Showing papers on "Airfoil published in 2000"

Principles of Helicopter Aerodynamics

Abstract: Preface Acknowledgements List of main symbols List of figures List of tables 1. Introduction: a history of helicopter flight 2. Fundamentals of rotor aerodynamics 3. Blade element analysis 4. Rotating blade motion 5. Basic helicopter performance 6. Conceptual design of helicopters 7. Rotor airfoil aerodynamics 8. Unsteady aerodynamics 9. Dynamic stall 10. Rotor wakes and tip vortices Appendix Index.

Low Reynolds Number Aerodynamics of Low-Aspect-Ratio, Thin/Flat/Cambered-Plate Wings

Abstract: The design of micro aerial vehicles requires a better understanding of the aerodynamics of small low-aspect-ratio wings An experimental investigation has focused on measuring the lift, drag, and pitching moment about the quarter chord on a series of thin flat plates and cambered plates at chord Reynolds numbers varying between 60,000 and 200,000 Results show that the cambered plates offer better aerodynamic characteristics and performance It also appears that the trailing-edge geometry of the wings and the turbulence intensity in the wind tunnel do not have a strong effect on the lift and drag for thin wings at low Reynolds numbers Moreover, the results did not show the presence of any hysteresis, which is usually observed with thick airfoils/wings

Proper Orthogonal Decomposition Technique for Transonic Unsteady Aerodynamic Flows

Abstract: A new method for constructing reduced-order models (ROM) of unsteady small-disturbance e ows is presented. The reduced-order models are constructed using basis vectors determined from the proper orthogonal decomposition (POD) of an ensemble of small-disturbance frequency-domain solutions. Each of the individual frequencydomain solutions is computed using an efe cient time-linearized e ow solver. We show that reduced-order models can be constructed using just a handful of POD basis vectors, producing low-order but highly accurate models of the unsteady e ow over a wide range of frequencies. We apply the POD/ROM technique to compute the unsteady aerodynamic and aeroelastic behavior of an isolated transonic airfoil and to a two-dimensional cascade of airfoils.

Aerodynamics of Gurney Flaps on a Single-Element High-Lift Wing

Abstract: The trailing-edge region of a single-element wing fitted with Gurney flaps has been studied. Measurements include surface pressure, force, and velocity by laser Doppler anemometry (LDA). The mean-velocity vectors and streamlines suggest a twin vortex structure downstream of the Gurney flap. Spectral analysis of the LDA data indicates that the wake consists of a von Karman vortex street of alternately shed vortices, and this flow structure is confirmed by smoke visualization of the flow downstreamof the Gurney flap. The vortex shedding increases the trailing-edge suction of the aerofoil, whereas the upstream face of the device decelerates the flow at the trailing edge of the pressure surface. These two changes result in a pressure difference acting across the trailing edge, and it is this that generates the increase in circulation.

Investigating Three-Dimensional and Rotational Effects on Wind Turbine Blades by Means of a Quasi-3D Navier-Stokes Solver

Abstract: Three-dimensional and rotational viscous effects on wind turbine blades are investigated by means of a quasi-3D Navier-Stokes model. The governing equations of the model are derived from the 3-D primitive variable Navier-Stokes equations written in cylindrical coordinates in the rotating frame of reference. The latter are integrated along the radial direction and certain assumptions are made for the mean values of the radial derivatives. The validity of these assumptions is cross-checked through fully 3-D Navier-Stokes calculations. The resulting quasi-3D model suggests that three-dimensional and rotational effects be strongly related to the local chord by radii ratio and the twist angle. The equations of the model are numerically integrated by means of a pressure correction algorithm. Both laminar and turbulent flow simulations are performed. The former is used for identifying the physical mechanism associated with the 3-D and rotational effects, while the latter for establishing semiempirical correction laws for the load coefficients, based on 2-D airfoil data. Comparing calculated and measured power curves of a stall controlled wind turbine, it is shown that the suggested correction laws may improve significantly the accuracy of the predictions.

