over a range of rotational speeds and collective blade pitch angles. The measurements showed that the rotor had low aerodynamic efficiency, with a maximum attainable figure of merit of only about 0.35 for out of ground effect operating conditions. An analysis of the measurements showed that the performance of this rotor is dominated by relatively high induced losses, with an average induced power factor of over two, mainly because of the use of low-aspect-ratio blades. Extrapolation of the performance polar back to zero thrust with the aid of modified momentum theory suggests that the average profile drag coefficient for the blade sections was of the order of 0.1. However, even with the benefits of a proper airfoil section of lower drag, it is shown that without also using a blade design of higher aspect ratio, the figure of merit of the Cornu rotor design could not be improved much above 0.4. The measured results confirm previous modeling assumptions used for the aerodynamic and performance analysis of the Cornu rotor design and conclusions drawn regarding the unfeasibility of successful free flights of the Cornu helicopter.

Three‐time scale singular perturbation control and stability analysis for an autonomous helicopter on a platform

SUMMARYA three‐time scale singular perturbation control law is designed for a nonlinear helicopter model in vertical flight. The proposed control law is based on time scale decomposition and is able to achieve the desired altitude by selecting a desired angular velocity and the associated collective pitch angle of the blades. The stability of the system is performed by presenting a stability analysis for generic three‐time scale singularly perturbed systems, which allows to construct a composite Lyapunov function for the resultant closed‐loop system by using time scale separation and also providing mathematical expressions for the upper bounds of the singularly perturbed parameters that define the three‐time scale. Numerical results on both, the singular perturbation control strategy and the stability analysis, are also presented for the studied nonlinear highly coupled helicopter model. Copyright © 2012 John Wiley & Sons, Ltd.

H∞ Based Control for Load Mitigation in Wind Turbines

This article demonstrates a strategy to design multivariable and multi-objective controllers based on the H∞ norm reduction applied to a wind turbine. The wind turbine model has been developed in the GH Bladed software and it is based on a 5 MW wind turbine defined in the Upwind European project. The designed control strategy works in the above rated power production zone and performs generator speed control and load reduction on the drive train and tower. In order to do this, two robust H∞ MISO (Multi-Input Single-Output) controllers have been developed. These controllers generate collective pitch angle and generator torque set-point values to achieve the imposed control objectives. Linear models obtained in GH Bladed 4.0 are used, but the control design methodology can be used with linear models obtained from any other modelling package. Controllers are designed by setting out a mixed sensitivity problem, where some notch filters are also included in the controller dynamics. The obtained H∞ controllers have been validated in GH Bladed and an exhaustive analysis has been carried out to calculate fatigue load reduction on wind turbine components, as well as to analyze load mitigation in some extreme cases. The analysis compares the proposed control strategy based on H∞ controllers to a baseline control strategy designed using the classical control methods implemented on the present wind turbines.

정지비행 조건에서의 축소 로터 실험을 통한 소음 예측 기법 검증

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.

Performance Comparison of Two Airfoil Rotor Designs for an Agricultural Unmanned Helicopter

Purpose: The most important element of an agricultural helicopter is the rotor blade realizing lift force. In order to improve the performance of the rotor blades, two types (KA152313 and KB203611) of airfoils were designed and compared. Methods: The nose shape of the KB203611 airfoil was 'drooped' and 'sharp' compared to the leading edge of the KA152313 airfoil. The performance of the experimental airfoils was simulated using CFD-ACE program, and lifts were measured in situ using the 'AgroHeli-4G', a prototype helicopter. Results: Simulated lifts of the blade with the KA152313 airfoil showed proper values for a wide range of angles of attack between <TEX>$14^{\circ}{\sim}18^{\circ}$</TEX>, while the simulated lift of the KB203611 blade exhibited maximum values near <TEX>$13^{\circ}{\sim}14^{\circ}$</TEX>. In the lift measurements, the range of operable angles of attack was a collective pitch angle at the grip (GP) of <TEX>$12^{\circ}{\sim}18^{\circ}$</TEX> for the KA152313 blade. On the other hand, the range of angles of attack for the KB203611 blade was a GP of <TEX>$12^{\circ}{\sim}14^{\circ}$</TEX>. Conclusions: The blade of KA152313 performed well over a wide range of AoAs and the blade of KB203611 performed better at low AoAs. In this study, a variative airfoil blade, gradually emerging from grip to tip using the two different airfoils, was suggested.

