Showing papers on "Airfoil published in 1998"

Experimental and Computational Investigation of the Knoller-Betz Effect

Abstract: The ability of a sinusoidally plunging airfoil to produce thrust, known as the Knoller-Betz or Katzmayr effect, is investigated experimentally and numerically. Water-tunnel experiments are performed providing flow visualization and laser Doppler velocimetry data of the unsteady wakes formed by the plunging foils. Vortical structures and time-averaged velocity profiles in the wake are compared with numerical computations from a previously developed inviscid, unsteady panel code that utilizes a nonlinear wake model

A 3-D stall-delay model for horizontal axis wind turbine performance prediction

Abstract: Most design and analysis methods widely used for horizontal axis wind turbine performance prediction, such as the PROP code, are based on the traditional 2-D blade element/momentum theory (BEMT) methods, which are inadequate and underpredict the wind turbine rotor power output in the high-wind/peak-power condition, owing to effects of rotation on the wind turbine blade boundary layer. Although the deficiencies of the methods have been known for some time, this area has been neglected. The continued development of viable and well-established stall-regulated wind-turbine technology makes this research topic timely and particularly relevant to reducing the cost of wind energy. The main aim of the present paper is to describe and analyze the fundamental flow phenomena that characterize the boundary layer on rotating blades, and to develop a preliminary stall-delay model that modifies the 2-D airfoil data so as to simulate the 3-D stall-delay effects. The following steps were taken in the development of the model: 1) analysis of the 3-D integral boundarylayer equations for a reference system rotating with the blade, 2) description of the effects of rotor rotation on the separation point and its causes, and 3) determination of a simple correction formula to obtain rotating rotor lift coefficient Ci(a) and drag coefficient Cd(a) data from measured 2-D airfoil data. The preliminary 3-D stall-delay model consists of two key parameters (the ratio of local chord to local radius c/r the ratio of rotation speed to freestream velocity A) and three empirical correction factors Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. and the American Society of Mechanical Engineering. All rights reserved. ^Visiting Scholar. 'Assistant Professor, Senior Member AIAA. (a, b, d}. The stall-delay model is consistent with the blade element/momentum theory method and the Viterna/Tangler model, and the 3-D stall-delay model can be incorporated into the state of the art performance prediction codes, such as PROP. Through comparison with the field test data, the new model for 3-D stall-delay shows good agreement between predictions and experiments. The new model should be of great use in existing codes for horizontal axis wind turbine design and analysis.

Modification of lifting body aerodynamics using synthetic jet actuators

Abstract: The control of separated flow on an unconventional airfoil using synthetic jet actuators was investigated experimentally. A symmetric airfoil based on the aft portion of a NACA four-digit series airfoil with a cylindrical leading edge was used in the experiment. The tests were conducted at Rec=3(10)5. For a>5°, the flow separated from the airfoil surface. Applying synthetic jet control near the leading edge, upstream of the separation point, reattached the separated flow fixangle of attack up to 18°. The effect of control location and amplitude was investigated for different angles of attack. Hot wire measurements in the nearwake of the airfoil revealed a transient passing of vortices associated with the transition from separated to reattached flow on the airfoil.

Wind Turbine Performance Under Icing Conditions

Abstract: The effects of rime ice on horizontal axis wind turbine performance were estimated. For typical supercooled fog conditions found in cold northern regions, four rime ice accretions on the S809 wind turbine airfoil were predicted using the NASA LEWICE code. The resulting airfoil/ice profile combinations were wind tunnel tested to obtain the lift, drag, and pitching moment characteristics over the Reynolds number range 1--2 {times} 10{sup 6}. These data were used in the PROPID wind turbine performance prediction code to predict the effects of rime ice on a 450-kW rated-power, 28.7-m diameter turbine operated under both stall-regulated and variable-speed/variable-pitch modes. Performance losses on the order of 20% were observed for the variable-speed/variable-pitch rotor. For the stall-regulated rotor, however, a relatively small rime ice profile yielded significantly larger performance losses. For a larger 0.08c-long rime ice protrusion, however, the rated peak power was exceeded by 16% because at high angles the rime ice shape acted like a leading edge flap, thereby increasing the airfoil C{sub l,max} and delaying stall.