Aerodynamic Measurements at Low Raynolds Numbers for Fixed Wing Micro-Air Vehicles

Abstract: : A description of the micro-air vehicle (MAV) concept and design requirements is presented These vehicles are very small and therefore operate at chord Reynolds numbers below 200,000 where very little data is available on the performance of lifting surfaces, ie, airfoils and low aspect-ratio wings This paper presents the results of a continuing study of the methods that can be used to obtain reliable force and moment data on thin wings in wind and water tunnels To this end, a new platform force and moment balance, similar to an already existing balance, was designed and built to perform lift, drag and moment measurements at low Reynolds numbers Balance characteristics and validation data are presented Results show a good agreement between published data and data obtained with the new balance Results for lilt, drag and pitching moment about the quarter chord with the existing aerodynamic balance on a series of thin flat plates and cambered plates at low Reynolds numbers are presented They show that the cambered plates offer better aerodynamic characteristics and performance Moreover, it appears that the trailing-edge geometry of the wings and the turbulence intensity up to about 1% in the wind tunnel do not have a strong effect on the lilt and drag for thin wings at low Reynolds numbers However, the presence of two endplates for two-dimensional tests and one endplate for the semi-infinite tests appears to have an undesirable influence on the lift characteristics at low Reynolds numbers

Airfoil design optimization using reduced order models based on proper orthogonal decomposition

Abstract: This paper presents a method for inviscid airfoil analysis and design optimization that uses reduced order models to reduce the cost of computation. Strong emphasis is placed on obtaining reasonably accurate solutions to the Euler equations with computational costs which are far lower than those required by traditional Computational Fluid Dynamics (CFD) techniques. The design procedure presented here begins by computing a series of flow solutions (snapshots) in which the design variables of interest are perturbed using a Design of Experiments approach. Proper Orthogonal Decomposition (POD) is then used to produce the optimal linear representation of these snapshots using a finite series of basis functions or modes. These basis modes are then used to construct arbitrary solutions to the Euler equations about modified airfoil geometries with very small computational expense. The flow solution problem is reduced in this way to a non-linear least squares fit problem with a small number of variables that can be solved efficiently. For design purposes, a gradient-based optimization procedure is used with the information supplied by the reduced order model. Results for both direct airfoil analysis and for an inverse design optimization problem are presented. Observations regarding the useability of this technique in a design environment are also discussed. Nomenclature aj generic coefficient of the j-th POD mode E total energy (internal plus kinetic) f , g Euler flux vectors H total enthalpy M number of modes used in approximation p static pressure R(x, x) autocorrelation function R autocorrelation tensor, finite-volume residual R autocorrelation matrix for method of snapshots u x-component of velocity v y-component of velocity u arbitrary function to be generated x vector of independent variables λ Lagrange multiplier, an eigenvalue ηi coefficient of the ith mode in a function expansion Ω domain of interest ρ density ϕ j (x) j-th POD basis mode

A proof of existence of perturbed steady transonic shocks via a free boundary problem

Abstract: We prove the existence of a solution of a free boundary problem for the transonic small-disturbance equation The free boundary is the position of a transonic shock dividing two regions of smooth flow Assuming inviscid, irrotational flow, as modeled by the transonic small-disturbance equation, the equation is hyperbolic upstream where the flow is supersonic, and elliptic in the downstream subsonic region To study the stability of a uniform planar transonic shock, we consider perturbation by a steady C1+ϵ upstream disturbance If the upstream disturbance is small in a C1 sense, then there is a steady solution in which the downstream flow and the transonic shock are Holder-continuous perturbations of the uniform configuration This result provides a new use of inviscid perturbation techniques to demonstrate, in two dimensions, the stability of transonic shock waves of the type that appear, for example, over the wing of an airplane, along an airfoil, or as bow shocks in a flow with a supersonic free-stream velocity © 2000 John Wiley & Sons, Inc

