Showing papers on "Blade element theory published in 2004"

Biologically Inspired Design of Small Flapping Wing Air Vehicles Using Four Bar Mechanisms and Quasi-Steady Aerodynamics

Abstract: In this paper, the energetics of a flapping wing micro air vehicle is analyzed with the objective of design of flapping wing air vehicles. The salient features of this study are: (i) design of an energy storage mechanism in the air vehicle similar to an insect thorax which stores part of the kinetic energy of the wing as elastic potential energy in the thorax during a flapping cycle; (ii) inclusion of aerodynamic wing models using blade element theory and inertia of the mechanism using rigid body modeling techniques; (iii) optimization of parameters of the energy storage mechanism using the dynamic models so that the peak power input from the external actuators during a flapping cycle is minimized. A series of engineering prototypes based on these studies have been fabricated which justify the use of these mathematical techniques.

Dual Purpose Design of Small Wind Turbine Blades

Abstract: Experimental studies of the starting behaviour of small wind turbines have shown that the initial "idling period", characterised by small rotor acceleration, is usually much longer than the subsequent period of rapid accleration to reach operational rotor speed. Idling obviously reduces the power generation potential of any turbine. The experimental results also imply that most starting torque is generated near the hub, whereas power-producing torque is concentrated near the tip. Therefore this paper considers the computational design of blades for small wind turbines for the dual purposes of (i) efficient power production at rated wind speeds and (ii) rapid starting at smaller wind speeds. Standard blade element theory is used to determine the power coefficient, which is the first objective function to be maximised. A modified blade element method gives the start, whose inverse is the second objective function. During the idling period, the blade angles of attack are relatively large, allowing the lift a...

Helicopter vibration reduction in forward flight with induced-shear based piezoceramic actuation

Abstract: Governing equations are obtained for helicopter rotor blades with surface bonded piezoceramic actuators using Hamilton's principle. The equations are then solved for dynamic response using finite element discretization in the spatial and time domains. A time domain unsteady aerodynamic model is used to obtain the airloads. The nonlinear relationship between the piezoelectric shear coefficient and applied ac field is represented as a polynomial curve fit. The nonlinear effects are investigated by applying a sinusoidal voltage to the helicopter rotor blade. The rotor blade is modeled as a two-cell box section with piezoelectric layers surface bonded to the top and bottom of the box beam. Comparison of results with linear and nonlinear shear coefficients is presented. Use of a nonlinear relationship (compared to linear) to achieve targeted reductions in strains or displacements results in a reduction in the requirement of applied amplitude of the sinusoidal field. A rate feedback control law is implemented which feeds back the higher harmonics of the time rate of change of strain in the azimuthal direction. The sensed voltage is then applied to the rotor blade, resulting in a vibration reduction of approximately 43% for a four-bladed, soft-in-plane hingeless rotor in forward flight.

A near wake model for trailing vorticity compared with the blade element momentum theory

Abstract: A near wake model for trailing vorticity originally proposed by Beddoes for high-resolution helicopter blade vortex interaction computations has been implemented and compared with the usual blade element momentum models used for wind turbine calculations. The model is in principle a lifting line model for the rotating blade, where only a quarter revolution of the wake system behind the blade is taken into account. This simplification of the wake enables a fast computation of the downwash from the trailed vortex system along the blade using the indicial function method and thus makes it realistic to use the model in aeroelastic time simulations. The downwash from the shed vorticity is also computed with a fast indicial function algorithm. In particular the model is investigated for use in calculations of aerodynamic damping for the different mode shapes of an operating wind turbine. Numerical results for the downwash of a wing in straight flow with elliptical circulation are compared with analytical results. Further, the downwash distribution of a 40 m long rotating blade is computed. Aerodynamic damping of the blade in axial harmonic translation and in the first flapwise mode is computed with the near wake model and compared with the results of a standard momentum model including a model for dynamic inflow. Copyright © 2004 John Wiley & Sons, Ltd.

Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor—Part I: Unsteady Profile Pressures and the Effect of Clocking

Abstract: This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage Low-Speed Research Compressor of Dresden. In part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are calculated from the measured profile pressure distributions. The unsteady pressure distributions were analysed in the first, a middle and the last compressor stage both on the rotor and stator blades. The measurements were carried out on pressure side and suction side at midspan. Several operating points were investigated. A complex behaviour of the unsteady profile pressures can be observed, resulting from the superimposed influences of the wakes and the potential effects of several up- and downstream blade rows of the four-stage compressor. The profile pressure changes nearly simultaneously along the blade chord if a disturbance arrives at the leading edge or the trailing edge of the blade. Thus the unsteady profile pressure distribution is nearly independent of the convective wake propagation within the blade passage. A phase shift of the reaction of the blade to the disturbance on the pressure and suction side is observed. In addition clocking investigations were carried out to distinguish between the different periodic influences from the surrounding blade rows. For this reason the unsteady profile pressure distribution on rotor 3 was measured, while stator 1–4 were separately traversed stepwise in the circumferential direction. Thus the wake and potential effects of the up- and downstream blade rows on the unsteady profile pressure could clearly be distinguished and quantified.© 2004 ASME

Effects of Auxiliary Lift and Propulsion on Helicopter Vibration Reduction and Trim

Abstract: This paper examines the vibration reductions caused by the introduction of auxiliary lift and propulsion, individually, as well as in combination, on a light [5800-lb (2640 kg)] helicopter with a four-bladed hingeless rotor, at flight speeds close to the maximum cruise velocity of the baseline helicopter. The changes in trim (vehicle orientations and control settings) because of auxiliary lift and propulsion are also examined in detail, and the fundamental mechanisms that produce the changes in trim and associated vibration reductions are identified. Based on results using a comprehensive aeroelastic analysis, it was concluded that auxiliary lift, alone, produces relatively small reductions in vibration. On the other hand, significant vibration reductions were obtained through auxiliary propulsion alone. A combination of lift and propulsion was most effective and reduced the vibration index by over 90%. It was also observed that auxiliary lift significantly reduces the main rotor thrust but increases the nose-down pitch attitude and tip-path-plane forward tilt to provide the required propulsive force. This increases the downwash through the rotor disk and requires a larger rotor longitudinal cyclic pitch input. In contrast, auxiliary propulsion that minimizes vibration produces little reduction in main rotor thrust, but results in a slightly nose-up pitch attitude (the auxiliary propulsion exceeds vehicle drag) along with a backward tilt of the tip-path plane. This decreases the downwash through the rotor disk and requires a smaller rotor longitudinal cyclic pitch input. A combination of auxiliary lift and propulsion minimizes vibration results in an even larger backward tilt of the tip-path plane and a net upwash through the rotor disk. The rotor collective pitch undergoes little change as a result of auxiliary lift, even though the main rotor thrust is decreased. In contrast, for auxiliary propulsion it decreases significantly even though the rotor thrust undergoes only small reductions. This counterintuitive observation is explained. The reduced downwash with auxiliary propulsion, or upwash with combined lift and propulsion, puts the rotor in a partial autorotation state, drastically reducing the induced drag, main rotor torque, and power. Auxiliary lift produces modest reductions in main rotor power, primarily because of a reduced profile drag associated with lower rotor loading. Because the rotor loading is lower with auxiliary lift than with auxiliary propulsion, but larger vibration reductions are produced with the latter, it can be deduced that vibration reductions are less a result of “unloading” of the rotor per se and more because of overall changes in trim, especially the reduction in longitudinal cyclic pitch (seen with auxiliary propulsion).

Unsteady Load of Partial Admission Control Stage Rotor of a Large Power Steam Turbine

Abstract: Partial admission flow in the control stage of a 200MW steam turbine is investigated with the help of a RANS solver with k-ω SST turbulence model in the code Fluent. A 2D model of flow at the mid-span section of the full annulus is assumed. The results exhibit interesting details of the process of expansion in the control stage. Unsteady forces acting on the single rotor blades of the control stage are calculated, and are subject to Fourier analysis. Single blade forces are summed up to obtain the unsteady load at the rotor (forces acting at the rotor disc are neglected due to the assumed 2D model). The calculations take into account pressure pulsations at the entry to the nozzle boxes and rotor blade mistuning / geometrical imperfections.Copyright © 2004 by ASME

Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor—Part II: Unsteady Aerodynamic Forces of Rotor and Stator Blades

Abstract: This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage Low-Speed Research Compressor of Dresden. In part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are determined from the measured profile pressure distributions. A method to calculate the aerodynamic blade forces on the basis of the experimental data is presented. The resulting aerodynamic blade forces are discussed for the rotor and stator blade rows of the first stage and the third stage of the compressor. Different operating points between design point and stability limit of the compressor were chosen to investigate the influence of loading on the aerodynamic force excitation. The time traces and the frequency contents of the unsteady aerodynamic blade force are discussed. Strong periodic influences of the incoming wakes and of potential effects of downstream blade rows can be observed. The amplitude and shape of the unsteady aerodynamic blade force depend on the interaction of the superimposed influences of the blade rows.© 2004 ASME