Use of Piezoelectric Actuators for Airfoil Separation Control

Abstract: Surface-mounted piezoelectric actuators are used to excite the turbulent boundary layer upstream of separation, where the actuators interact directly with the boundary layer. The actuators are rigid and do not attenuate with increased aerodynamic loading up to the maximum tested speed of 30 m/s

Active flow control applied to an airfoil

Abstract: Results of numerical simulations of active flow control applied to an airfoil using the Reynoldsaveraged Navier-Stokes equations are presented. The simulations are first compared with the poststall separation control experiments of Seifert et al.1'12 on a NACA0015 at 1.2 x 106 chord Reynolds number. The jet is introduced tangential to the surface at the leading edge of the airfoil. The calculated lift increments are in good agreement with the experimental data. Two flow-control techniques for a NACA0012 airfoil at a chord Reynolds number of 8.5 x 10 6 are investigated. The first technique utilizes a small, 0.5% chord, steady jet, and the second method employs a synthetic jet of a similar scale. Performance benefits are obtained by placing the actuators very near the airfoil leading edge on the suction surface. A significant increase in lift (29%) is obtained using the synthetic jet actuator in the post-stall regime. At lower lift, the steady jet actuator significantly reduces drag by rotating the lift vector upstream.

New Airfoils for Small Horizontal Axis Wind Turbines

Abstract: In a continuing effort to enhance the performance of small wind energy systems, one root airfoil and three primary airfoils were specifically designed for small horizontal axis wind turbines. These airfoils are intended primarily for 1--5 kW variable-speed wind turbines for both conventional (tapered/twisted) or pultruded blades. The four airfoils were wind-tunnel tested at Reynolds numbers between 100,000 and 500,000. Tests with simulated leading-edge roughness were also conducted. The results indicate that small variable-speed wind turbines should benefit from the use of the new airfoils which provide enhanced lift-to-drag ratio performance as compared with previously existing airfoils.

Gurney Flap Experiments on Airfoils, Wings, and Reflection Plane Model

Abstract: The effect of Gurney e aps on two-dimensional airfoils, three-dimensional wings, and a ree ection plane model were investigated. There have been a number of studies on Gurney e aps in recent years, but these studies have been limited to two-dimensional airfoil sections. A comprehensive investigation on the effect of Gurney e aps for a wide range of cone gurations and test conditions was conducted at Wichita State University. A symmetric NACA 0011 and a cambered GA (W)-2 airfoil were used during the single-element airfoil part of this investigation. The GA (W)-2 airfoil was also used during the two-element airfoil study with a 25% chord slotted e ap dee ected at 10, 20, and 30 deg. Straight and tapered ree ection plane wings with natural laminar e ow (NLF) airfoil sections were tested for the three-dimensional wing part of this investigation. A fuselage and engine were attached to the tapered NLF wing for the ree ection plane model investigation. In all cases the Gurney e ap improved the maximum lift coefe cient compared to the baseline clean cone guration. However, there was a drag penalty associated with this lift increase.

Radiated noise from airfoils in realistic mean flows

Abstract: The long-term objective of the research described is to use computational aeroacoustics methodology and parallel computers to increase the understanding of broadband blade noise. In a systematic progression toward simulations of completely realistic configurations and conditions, some simplified problems that address the important features of the flow are investigated. A two dimensional Navier-Stokes code, implemented using the message passing library and Fortran 90 on the IBM SP2, is used to perform the calculations. Results are presented for the interaction of a vortical gust and NACA airfoils, including nonlinear effects. The influence of gust frequency and airfoil thickness is described. A multigrid method is used to obtain converged steady-state solutions before the gust is introduced in a source region inside the domain.

Methods for repairing and reclassifying gas turbine engine airfoil parts

Abstract: A method for repairing gas turbine engine airfoil parts. The dimensional differences between pre-repaired dimensions of a turbine engine airfoil part and desired post-repair dimensions of the turbine engine airfoil part are determined. A build-up thickness of coating material required to obtain the desired post-repair dimensions of the turbine engine airfoil part is determined. A high-density coating process, such as HVOF, is used to coat the turbine engine airfoil part with a coating material to the determined build-up thickness of coating material effective to obtain the desired post-repair dimensions after performing a sintering heat treatment and a hot isostatic pressing treatment, and, if performed, a re-application of a protective coating. The coating material comprises a metal alloy capable of forming a diffusion bond with the substrate of the turbine engine airfoil part. After the coating material is applied, the sintering heat treatment process is performed to prevent gas entrapment of the coating material and/or the diffusion bonding area during the hot isostatic pressing process. Then, the hot isostatic pressing process is performed to obtain a post-repair turbine engine airfoil part having the desired post-repair dimensions and having diffusion bonding between the coating material and the turbine engine airfoil substrate. A protective coating may be first removed from the turbine engine airfoil part prior to performing the high-density coating process. Typically, this protective coating is present on an airfoil part to protect it from the hot corrosive environment it experiences during service. After performing the hot isostatic pressing process, the protective coating may be re-applied. In this case, the build-up thickness may determined to take into consideration the additional thickness of the post-repaired part due to the addition of the protective coating.