Computational Study of Flapping Airfoil Aerodynamics

Abstract: Unsteady, viscous, low-speed e ows over a NACA 0012 airfoil oscillated in plungeand/orpitch at various reduced frequency,amplitude, andphaseshift arecomputed. Vortical wakeformations, boundary-layere owsat theleading edge, the formation of leading-edge vortices and their downstream convection are presented in terms of unsteady particletraces.Flowseparationcharacteristicsandthrust-producingwakeproe lesareidentie ed.Computedresults compare well with water tunnel e ow visualization and force data and other computational data. The maximum propulsive efe ciency is obtained for cases where the e ow remains mostly attached over the airfoil oscillated in a combined pitch and plunge.

Recent Turbulence Model Advances Applied to Multi-Element Airfoil Computations

Abstract: A one-equation linear turbulence model and two-equation nonlinear explicit algebraic stress model (EASM) are applied to the flow over a multielement airfoil. The effect of the K-epsilon and K-omega forms of the two-equation model are explored, and the K-epsilon form is shown to be deficient in the wall-bounded regions of adverse pressure gradient flows. A new K-omega form of EASM is introduced. Nonlinear terms present in EASM are shown to improve predictions of turbulent shear stress behind the trailing edge of the main element and near midflap. Curvature corrections are applied to both the one- and two-equation turbulence models and yield only relatively small local differences in the flap region, where the flow field undergoes the greatest curvature. Predictions of maximum lifts are essentially unaffected by the turbulence model variations studied.

Measurements in Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions

Abstract: Detailed velocity measurements were made along a flat plate subject to the same dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil. Reynolds numbers based on wetted plate length and nominal exit velocity were varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low and high inlet free-stream turbulence intensities (0.2% and 7%) were set using passive grids. The location of boundary-layer separation does not depend strongly on the free-stream turbulence level or Reynolds number, as long as the boundary layer remains non-turbulent prior to separation. Strong acceleration prevents transition on the upstream part of the plate in all cases. Both free-stream turbulence and Reynolds number have strong effects on transition in the adverse pressure gradient region. Under low free-stream turbulence conditions transition is induced by instability waves in the shear layer of the separation bubble. Reattachment generally occurs at the transition start. At Re = 50,000 the separation bubble does not close before the trailing edge of the modeled airfoil. At higher Re, transition moves upstream, and the boundary layer reattaches. With high free-stream turbulence levels, transition appears to occur in a bypass mode, similar to that in attached boundary layers. Transition moves upstream, resulting in shorter separation regions. At Re above 200,000, transition begins before separation. Mean velocity, turbulence and intermittency profiles are presented.

An accurate four-quadrant nonlinear dynamical model for marine thrusters: theory and experimental validation

Abstract: This paper reports two specific improvements in the finite-dimensional nonlinear dynamical modeling of marine thrusters. Previously reported four-quadrant models have employed thin airfoil theory considering only axial fluid flow and using sinusoidal lift/drag curves. First, we present a thruster model incorporating the effects of rotational fluid velocity and inertia on thruster response. Second, we report a novel method for experimentally determining nonsinusoidal lift/drag curves. The model parameters are identified using experimental thruster data (force, torque, and fluid velocity). The models are evaluated by comparing experimental performance data with numerical model simulations. The data indicates that thruster models incorporating both reported enhancements provide superior accuracy in both transient and steady-state responses.

Trailing-Edge Noise Prediction Using Large-Eddy Simulation and Acoustic Analogy

Abstract: The filtered Navier-Stokes equations are used to perform the large-eddy simulation of the unsteady incompressible flow around the blunt trailing edge of a thick flat plate. The computed flow exhibits a three-dimensional vortex shedding mechanism. The frequency domain of this mechanism is in agreement with experiments and theory. In that frequency domain normalized wall-pressure levels are favorably compared to spectra measured at the blunted trailing edge of an airfoil. The far-field radiated noise is first computed via Curle's formulation, the solution of Lighthill's equation for flows embedding solid bodies. Then the theory developed by Ffowcs Williams and Hall is considered. This formulation expresses the noise generated by turbulence passing over the edge of an infinite half-plane. The suitability of this theory to the case for a thick plate is discussed. The normalized spectra of the radiated noise predicted by both methods are compared, in the frequency domain of the vortex-shedding mechanism, to airfoil noise measurements in an anechoic facility.