Optimum Design Framework for Integrally Twisted Helicopter Blades

Abstract: This paper presents the initial development of an optimization framework for designing active twist helicopter rotor blade cross sections with embedded anisotropic piezocomposite actuators. Optimum design of active twist blades is a complex task, since it involves a rich design space with tightly coupled design variables, e.g., the simple orientation of the actuators aects the blade natural frequencies. Therefore, it becomes advantageous to apply the principle of mathematical optimization to the design task. In the proposed framework, the blade cross-sectional internal layout is designed to maximize the static twist actuation while satisfying a series of blade requirements. These requirements are associated with locations of the center of gravity and elastic axis, blade mass per unit span, fundamental rotating blade frequencies, and the blade strength based on local threedimensional stress and strain fields under worst loading conditions. An active composite cross-sectional analysis and a geometrically exact one-dimensional beam analysis, along with other related analysis routines, are combined with a gradient-based optimizer within MATLAB. The developed optimization framework is exemplified by using the NASA/Army/MIT Active Twist Rotor blade and its baseline design.

Wind turbine rotor blade with wing tip

Abstract: The rotor blade (2) is provided at its free end section (1) with an end projection (7) having an aerodynamic cross-sectional profile, which lies in a plane (10) extending at an angle to the rotor blade plane (9). The end projection is asymmetric to the central longitudinal axis of the rotor blade, with a progressive or stepped reduction in the blade thickness at the transition between the end projection and the remainder of the rotor blade.

Understanding Ducted Rotor Antitorque and Directional Control Characteristics Part II: Unsteady Simulations

Abstract: Understanding the dynamic relationship between the antitorque thrust moment and the applied collective pitch angle is crucial, especially for directional control sensitivity analyses. Although there are many studies in the literature on the steady-state behavior of the FANTAIL TM , little is known about the transient response and thrust buildup, which is the primary focus of this paper. Computational fluid dynamics is used for the solutions here because it provides a more complete flowfield prediction, especially in low-power, near edgewise conditions. The flowfield is assumed to be inviscid, and the Euler equations are solved with a blade-element model for the FANTAIL. The main rotor is excluded in this study. Solutions are obtained by modifying the computer code PUMA2 (Parallel Unstructured Maritime Aerodynamics) and using an unstructured grid of 2.8 million cells. The code was run on Beowulf PC clusters. Dynamic fan thrust and moment response to applied collective pitch in hover and forward flight are presented and discussed. Nomenclature Cp = pressure coefficient L = total length of the helicopter N =y awing moment Tfan =f an thrust t = time V = freestream velocity v � =a verage induced velocity y = helicopter spanwise station θ.75 = collective pitch angle

Transient Helicopter Rotor Wakes in Response to Time-Dependent Blade Pitch Inputs

Abstract: A time-accurate, free-vortex method was used to predict the evolution of a helicopter rotor wake and the corresponding unsteady rotor airloads in response to time-varying changes in blade pitch. Both steady and maneuvering flight conditions were examined. The modeling was validated using measured rotor responses to transient increases in collective pitch and also for oscillatory collective and cyclic blade pitch inputs. In each case the results showed that there was a temporal lag in the growth and convection of vorticity into the rotor wake, causing significant unsteady effects at the rotor. For transient blade pitch inputs the calculated results showed the bundling of individual vortex filaments below the rotor into vortex rings, a result also verified experimentally. These vortex rings, however, subsequently break down through the development of Kelvin waves. A simulated piloted pull-up maneuver from descending flight was studied, producing evidence that maneuvers can also cause wake vorticity to bundle below the rotor. Large unsteady rotor airloads were produced as the blades encountered this accumulated wake vorticity

Dynamic modelling and stability of hingeless helicopter blades with a smart spring

Abstract: The aeroelastic stability of a uniform, untwisted hingeless ‘smart’ helicopter rotor blade in hover has been analysed. The concept of a ‘smart’ blade is achieved by implementing a piezoelectric stack at an appropriate location along a host blade such that upon actuation it enters the load path becoming an integral part of the host structure. Thus, the stiffness characteristics of the rotor are altered causing modal damping augmentation of the blade. The perturbation equations of motion for the ‘smart’ blade that describe the unsteady blade motion about the equilibrium operating condition are obtained using Galerkin’s method. These differential equations with periodic time coefficients are analysed for stability utilising the Floquet method. Six different regimes of actuation are investigated, and a parametric study is carried out by considering six different design cases. It is shown that, compared to a ‘host’ blade the stability characteristics of the ‘smart’ blade are not affected adversely. In fact, a judicious design and actuation of the ‘smart’ spring has the potential of improving the stability boundaries of individual blades.