Singular perturbation control strategy for a three-time-scale autonomous helicopter on a platform

A three-time-scale singular perturbation control law is designed for a nonlinear helicopter model in vertical flight. The proposed control law is based on time-scale decomposition, and is able to achieve the desired altitude by selecting a desired angular velocity, and the associated collective pitch angle of the blades. Numerical results for the singular perturbation control strategy, along with comparison results with the literature, are also presented for the studied nonlinear highly coupled helicopter model.

Intelligent Control of a Large Variable Speed Wind Turbine

We present two intelligent controllers for large and flexible wind turbines operating in high-speed winds, a Fuzzy-P + I and an adaptive neuro-fuzzy controller. The control objective is to regulate the rotor speed at the given rated power in region 3 (full load) via collective blade pitch angle. The modeled turbine is a three-bladed, upwind machine with a flexible blade and tower. We use the particle swarm optimization method in off-line training for our adaptive neuro-fuzzy controller. Numerical simulations are performed using wind inflow step change with a set of input–output data of a nonlinear wind turbine model. We compare the performance of the proposed controllers with the baseline PI-controller. Simulation results confirm successful performance of the proposed controllers.

Experimental hover performance evaluation on a small-scale rotor using a rotor test stand

This paper presents the work being carried out in order to deduce hover performance of a small-scale single rotor blade as a preliminary study of a small coaxial rotor helicopter development. As an initial research, a test stand capable of measuring thrust and torque of a small-scale rotor blade in hover state was constructed and fabricated. The test stand consists of three parts; a rotating device, a load measuring sensor and a data acquisition system. Thrust and torque were measured with varying collective pitch angle at fixed RPM. Through this research, hover performance tests were conducted for a small-scale single rotor blade operating in low Reynolds numbers (Re ≈ 3 × 105). The rotor blades investigated in this paper have maximum FM values varying from 0.59 to 0.65, which are low relative to modern full-scale helicopters. From these differences in FM between a small and a full-scale helicopter, the induced power factor is determined as varying from 1.35 to 1.42. Through this study, tests of hover performance were conducted for a single small-scale rotor blade, as well as verifying the test stand itself for the acquisition of hover performance.

SPR 기법을 이용한 회전 블레이드의 변형 및 모션 측정

A measurement system using stereo pattern recognition (SPR) method was configured to measure the rotor blade deformations and motions. An SPR-based measurement system was prepared using six stereo cameras. Through a series of experiments to evaluate the system measurement uncertainty, it was verified that the SPR system had less than 0.2mm standard uncertainty. The combined standard uncertainties for the lead-lag, flapping, and pitching motions were estimated as 0.296mm, 0.209mm, and , respectively. The SPR system was installed at a general small-scaled rotor test system at Korea Aerospace Research Institute. The blade motions and elastic deformation were successfully measured under the conditions with rotating speeds of 360rpm or 589rpm, and collective pitch angles of , , or . The advantages of the SPR system was analyzed in comparison with the measurement system used in Higher Harmonic Control Aeroacoustic Rotor Test -II.

부분적 지면조건 하에서의 소형 로터 블레이드 제자리 비행 성능에 대한 실험적 연구

본 논문에서는 부분적인 지면효과 하에서의 단일로터 블레이드에 대한 제자리 비행 성능 특성을 확인하였다. 이를 위해 소형 로터 블레이드 회전 시험장치와 이동식 바닥면을 이용하여 고정된 회전속도에서 콜렉티브 피치각을 변화시키면서 추력과 토크를 측정하였다. 로터면과 지면이 겹치는 거리를 d, 로터의 직경을 D라 하면 지면효과가 없을 때부터 d/D가 0.25일 때까지는 지면효과가 로터의 성능에 거의 영향을 미치지 않는 것으로 확인되었다. 로터 깃의 개수가 4개일 때 지면 효과의 영향을 더 크게 받아 2개일 때보다 추력이 증가하고 동력이 감소하였다. 또한 피치각이 작을수록 지면효과로 인하여 추력이 증가하는 경향을 보였다. 이와 같은 실험 결과를 바탕으로 로터 깃의 개수와 부분적인 지면 효과를 고려한 새로운 경험식을 도출하였다. This paper focuses on the hover performance experiment of a small-scale single rotor in partial ground conditions. In this study, small-scale rotor blade rotating device and floor panel are used to include partial ground effect. Thrust and torque were measured with varying collective pitch angles at fixed rotor rotating speed. The overlap distance between rotor and ground is d, the rotor diameter is D. It was shown that the ground effects have little effect on the rotor performance until d/D is 0.25. Four blade rotor has more increased thrust and more reduced power than those of two blade rotor because of stronger ground effect. In addition, it was also found that the thrust increases as a collective pitch angle become smaller. Based on these experiment results, we deduced new empirical equation considered blade number and partial ground effect.