Large vertical-axis variable-pitch wind turbine

Abstract: Several improvements are invented upon a known and well tested selfstarting vertical-axis wind turbine, for economically competitive power production by driving large grid-connected AC generators. It comprises: variable blade pitch-angle from 0 to 60 degrees, following variable wind speed for maximum efficiency and to keep constant turbine speed, variable blade camber limited to optimal lift-to-drag ratio, controled by pitch and cyclical variation of incidence-angle; improved airfoil shape of cambered blades; low cost automatic gear-train for two constant turbine speeds; protection against overload and prevention of power surge during wind gusts; low stress three-legged high tower assembled with nacelle and tail structure on ground level and erected by assembling segments of third leg one section at a time, and reducing costs of concrete footing, tower structure and, hoisting and assembling at ground level, enables a tower to be built to any height required to harness maximum wind energy.

Shaped cooling hole for an airfoil

Abstract: A cooling hole configuration for an air-cooled component, such as a gas turbine engine airfoil. The cooling hole is configured to have cross-sectional variations and a noncircular-shaped diffuser-type opening that significantly improve the cooling film distribution across the external surface of an airfoil, with the result that heat transfer from the surrounding environment to the airfoil is reduced. The cooling hole is configured to have its central axis at an acute angle to the exterior surface of the airfoil, and defines a noncircular-shaped opening at the airfoil surface. The cooling hole generally has a first region adjacent the airfoil surface and a second region interior to the airfoil. The cooling hole is configured such that the second region has an oblong or oval-shaped cross-section, with a major diameter approximately equal to the major diameter of the first region and a minor diameter less than the major diameter of the first region, with a smooth transition existing between the first and second regions. The oblong shape of the second region is preferably the result of a recess being present in the wall of the cooling hole opposite the direction in which the cooling hole extends toward the airfoil surface, causing the central axis of the cooling hole to have an arcuate shape in which the central axis is disposed at a lesser angle to the airfoil surface in the first region than the angle in the second region. The cooling hole can be generated using a water jet technique.

Turbine blade for gas turbine engine and method of cooling same

Abstract: A turbine blade includes a platform and an airfoil extending radially from the platform. An internal cooling circuit is formed in the airfoil for circulating a coolant therethrough to cool the airfoil. At least one supply passage is provided to direct some of the coolant that has passed at least partially through the internal cooling circuit onto the platform for cooling the platform.

Navier-Stokes Airfoil Computations with e Transition Prediction Including Transitional Flow Regions

Abstract: The e ow around laminar airfoils is computed using a Navier ‐Stokes method coupled to a transition prediction method based on thee N approach. Applying point transition at the predicted transition location produces a strong viscous/inviscid interaction region that prevents the coupled system from converging, whereas the introduction of transitional e ow regions resolves that problem. The emphasis is not placed on the development of new transitional e ow models but primarily on producing convergence, applying modie ed, available models. A comprehensive computational study is performed in a strong adverse and zero pressure gradient airfoil e ow region, as the transitional lengths differ considerably for the different models. A conventional model, which is applicable in e ow regions wheretransition is predicted well upstream of laminar separation, is proposed, together with a special transitional length model fore owswheretheboundary layerstays laminarup to separation. Thee owsovertheDoAL3 and the NLF(1)-0416 laminar airfoils are investigated. The coupled Navier ‐Stokes and e N methods are shown to produce converged results. Furthermore, the values for lift and drag are in excellent agreement with the free transition measurements.

