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Article
Publication date: 6 July 2010

A. Arun Kumar, S.R. Viswamurthy and R. Ganguli

This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.

Abstract

Purpose

This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.

Design/methodology/approach

Aeroelastic analysis of helicopter rotor with elastic blades based on finite element method in space and time and capable of considering higher harmonic control inputs is carried out. Moderate deflection and coriolis nonlinearities are included in the analysis. The rotor aerodynamics are represented using free wake and unsteady aerodynamic models.

Findings

Good correlation between analysis and HART‐II wind tunnel test data is obtained for blade natural frequencies across a range of rotating speeds. The basic physics of the blade mode shapes are also well captured. In particular, the fundamental flap, lag and torsion modes compare very well. The blade response compares well with HART‐II result and other high‐fidelity aeroelastic code predictions for flap and torsion mode. For the lead‐lag response, the present analysis prediction is somewhat better than other aeroelastic analyses.

Research limitations/implications

Predicted blade response trend with higher harmonic pitch control agreed well with the wind tunnel test data, but usually contained a constant offset in the mean values of lead‐lag and elastic torsion response. Improvements in the modeling of the aerodynamic environment around the rotor can help reduce this gap between the experimental and numerical results.

Practical implications

Correlation of predicted aeroelastic response with wind tunnel test data is a vital step towards validating any helicopter aeroelastic analysis. Such efforts lend confidence in using the numerical analysis to understand the actual physical behavior of the helicopter system. Also, validated numerical analyses can take the place of time‐consuming and expensive wind tunnel tests during the initial stage of the design process.

Originality/value

While the basic physics appears to be well captured by the aeroelastic analysis, there is need for improvement in the aerodynamic modeling which appears to be the source of the gap between numerical predictions and HART‐II wind tunnel experiments.

Details

Aircraft Engineering and Aerospace Technology, vol. 82 no. 4
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 2 January 2018

Yu Hu, Hailang Zhang and Gengqi Wang

This paper aims to investigate the mechanisms lying behind the cycloidal rotor under hovering status.

Abstract

Purpose

This paper aims to investigate the mechanisms lying behind the cycloidal rotor under hovering status.

Design/methodology/approach

Experiments were conducted to validate the numerical simulation results. The simulations were based on unsteady Reynolds-averaged Navier–Stokes (URANS) equations solver and the sliding mesh technique was used to model the blade motion. 2D and 2.5D simulations were made to investigate the 3D effects of turbulence. The effects of pressure and viscosity were compared to study the significance of the blade motion on force generation.

Findings

The 2.5D numerical simulation cannot produce more accurate results than the 2D counterpart. The pitching motion of the blade results in dynamic stall. The dynamic stall vortices induce parallel blade vortex interaction (BVI) upon downstream blades. The interactions between the blades delay the stall of the blade which is beneficial to the thrust generation. The blade pitching motion is the dominant contributor to the force generation and the turbulence is the secondary. Strong downwash in the rotor cage varied the inflow velocity as well as the effective angle of attack (AOA) of the blade.

Practical implications

Cycloidal rotor is a propulsion device that can provide omni-directional vectored thrust with high efficiency and low noise. To understand the mechanisms lying behind the cycloidal rotor helps the authors to design efficient cycloidal rotors for aircraft.

Originality/value

The authors discovered that the blade pitching motion plays primary role in force generation. The effects of the dynamic stall and BVI were studied. The reason why cycloidal rotor can be more efficient was discussed.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 1
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 31 January 2020

Zehba A.S. Raizah

The purpose of this study is to apply the incompressible smoothed particle hydrodynamics method for simulating the natural convection flow inside a cavity including cross…

Abstract

Purpose

The purpose of this study is to apply the incompressible smoothed particle hydrodynamics method for simulating the natural convection flow inside a cavity including cross blades or circular cylinder cylinder.

Design/methodology/approach

The base fluid is water and copper-water nanofluid is treated as a working fluid. The left and rights walls are maintained at a cool temperature, the horizontal cavity walls are isolated and the inner shape was heated. The physical parameters are the length of the blades L_Blade, the number of cross blades, circular cylinder radius L_R, Rayleigh number Ra and the nanoparticles volume fraction.

Findings

The results reveal that the lengths of the cross blade, number of the blades and radius of the circular cylinder is working as an enhancement factor for heat transfer and fluid flows inside a cavity. Adding nanoparticles augments heat transfer and reduces the fluid flow intensity inside a cavity. The best case for buoyancy-driven flow was obtained when the inner shape is the circular cylinder at a higher Rayleigh number.

Originality/value

This work uses a distinctive numerical method to study the natural convection heat from cross blades inside a cavity filled with nanofluid. It provides a new analysis of this issue and presented good results.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 30 no. 10
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 14 January 2019

Jaroslaw Stanislawski

The purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon.

Abstract

Purpose

The purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon.