Tapered tip turbine blade

Abstract: A turbine blade includes a hollow airfoil extending from an integral dovetail. The airfoil includes sidewalls extending between leading and trailing edges and longitudinally between a root and a tip. The sidewalls are spaced apart to define a flow channel for channeling cooling air through the airfoil. The tip is tapered longitudinally above at least one of the sidewalls and decreases in thickness.

Aerodynamic control of bridge deck flutter by active surfaces

Abstract: An active aerodynamic control method of suppressing flutter of a very-long-span bridge is presented in this paper. The control system consists of additional control surfaces attached to the bridge deck; their torsional movement, commanded via feedback control law, is used to generate stabilizing aerodynamic forces. The frequency independent formulation of unsteady aerodynamic forces acting on the bridge deck as well as the control surfaces is derived through rational function approximation. The high precision of approximation is ensured by multilevel linear and nonlinear constrained optimization. Although the proposed mathematical model of aeroservoelastic system is augmented by new aerodynamic states, it is in the form of a set of constant coefficient differential equations that are particularly convenient for control law synthesis. The obtained equations of motion are functions of mean wind speed so the efficiency of application of the conventional constant gain optimal feedback control is limited. To cope with the system dependence on wind speed, a variable-gain control is proposed. The static output variable-gain approach is formulated in terms of a mathematical optimization problem and the necessary conditions are derived. Application of the variable-gain control provides variation of control strategies in different wind velocities and is found to be efficient for the studied aerodynamic active control of bridge deck flutter.

Drag Reduction and Shape Optimization of Airship Bodies

Abstract: A tool for the numerical shape optimization of axisymmetric bodies submerged in incompressible flow at zero incidence has been developed. Contrary to the usual approach, the geometry of the body is not optimized in a direct way with this method. Instead, a source distribution on the body axis was chosen to model the body contour and the corresponding inviscid flowfield, with the source strengths being used as design variables for the optimization process. Boundary-layer calculation is performed by means of a proved integral method. To determine the transition location, a semiempirical method based on linear stability theory (e n method) was implemented. A commercially available hybrid optimizer as well as an evolution strategy with covariance matrix adaption of the mutation distribution are applied as optimization algorithms. Shape optimizations of airship hulls were performed for different Reynolds number regimes. The objective was to minimize the drag for a given volume of the envelope and a prescribed airspeed range

Measurement of unsteady boundary layer developed on an oscillating airfoil using multiple hot-film sensors

Abstract: The spatial-temporal progressions of the leading-edge stagnation, separation and reattachment points, and the state of the unsteady boundary layer developed on the upper surface of a 6 in. chord NACA 0012 airfoil model, oscillated sinusoidally within and beyond the static-stall angle, were measured using 140 closely-spaced, multiple hot-film sensors (MHFS). The MHFS measurements show that (i) the laminar separation point and transition were delayed with increasing α and the reattachment and relaminarization were promoted with decreasing α, relative to the static case, (ii) the pitchup motion helped to keep the boundary layer attached to higher angles of attack over that could be obtained statically, (iii) the dynamic stall process was initiated by the turbulent flow separation in the leading-edge region as well as by the onset of flow reversal in the trailing-edge region, and (iv) the dynamic stall process was found not to originate with the bursting of a laminar separation bubble, but with a breakdown of the turbulent boundary layer. The MHFS measurements also show that the flow unsteadiness caused by airfoil motion as well as by the flow disturbances can be detected simultaneously and nonintrusively. The MHFS characterizations of the unsteady boundary layers are useful in the study of unsteady separated flowfields generated by rapidly maneuvering aircraft, helicopter rotor blades, and wing energy machines.

Gyro stabilized triple mode aircraft

Abstract: A triple mode aircraft which can take off as a helicopter, or in gyrocopter mode with no power to the rotors or as a conventional aircraft obtaining lift from a circular wing and in another embodiment from short stub wings; a canard wing and high lift tailplane. So it combines the flexibility of a helicopter with the same efficiency and safety of a gyrocopter and a fixed wing aircraft, also has the same simplicity and efficiency of flying a helicopter that doesn't have a tail rotor to worry about. The rotor craft includes two counter-rotating rotors with weighted tips on one set of rotors or a circular airfoil (CA) attached to at least one set of rotors. This CA gives the lift to function as a conventional aircraft, it also weights the tips of the rotor to give a gyro-stabilizing effect to the whole aircraft. Also there is a down draft rudder that functions as a rudder in horizontal flight or catches the down draft from the rotors for directional control. There is a conventional horizontal tail with elevator controls which can be differentially operated for additional control.