소형 로터 블레이드의 제자리 비행 성능 시험장치 개발 및 검증

본 논문은 동축반전 로터 블레이드의 성능 특성 파악을 위한 선행연구로서, 단일 로터 블레이드에 대한 공력 데이터 획득 및 성능 특성을 확인하기 위한 연구에 중점을 두었다. 이를 위해 제자리 비행 상태에서 추력과 토크의 측정이 가능한 소형 로터 블레이드용 회전 시험장치를 구성하고 제작하였다. 로터 회전 시험장치는 회전장치부, 센서부, 그리고 데이터 획득 시스템으로 구성되었으며, 고정된 회전속도에서 콜렉티브 피치각을 변화시키면서 추력과 토크를 측정할 수 있도록 하였다. 이를 통하여 저 레이놀즈 수 (<TEX>$Re{\approx}3{\times}10^5$</TEX>) 영역에서 운용되는 단일 로터의 제자리 비행 성능 시험을 수행하였고 소형 로터 블레이드의 제자리 비행 성능을 획득하여 로터 회전 시험장치의 성능을 검증하였다. This paper presents the work being carried out in order to deduce hover performance of a small-scale single rotor blade as a preliminary study of a small coaxial rotor helicopter development. As an initial research, a test stand capable of measuring thrust and torque of a small-scale rotor blade in hover state was constructed and fabricated. The test stand consists of three parts; a rotating device, a load measuring sensor and a data acquisition system. Thrust and torque were measured with varying collective pitch angle at fixed RPM. Through this research, hover performance tests were conducted for a small-scale single rotor blade operating in low Reynolds number (<TEX>$Re\;{\approx}3{\times}10^5$</TEX>), as well as for verifying the test stand itself for acquiring hover performance.

Numerical Simulation of Autorotation in Forward Flight

Autorotation in forward flight is an important subject for gyroplanes or compound helicopters. By integrating the flapping and rotational equations of motion, we investigate in this work the steady-state autorotation of a three-bladed rotor for a given set of three independent variables: forward velocity, collective pitch angle, and shaft angle. Two-dimensional aerodynamic coefficients as functions of the angle of attack and the Reynolds number from the Navier-Stokes simulation and Pitt/Peters inflow model to determine induced flowfield are adopted to implement the simulation. Transient behavior of the solution from arbitrary initial conditions to a steady-state periodic solution is described. Variations of rotor speed and flapping angle of autorotation with the shaft angle and the collective pitch angle for a given forward velocity are presented and discussed in detail. Drastic changes in the region of autorotation near a specific shaft angle is found to exist, below which the range of collective pitch angle for stable autorotation becomes very limited and the flapping angle changes abruptly with the pitch angle.

Investigation of the Downwash Induced by Rotary Wings in Ground Effect

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.

Sizing and Performance-Prediction Method for Rotary-Wing Ejector Nozzles

A reaction-drive rotor configuration including an ejector nozzle is studied to determine possible exhaust-jet velocity reduction for the purpose of noise suppression. A vortex-wake rotor model is compared with an experimental rotor test, determining the model validity for a hovering rotor. A two-dimensional computational fluid dynamic model of an ejector flow is compared with an experimental ejector-nozzle test to determine the appropriate grid coarseness and turbulence model relevant to lift and drag predictions supporting the hovering-rotor aerodynamic model. The reaction-drive-rotor tip-jet-ejector nozzle geometry is sized by the analysis method put forward here, implementing coupled rotor aerodynamic, reaction-drive tip-jet-ejector thermodynamic, and computational fluid dynamic models. A performance version of the proposed analysis method is developed to characterize this rotor configuration over a range of blade collective pitch angles operating in a hovering-flight condition.

STUDY THE EFFECT OF QUADRATIC BLADE SHAPE ON THE VERTICAL AUTOROTATION PERFORMANCE OF A SMALL SCALE MODEL HELICOPTER

This paper presents the results of the numerical and experimental investigation of reducing the steady rate of descent of a typical small scale model helicopter. Power-off after hovering manoeuvre was picked for a case of study because of their critical induced velocity state. Contribution in every blade element in the effect of angle of attack, thrust and torque has been taken into consideration by blade element momentum theory for better results. In this analysis, test data of two types of blades with different driving region areas were obtained. The results of this study indicated that the expanding of the driving region area did reduce the steady rate of descent in vertical autorotation slightly but give lower rotor rotational speed compared to original shape blade. The study also had shown that zero collective pitch angle correction in the autorotation entry gives the best autorotation performance.