Aerodynamic and Aeroacoustic Optimization of Rotorcraft Airfoils via a Parallel Genetic Algorithm

Abstract: A parallel genetic algorithm (GA) methodology was developed to generate a family of two-dimensional airfoil designs that address rotorcraft aerodynamic and aeroacoustic concerns The GA operated on 20 design variables, whichconstitutedthecontrolpointsforasplinerepresentingtheairfoilsurfaceTheGAtookadvantageofavailable computer resources by operating in either serial mode, where the GA and function evaluations were run on the same processor or “ manager/worker” parallel mode, where the GA runs on the manager processor and function evaluations areconducted independently on separate workerprocessors The multiple objectives of this work were to minimizethedrag and overall noiseof the airfoil Constraintswereplaced on liftcoefe cient, moment coefe cient, andboundary-layerconvergenceTheaerodynamicanalysiscodeXFOILprovidedpressureandsheardistributions in addition to liftand drag predictions Theaeroacousticanalysis code, WOPWOP, provided thicknessand loading noise predictions The airfoils comprising the resulting Pareto-optimal set exhibited favorable performance when compared with typical rotorcraft airfoils under identical design conditions using the same analysis routines The relationship between the quality of results and the analyses used in the optimization is also discussed The new airfoil shapes could provide starting points for further investigation

A Time-Linearized Navier–Stokes Analysis of Stall Flutter

Abstract: A computational method for predicting unsteady viscous flow through two-dimensional cascades accurately and efficiently is presented. The method is intended to predict the onset of the aeroelastic phenomenon of stall flutter. In stall flutter, viscous effects significantly impact the aeroelastic stability of a cascade. In the present effort, the unsteady flow is modeled using a time-linearized Navier-Stokes analysis. Thus, the unsteady flow field is decomposed into a nonlinear spatially varying mean flow plus a small-perturbation harmonically varying unsteady flow. The resulting equations that govern the perturbation flow are linear, variable coefficient partial differential equations. These equations are discretized on a deforming, multiblock, computational mesh and solved using a finite-volume Lax-Wendroff integration scheme. Numerical modeling issues relevant to the development of the unsteady aerodynamic analysis, including turbulence modeling, are discussed. Results from the present method are compared to experimental stall flutter data, and to a nonlinear time-domain Navier-Stokes analysis. The results presented demonstrate the ability of the present time-linearized analysis to model accurately the unsteady aerodynamics associated with turbomachinery stall flutter.

Aerodynamic and aeroacoustic optimization of rotorcraft airfoils via a parallel genetic algorithm

Abstract: A parallel genetic algorithm (GA) methodology was developed to generate a family of two-dimensional airfoil designs that address rotorcraft aerodynamic and aeroacoustic concerns. The GA operated on 20 design variables, which constituted the control points for a spline representing the airfoil surface. The GA took advantage of available computer resources by operating in either serial mode, where the GA and function evaluations were run on the same processor or manager/worker parallel mode, where the GA runs on the manager processor and function evaluations are conducted independently on separate worker processors. The multiple objectives of this work were to minimize the drag and overall noise of the airfoil. Constraints were placed on lift coefficient, moment coefficient, and boundary-layer convergence. The aerodynamic analysis code XFOIL provided pressure and shear distributions in addition to lift and drag predictions. The aeroacoustic analysis code, WOPWOP, provided thickness and loading noise predictions. The airfoils comprising the resulting Pareto-optimal set exhibited favorable performance when compared with typical rotorcraft airfoils under identical design conditions using the same analysis routines

Using Sweep and Dihedral to Control Three-Dimensional Flow in Transonic Stators of Axial Compressors