Autonomous Autorotation Landing of Small Unmanned Helicopter

Abstract: In this paper, we derive a linear and a nonlinear model of helicopter autorotation maneuver focusing on main rotor revolution and decent rate dynamics. The nonlinear model for above dynamics is derived from Blade Element Theory combined with the induced velocity linear approximation proposed in this paper. The linear model is derived by implementing a neural network optimization method based on experimental data. The derived nonlinear and linear models are verified by comparing simulated data of the model with experimental one. Finally we carry out preliminary experiment of autorotation landing with a simple PI controller. Rotor revolution and decent rate are found to be well controlled by the designed controller.

Rotor blade connection

Abstract: The invention relates to a rotor blade for a wind power plant provided with several holes which are embodied in the base thereof and designed in the form of passage holes extending essentially vertically with respect to the longitudinal axis of said rotor blade which also comprises transversal bolts inserted into said holes and strength members connectable to the transversal bolts. Usually, the strength members extend in the base of a rotor blade thereby weakening material in said area. The aim of said invention is to solve the problem by simplifying the structure. For this purpose, the inventive rotor blade is characterised in that the strength members extends outwards the rotor blade base in such a way that said area is aerodynamically modified in a disadvantageous manner at a hub area without detrimentally affecting the acoustic behaviour and other properties of the device since the blade base is covered by a blade spinner or arranged at least in the low-rotating rotor area.

On the Formation and Decay of Rotor Blade Tip Vortices

Abstract: Models for the rate of decay of rotor tip vortices, based on near-wake data and computations, are in conflict with the observed persistence of rotor tip vortices for many revolutions. A correlation of data from the literature shows that the measured tip vortex strength is only 40% of the strength expected from the peak of the blade bound circulation, for blades with sharply-cut straight-edged tips, over a wide range of tip Mach and Reynolds numbers. Explanations of this drop in the near wake through turbulence models, are seen to be both counter-intuitive and in conflict with measurements of turbulent fluctuations. Using flow visualization and velocimetry in the wake of a 2-bladed rotor in low-speed forward flight, explanations are shown for these mysteries. Shear layer development at the blade tip, and the twist distribution of the blade have a strong influence on the fraction of the blade tip vorticity lost into the inboard counter-rotating wake. The observations of elliptic core shape and of multiple local extrema of axial velocity deficit, are shown to be features of the shear-layer separation and rollup processes. Blade-to-blade differences in vortex evolution from the same rotor are traced to surface texture differences at the blade tip. Results in the literature reporting 90 t0 100% tip circulation recovery for such blades, are seen to be in cases where the induced velocity due to the tip vortex could not be distinguished from that due to the inboard sheet.

Rotor blade conncection

Abstract: The present invention concerns a rotor blade for a wind power installation comprising a plurality of holes which are arranged in the region of the rotor blade root and which are in the form of through holes which extend substantially transversely with respect to the longitudinal axis of the rotor blade, transverse pins which are fitted into the holes and tension elements which can be connected to the transverse pins. In that arrangement the tension elements in the state of the art extend within the rotor blade root and thus weaken the material in that region. In order to eliminate those disadvantages by means of a structural simplification, the rotor blade according to the invention is characterized by tension elements extending outside the rotor blade root. In that respect the invention is based on the realization that in that way the region of the rotor blade root at the hub is admittedly altered in an aerodynamically disadvantageous fashion, but that does not have any detrimental influence on the acoustic characteristics and the other properties of the installation because that region of the rotor blade is either covered by the spinner or is disposed at least in the part of the rotor, which rotates most slowly.