Unsteady Stall Control Using Dynamically Deforming Airfoils

Abstract: Introduction An experiment documenting the compressible flow over a dynamically deforming airfoil is presented. This airfoil, which has a leading edge radius that can be dynamically changed, was tested at various defor- mation rates for fixed airfoils angle of attack. Selected leading-edge shapes were also tested during airfoil os- cillation. These tests show that for a range of Mach numbers observed on the retreating blades of heli- copter rotors the dynamic stall vortex can be avoided by the judicious variation of leading-edge curvature

Dynamic stall and aerodynamic damping

Abstract: A dynamic stall model is used to analyze and reproduce open air blade section measurements as well as wind tunnel measurements. The dynamic stall model takes variations in both angle of attack and flow velocity into account. The paper gives a brief description of the dynamic stall model and presents results from analyses of dynamic stall measurements for a variety of experiments with different airfoils in wind tunnel and on operating rotors. The wind tunnel experiments comprises pitching as well as plunging motion of the airfoils. The dynamic stall model is applied for derivation of aerodynamic damping characteristics for cyclic motion of the airfoils in flapwise and edgewise direction combined with pitching. The investigation reveals that the airfoil dynamic stall characteristics depend on the airfoil shape, and the type of motion (pitch, plunge). The aerodynamic damping characteristics, and thus the sensitivity to stall induced vibrations, depend highly on the relative motion of the airfoil in flapwise and edgewise direction, and on a possibly coupled pitch variation, which is determined by the structural characteristics of the blade.

Experimental Active Control of a Typical Section Using a Trailing-Edge Flap

Abstract: This paper presents an experimental implementation of an active control system used to suppress flutter in a typical section airfoil. The H2 optimal control system design is based on experimental system identifications of the transfer functions between three measured system variables - pitch, plunge, and flap position - and a single control signal that commands the flap of the airfoil. Closed-loop response of the airfoil demonstrated gust alleviation below the open-loop flutter boundary. In addition, the flutter boundary was extended by 12.4% through the application of active control. Cursory robustness tests demonstrate stable control for variations in flow speed of ± 10%.

Trapped insert turbine airfoil

Abstract: A gas turbine engine airfoil is manufactured by forming an internal retention seat in two complementary airfoil parts. An insert is fabricated for retention in the seat. The two parts are assembled with the insert disposed in the seat therebetween. The parts are then bonded together to trap the insert therein to collectively define the airfoil. The insert and seat may be precisely fabricated for improving the efficiency of the airfoil.

Flutter of an airfoil with a cubic nonlinear restoring force

Abstract: In this paper, the effect of a cubic structural restoring force on the flutter characteristics of a twodimensional airfoil placed in an incompressible flow is investigated. The aeroelastic equations of motion are written as a system of eight first-order ordinary differential equations. Given the initial values of plunge and pitch displacements and their velocities, the system of equations is integrated numerically using a 4 order Runge-Kutta scheme. Results for softand hard-springs are presented for a pitch degree-of-freedom nonlinearity. The study shows the dependence of the divergence flutter boundary on initial conditions for a soft spring. For a hard spring, the nonlinear flutter boundary is independent of initial conditions for the spring constants considered. The flutter speed is identical to that for a linear spring. Divergent flutter is not encountered, but instead limit cycle oscillation occurs for velocities greater than the flutter speed. The behaviour of the airfoil is also analyzed using analytical techniques developed for nonlinear dynamical systems. The Hopf-bifurcation point is determined analytically and the amplitude of the limit cycle oscillation in postHopf-bifurcation for a hard spring is predicted using an asymptotic theory. The frequency of the limit cycle *Principal Research Officer and Head, Experimental Aerodynamics and Aeroelasticity Group. Also adjunct professor, Depl. of Mathematical Sciences, University of Alberta. Associate Fellow AIAA. "Research Associate, Experimental Aerodynamics and Aeroelasticity Group. 'Professor, Dept. of Mathematical Sciences. Copyright © 1998 by B.H.K. Lee. L.Y. Jiang and Y.S. Wong, Published by the American Inst i tute of Aeronautics and Astronautics Inc. with permission oscillation is estimated from an approximate method. Comparisons with numerical simulations are carried out and the accuracy of the approximate method is discussed. The analysis can readily be extended to study limit cycle oscillation of airfoils with nonlinear polynomial spring forces in both plunge and pitch degrees of freedom. NOMENCLATURE ah non-dimensional distance from airfoil midchord to elastic axis b airfoil semi-chord CL aerodynamic lift coefficient CM pitching moment coefficient h plunge displacement m airfoil mass R response amplitude of pitch motion r response amplitude of plunge motion ra radius of gyration about elastic axis t time U free stream velocity U non-dimensional velocity, U/bcoa UL non-dimensional linear flutter speed xa non-dimensional distance from elastic axis to centre of mass X system variable vector XE system equilibrium point y variable vector a pitch angle of airfoil OCA pitch angle amplitude of l imit cycle oscillation EI, £2 constants in Wagner's function Pa, p= coefficients of cubic spring in pitch and plunge C,a, viscous damping ratios in pitch and plunge H airfoil/air mass ratio, m/Kpb