Design/methodology/approach

The considered physical model of helicopter standing on ground with rotating rotor consists of fuselage and main transmission gear treated as stiff bodies connected by elastic elements. The fuselage is supported on landing gear modeled by spring-damper units. The main rotor blades are treated as set of elastic axes with lumped masses distributed along blade radius. Due to Galerkin method, parameters of blades motion are assumed as a combination of bending and torsion eigen modes. A Runge–Kutta method is applied to solve equations of motions of rotor blades and helicopter fuselage.

Findings

The presented simulation method may be applied in preliminary stage of helicopter design to avoid ground resonance by proper selection of landing gear units and blade damper characteristics.

Practical implications

Ground resonance may occur in form of violently increasing mutual oscillations of helicopter fuselage and lead-lag motion of rotor blades. According to changes of stiffness and damping characteristics, simulations show stable behavior or arising oscillations of helicopter. The effects of different blade balance or defect of blade damper are predicted.

Originality/value

The simulation method may help to determine the envelope of safe operation of helicopter in phase of take-off or landing. The effects of additional disturbances as results of blades pitch control as swashplate deflection are introduced.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 3
Type: Research Article
ISSN: 1748-8842

Keywords

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Article
Publication date: 6 July 2015

Chengwei Fei, Wenzhong Tang, Guangchen Bai and Shuang Ma

This paper aims to reasonably quantify the radial deformation of turbine blade from a probabilistic design perspective. A probabilistic design for turbine blade radial…

Abstract

Purpose

This paper aims to reasonably quantify the radial deformation of turbine blade from a probabilistic design perspective. A probabilistic design for turbine blade radial deformation considering non-linear dynamic influences can quantify risk and thus control blade tip clearance to further develop the high performance and high reliability of aeroengine. Moreover, the need for a cost-effective design has resulted in the development of probabilistic design method with high computational efficiency and accuracy to quantify the effects of these uncertainties.

Design/methodology/approach

An extremum response surface method-based support vector machine (SVM-ERSM) was proposed based on SVM of regression to improve the computational efficiency and precision of blade radial deformation dynamic probabilistic design regarding non-linear material properties and dynamically thermal and mechanical loads.

Findings

Through the example calculation and comparison of methods, the results show that the blade radial deformation reaches at the maximum at t = 180 s; the probabilistic distribution and inverse probabilistic features of output parameters and the major factors (rotor speed and gas temperature) are gained; besides, the SVM-ERSM holds high computational efficiency and precision in the non-linear dynamic probabilistic design of aeroengine typical components.

Practical implications

The present efforts provide a method to design turbine besides other aeroengine components considering dynamic and non-linear factors base on probabilistic design for further research.

Social implications

Moreover, the present study provides a way to design dynamic (motion) structures from a probabilistic perspective.

Originality/value

It is proved that the dynamic probabilistic design-based SVM-ERSM could produce a more reasonable blade radial deformation while maintaining low failure probability, as well as offer a useful reference for blade-tip clearance control and a promising insight to the optimal design of aeroengine typical components.

Details

Aircraft Engineering and Aerospace Technology: An International Journal, vol. 87 no. 4
Type: Research Article
ISSN: 0002-2667

Keywords

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Article
Publication date: 8 May 2018

Fernando Tejero Embuena, Piotr Doerffer, Pawel Flaszynski and Oskar Szulc

Helicopter rotor blades are usually aerodynamically limited by the severe conditions present in every revolution: strong shock wave boundary layer interaction on the…

Abstract

Purpose

Helicopter rotor blades are usually aerodynamically limited by the severe conditions present in every revolution: strong shock wave boundary layer interaction on the advancing side and dynamic stall on the retreating side. Therefore, different flow control strategies might be applied to improve the aerodynamic performance.

Design/methodology/approach

The present research is focussed on the application of passive rod vortex generators (RVGs) to control the flow separation induced by strong shock waves on helicopter rotor blades. The formation and development in time of the streamwise vortices are also investigated for a channel flow.

Findings

The proposed RVGs are able to generate streamwise vortices as strong as the well-known air-jet vortex generators. It has been demonstrated a faster vortex formation for the rod type. Therefore, this flow control device is preferred for applications in which a quick vortex formation is required. Besides, RVGs were implemented on helicopters rotor blades improving their aerodynamic performance (ratio thrust/power consumption).

Originality/value

A new type of vortex generator (rod) has been investigated in several configurations (channel flow and rotor blades).

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 28 no. 5
Type: Research Article
ISSN: 0961-5539

Keywords

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Article
Publication date: 3 January 2017

Farid Shahmiri

The aim of this paper was to experimentally examine twin-rotor hover performance for different rotor overlap ratios at practical rotor loading.

Abstract

Purpose

The aim of this paper was to experimentally examine twin-rotor hover performance for different rotor overlap ratios at practical rotor loading.

Design/methodology/approach

The methodology was formed based on data measurements for a designed twin-rotor test model and development of hover performance mathematical models. Thus, measurements were made using a central composite test plan, and then mathematical models for thrust power required power loading (PL) and figure of merit (FM) as functions of collective pitch tip speed; rotor overlap ratio was obtained. In the present paper, the test model consisted of two three-bladed rotors with a diameter of 220 mm and a blade aspect ratio of 16.05. The blades were of a rectangular planform with NACA 0012 cross sections and had no twist or taper. The model was built such that the rear rotor was fixed on the fuselage, and the front rotor could move longitudinally for tests up to about 40 per cent overlap ratio in hover.