헬리콥터용 차세대 블레이드의 공력탄성학적 안정성에 관한 시험적 연구

This paper describes the aeroelastic stability test of the small-scaled `Next-generation Blade(NRSB)` with NRSH (next-generation hub system) and HCTH hingeless hub system in hover and forward flight conditions. Excitation tests of rotor system installed in GSRTS (general small-scale rotor test system) at KARI (Korea Aerospace Research Institute) were carried out to get lead-lag damping ratio of blades with flexures as hub flexure. MBA(moving block analysis) technique was used for the estimation of lead-lag damping ratio. First, NRSB-1F blades with HCTH hub system, then NRSB- 1F with NRSH hub system were tested. Second, NRSB-2F blades with NRSH hub system were tested. Tests were done on the ground and in the wind tunnel according to the test conditions of hover and forward flight, respectively. Non-rotating natural frequencies, non-rotating damping ratios and rotating natural frequencies were showed similar level for each cases. Estimated damping ratios of NRSB-1F, NRSB-2F with HCTH and NRSH were above 0.5%, and damping ratio increased by collective pitch angle increasement. Furthermore damping ratios of NRSB-2F were higher than damping ratios of NRSB-1F in high Pitch angle. It was confirmed that the blade design for noise reduction would give observable improvement in aeroelastic stability compared to paddle blade and NRSB-1F design.

마이크로폰 어레이를 이용한 회전하는 소음원 가시화에 관한 연구

Acoustic analysis of a moving sound source required that the measured sound signals be do-Dopplerized and restored as of the original emission signals. The purpose of this research is development of beamforming technique can be applied to the rotor noise source identification. For the do-Dopplerization and reconstruction of emitted sound wave, Forward Propagation Method is applied to the time domain beamforming technique. And validation test were performed using rotating sound source constructed by bended pipe and horn driver. In the validation test using sinusoidal sound wave, sufficient performance of signal processing can be seen, and the effect of measuring duration for accuracy was compared. In the prop-rotor measurements, the acoustic source locations were successfully verified in varying positions for different frequencies and collective pitch angle, in hover condition.

Embedded Laser Doppler Velocimetry Boundary-Layer Investigation on a Tilt-Rotor Blade

A fundamental study is made of a rotating boundary layer (BL) in the inboard region of a tilt-rotor blade, using an embedded laser Doppler velocimetry probe inside a rotating blade. Velocity measurements have been performed to get insight into the BL behavior around a blade section close to the rotor hub.,BL transition and separation phenomena are investigated on the basis of both chordwise and spanwise velocity profiles measured for different collective pitch angles on a two-bladed hovering rotor. The velocity. experimental data bank appears to be in good agreement with theoretical velocity profiles in attached How conditions (laminar, transitional, or turbulent). Moreover, validation of a transitional criterion based on the second thickness of energy delta'(3) of the BL is extended to the rotational flow configuration. Usually such a configuration classically exhibits a separation of the BL for high collective pitch angles. However, in the particular case of a high twisted tilt-rotor blade operating in, hover, centrifugal effects are shown to be responsible for an important delay in the BL separation process.

Numerical simulations of the unsteady flowfield of helicopter rotors on moving embedded grids

An Euler solver was developed for the prediction of 3-D unsteady flow phenomena that occur on helicopter rotors in forward flight, based upon a cell-centered finite volume approximation and dual-time method. In order to reduce false diffusion of vorticity and account for the rigid blade motions in rotation, flapping and pitching, a moving embedded grids methodology was adopted. Aiming at the key identification issue of donor elements in embedded grids, the present analysis implements a new searching scheme, the Pseudo-Searching Scheme of Donor Elements (PSSDE), for improvement purposes. The efficient dual-time method and moving embedded grids methodology make it feasible to compute the complicate flowfield of helicopter rotors on typical PC computers. Preliminary simulation results on surface pressure distributions were presented for a two-bladed Caradonna model rotor in forward flight at different collective pitch angles and flapping angles. Good agreements are found with measurements or other calculated results. Then unsteady lifting results, simulated for three different blade tip planforms, demonstrate the benefit of swept tips in suppressing supercritical flow. Additionally, the vorticity magnitude contours give some qualitative pictures on the evolvement of the distorted rotor wake.