A computational study on the dynamic stall of a flapping airfoil

Abstract: The dynamic stall boundaries of a NACA 0012 airfoil oscillating in either the pure plunge mode or in the combined pitch and plunge mode is computed using a thin-layer Navier-Stokes solver. Unsteady flowfields are computed at the free-stream Mach number of 0.3, the Reynolds number of 1 • 10, and the Baldwin-Lomax turbulence model is employed. It is found that the pure plunge oscillation leads to dynamic stall as soon as the non-dimensional plunge velocity exceeds the approximate value of 0.35. In addition, the power extraction capability of the airfoil operating in the wingmill mode is studied by computing the dynamic stall boundary for a combined pitch and plunge motion at the reduced frequency values of 0.1, 0.25 and 0.5.

Investigation of confluent boundary layers in high-lift flows

Abstract: The confluent boundary layers over a three-element high-lift airfoil are studied using both numerical and experimental approaches. The results suggest that wake prediction is crucial to the convergence and accuracy of the overall solution. At maximum lift, unsteadiness is observed in the experiment, which is not captured by computations. However, solutions at maximum lift indicate that, although the flow is attached over the flap, the separation bubble at the leading edge of the slat upper surface is coupled with inviscid flow reaching the compressibility limit. The thickened slat wake results in a displacement of near-surface flow over the main element and limits the main element from gaining more lift. The trends in the confluent boundary layers development require all aspects of the physics be modeled appropriately, including transition, turbulence, and inviscid-viscous interaction

Effects of Zero-Mass “Synthetic” Jets on the Aerodynamics of the NACA-0012 Airfoil

Abstract: A numerical study was conducted to investigate the effect of an array of zero-mass "synthetic" jets on the aerodynamic characteristics of the NACA-0012 airfoil. Flowfield predictions were made using a modified version of the NASA Ames "ARC2D" unsteady, two-dimensional, compressible Navier-Stokes flow solver. An unsteady surface transpiration boundary condition was enforced over a user-specified portion of the airfoil's upper or lower surface to emulate the time variation of the mass flux out from and into the airfoil's surface. Here, a sinusoidal function which describes the approximate time behavior of the mass flux across the airfoil's surface was used. Our numerical results have indicated that zero-mass jets can, with the careful selection of their peak amplitude and frequency, enhance the lift characteristics of airfoils (helicopter rotor blades, wings, etc.). Effects of the jet peak suction and blowing velocities, oscillation frequency and, jet placement on the time histories of the sectional lift, drag and moment are presented for two angles of attack and one free stream Mach number. Flowfield results which provide more insight into the mechanics of the interaction between the array of jets and the developing boundary layer over the airfoil are presented.

Aerodynamic and aero acoustic optimization of airfoils via a parallel genetic algorithm

Abstract: A parallel genetic algorithm (GA) was used to generate, in a single run, a family of aerodynamical ly efficient, low-noise rotor blade designs representing the Pareto optimal set. The n-branch tournament, uniform crossover genetic algorithm operates on twenty design variables, which constitute the control points for a spline representing the airfoil surface. The GA takes advantage of available computer resources by operating in either serial mode or "manager/work er" parallel mode. The multiple objectives of this work were to maximize lift-to-drag of a rotor airfoil shape and to minimize an overall noise measure including effects of loading and thickness noise of the airfoil. Constraints are placed on minimum lift coefficient, pitching moment and boundary layer convergence. The program XFOIL provides the aerodynamic analysis, and the code WOPWOP provides the aeroacoustic analysis. The Pareto-optimal airfoil set has been generated and is compared to the performance of a typical rotorcraft airfoil under identical flight conditions.

Low noise permeable airfoil

Abstract: A stator airfoil is subject to wake fluid from rotor blades in a gas turbine engine. The airfoil includes a plurality of cross channels extending therethrough between opposite sides thereof for bleeding the wake fluid therebetween to reduce differential pressure thereacross and reduce noise.