Blade for wind power generator and wind power generation device having the same

Abstract: PROBLEM TO BE SOLVED: To provide an excellent output controllability excellent in maintainability and low in cost. SOLUTION: A plurality of blades 15 are fixed to the hub 16 of the main shaft of a generator. Each of the blades 15 comprises a blade body 15a having an aerodynamic center Xa extending linearly in a rotating radial direction and a backward blade part 15b having an aerodynamic center Xb extending from the tip part of the blade body 15a to the rear side in a rotating direction. By producing a torsional force in the blade body 15a by a pressing force Fo produced on the backward blade part 15b by wind in a rotating axis direction, a pitch angle α can be adjusted. COPYRIGHT: (C)2006,JPO&NCIPI

Helicopter Vibration Reduction and Damping Enhancement Using Individual Blade Control

Abstract: A control law is presented for a BO-105 helicopter to reduce vibration and to increase damping by the use of individual blade control. H∞ control synthesis is used to develop a robust controller usable in different operating conditions with different helicopter flight speeds. In simulation, hub load vibration can be canceled (−99%) in three outputs simultaneously, for example in all three hub forces. A reduction in hub vibration, however, does not necessarily lead to reduced vibration in the cabin. Therefore, a finite element model of the flexible fuselage is coupled with the aeromechanical rotor model. The resulting coupled rotor‐fuselage model allows vibration to be calculated and controlled at locations in the cabin. A simultaneous vibration reduction of − −89% is achieved at the pilot and copilot seats. To reduce gust sensitivity, lag damping must be enhanced, which requires the lag rate of the blades to be fed back. However, the control law is developed with the constraint of no sensors and, consequently, no measurements in the rotating blades. The use of a model-based control strategy enables lag damping to be enhanced from 0.5 to 2‐3% critical damping by feedback of the observed lag rates, requiring only measurements of the hub loads.

Composite Fan Blade Design for Advanced Engine Concepts

Abstract: The aerodynamic and structural viability of composite fan blades of the revolutionary Exo-Skeletal engine are assessed for an advanced subsonic mission using the NASA EST/BEST computational simulation system. The Exo-Skeletal Engine (ESE) calls for the elimination of the shafts and disks completely from the engine center and the attachment of the rotor blades in spanwise compression to a rotating casing. The fan rotor overall adiabatic efficiency obtained from aerodynamic analysis is estimated at 91.6 percent. The flow is supersonic near the blade leading edge but quickly transitions into a subsonic flow without any turbulent boundary layer separation on the blade. The structural evaluation of the composite fan blade indicates that the blade would buckle at a rotor speed that is 3.5 times the design speed of 2000 rpm. The progressive damage analysis of the composite fan blade shows that ply damage is initiated at a speed of 4870 rpm while blade fracture takes place at 7640 rpm. This paper describes and discusses the results for the composite blade that are obtained from aerodynamic, displacement, stress, buckling, modal, and progressive damage analyses. It will be demonstrated that a computational simulation capability is readily available to evaluate new and revolutionary technology such as the ESE.

Design of a Transonic High-Pressure Turbine Stage 2D Section With Reduced Rotor/Stator Interaction

Abstract: In transonic turbine stages, the exit static pressure field of the vane is highly non-uniform in the pitchwise direction. The rotor traverses periodically this non-uniform field and large static pressure fluctuations are observed around the rotor section. As a consequence the rotor blade is submitted to significant variations of its aerodynamic force. This contributes to the high cycle fatigue and may result in unexpected blade failure. In this paper an existing transonic turbine stage section is redesigned in the view of reducing the rotor stator interaction, and in particular the unsteady rotor blade forcing. The first step is the redesign of the stator blade profile to reduce the stator exit pitchwise static pressure gradient. For this purpose, a procedure using a genetic algorithm and an artificial neural network is used. Next, two new rotor profiles are designed and analysed with a quasi 3D Euler unsteady solver in order to investigate their receptivity to the shock interaction. One of the new profiles allows reducing the blade force variation by 50%.Copyright © 2004 by ASME

Characteristics of Rotary Wings in Hovering Mode at an Ultra-Low Reynolds Number

Abstract: Thrust and torque generated by a model rotary wing were measured at an ultra-low Reynolds number, Re=4×103, for various aspect ratios with and without linear blade twist. The measured characteristics were compared with those calculated by the method which is well known to be effective for analyzing a rotary wing at a high Reynolds number. The method combines annular momentum theory and blade element theory. This calculation method can give the quantitative explanation of the effects of the aspect ratio and of linear blade twist on the characteristics of the rotary wings. The calculation results also indicate that the present calculation method has the capability of giving an accurate quantitative estimation of rotary wing performance with the blade aspect ratio larger than 10, operating at the ultra-low Reynolds number.