Findings

The best hover aerodynamic efficiency (maximum PL of 14.6 kg/kW) was achieved for non-overlapped rotors at a low value of disc loading (DL) and also at FM of 0.6 at that DL. This result was in agreement with blade element momentum theory predictions.

Practical implications

Results for the twin-rotor test model can be generalized for actual tandem helicopters through the Reynolds number transformation technique and also some modifications.

Originality/value

Design and construction of the twin-rotor test model and experimental measurements of hover performance based on an optimal test plan were performed for the first time.

Details

Aircraft Engineering and Aerospace Technology, vol. 89 no. 1
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 8 May 2018

Javier Martinez Suarez, Pawel Flaszynski and Piotr Doerffer

The purpose of this paper is to describe numerical investigations focused on the reduction of separation and the aerodynamic enhancement of wind turbine blades by a rod…

Abstract

Purpose

The purpose of this paper is to describe numerical investigations focused on the reduction of separation and the aerodynamic enhancement of wind turbine blades by a rod vortex generator (RVG).

Design/methodology/approach

A flow modelling approach through the use of a Reynolds-averaged Navier–Stokes solver is used. The numerical tools are validated with experimental data for the NREL Phase VI rotor and the S809 aerofoil. The effect of rod vortex generator’s (RVG) configuration on aerofoil aerodynamic performance, flow structure and separation is analysed. RVGs’ chordwise locations and spanwise distance are considered, and the optimum configuration of the RVG is applied to the wind turbine rotor.

Findings

Results show that streamwise vortices created by RVGs lead to modification of flow structure in boundary layer. As a result, the implementation of RVGs on aerofoil has proven to decrease the flow separation and enhance the aerodynamic performance of aerofoils. The effect on flow structure and aerodynamic performance has shown to be dependent on dimensions, chordwise location and spanwise distribution of rods. The implementation of devices with the optimum configuration has shown to increase aerodynamic performance and to significantly reduce separation for selected conditions. Application of rods to the wind turbine rotor has proven to avoid the spanwise penetration of flow separation where applied, leading to reduction of flow separation and to aerodynamic enhancement.

Originality/value

The proposed RVGs have shown potential to enhance the aerodynamic performance of wind turbine rotors and profiles, making devices an alternative solution to the classical vortex generators for wind turbine applications.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 28 no. 5
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 5 September 2016

Andrea Andrisani, Diego Angeli and Antonio Dumas

The purpose of this paper is to define an optimal pitching profile for the blades of a cycloidal rotor which minimizes the mean power consumption for a given mean thrust…

Abstract

Purpose

The purpose of this paper is to define an optimal pitching profile for the blades of a cycloidal rotor which minimizes the mean power consumption for a given mean thrust of the rotor.

Design/methodology/approach

A simple analytical model of the kinematics and aerodynamics of a cycloidal rotor is defined first to obtain expressions for thrust and power depending on the pitching profile and geometrical parameters of the rotor. Then, Lagrange optimization is applied to obtain the optimal pitching schedule under hovering conditions. Finally, results of the theoretical analysis are compared with those of a two-dimensional computational fluid dynamics (CFD) model.

Findings

Results of the optimization suggest that the optimal profile is a combination of sinusoidal functions. It is shown that the adoption of the optimal pitching schedule could improve the power efficiency of the rotor by approximately 25 per cent.

Practical implications

The possibility to increase the efficiency of a cycloidal rotor by acting on its pitching schedule could be a significant factor of success for this alternative propulsion concept.

Originality/value

The present work represents the first attempt at a definition of an optimal pitching profile for a cycloidal rotor. Moreover, although being carried out on the basis of simplified analytical considerations, the present investigation sets a methodological framework which could be successfully applied to the design of similar kinds of systems.

Details

Aircraft Engineering and Aerospace Technology, vol. 88 no. 5
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 1 October 2002

Jinwu Xiang, Guocai Hu and Xiaogu Zhang

An equivalent linear damping model is developed for forward flight condition, with the flap/lag/pitch kinematics and nonlinear characteristics of hydraulic damper taken…

Abstract

An equivalent linear damping model is developed for forward flight condition, with the flap/lag/pitch kinematics and nonlinear characteristics of hydraulic damper taken into account. Damper axial velocity is analyzed from the velocities of the damper‐to‐blade attachment point in time domain. For the case of blade lead‐lag oscillations without forced excitation and kinematics, the equivalent linear damping is calculated from transient response with energy balance method, Fourier series based moving block analysis and Hilbert transform based technology, respectively. Results indicate that equivalent linear damping decreases significantly with lead‐lag forced excitation and flap/lag/pitch kinematics, especially with the latter in flight condition.

Details

Aircraft Engineering and Aerospace Technology, vol. 74 no. 5
Type: Research Article
ISSN: 0002-2667

Keywords

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