Abstract: The paper describes an advanced three-dimensional blading concept for highly loaded transonic compressor stators. The concept takes advantage of the aerodynamic effects of sweep and dihedral. To the knowledge of the authors this is the first approach reported in the open literature that combines those two basic types of lean in an engine-worthy aerofoil design. The paper makes a contribution to the understanding of the endwall effect of both features with special emphasis put on sweep. The advanced three-dimensional blading concept was applied to an Engine Section Stator (ESS) of an aero-engine fan. In order to demonstrate how three-dimensional flow can be controlled, numerical analysis of the flow structure in a conventional and an advanced stator configuration was performed using a three-dimensional Navier-Stokes solver. The numerical analysis showed the advanced blade improving both radial loading distribution and the three-dimensional endwall boundary layer development. In particular, a strong hub corner stall could be largely alleviated. High-speed rig testing of the advanced ESS confirmed the concept and showed good qualitative agreement between measurement and prediction. The work presented was closely linked to the development of the BR710 engine on which the advanced ESS is in service today.

Double bowed compressor airfoil

Abstract: A compressor airfoil (12) includes pressure and suction sides (18,20) extending from root (22) to tip (24) and between leading and trailing edges (26,28). Transverse sections have respective chords and camber lines. Centers of gravity (34) of the sections are aligned along a double bowed stacking axis for improving performance.

Effects of Reynolds Number and free-stream Turbulence on Boundary Layer Transition in a Compressor Cascade

Abstract: An experimental and analytical study has been performed on the effect of Reynolds number and free-stream turbulence on boundary layer transition location on the suction surface of a controlled diffusion airfoil (CDA). The experiments were conducted in a rectilinear cascade facility at Reynolds numbers between 0.7 and 3.0 310 6 and turbulence intensities from about 0.7 to 4 percent. An oil streak technique and liquid crystal coatings were used to visualize the boundary layer state. For small turbulence levels and all Reynolds numbers tested, the accelerated front portion of the blade is laminar and transition occurs within a laminar separation bubble shortly after the maximum velocity near 35‐40 percent of chord. For high turbulence levels (Tu .3 percent) and high Reynolds numbers, the transition region moves upstream into the accelerated front portion of the CDA blade. For those conditions, the sensitivity to surface roughness increases considerably; at Tu54 percent, bypass transition is observed near 7 ‐10 percent of chord. Experimental results are compared to theoretical predictions using the transition model, which is implemented in the MISES code of Youngren and Drela. Overall, the results indicate that early bypass transition at high turbulence levels must alter the profile velocity distribution for compressor blades that are designed and optimized for high Reynolds numbers. @DOI: 10.1115/1.1413471#

Prediction of Hysteresis Associated with the Static Stall of an Airfoil

Abstract: Introduction T HE basic phenomenon of stall associated with airfoils is quite well understood and has now become standard textbook material. It is caused by massive flow separation resulting in sharp drop in lift and increase in the drag acting on the airfoil. In certain cases, hysteresis in the flow has been observed for angles of attack close to the stall angle. However, this phenomenon is not very well understood. Hoffmannl has reported the hysteresis loop in the data for aerodynamic coefficients for a NACA 0015 airfoil. He also studied the effect of freestream turbulence (FST) on the performance characteristics of the airfoil. The hysteresis in the data can be observed for low FST but disappears for high FST. The present work is an effort to study the behavior of the flow near stall by solving the governing flow equations numerically. Carefully conducted <:omputations are utilized to track the hysteresis loop in the aerodynamic data close to the stall angle. To the best of the knowledge of these authors, this is the first effort of its nature. The incompressible, Reynoldsaveraged Navier-Stokes (RANS) equations, in conjunction with the Baldwin-Lomax moder for turbulence closure are solved using stabilized finite element formulations. The finite element mesh consists of a structured mesh close to the body and an unstructured part, generated via Delaunay's triangulation, away from the body. This type of a grid has the ability of handling fairly complex geometries while still providing the desired resolution close to the body to effectively capture the boundary-layer flow, especially in the context of unsteady flows. Despite the simplicity of the Baldwin-Lomax model, its implementation with unst~ctured grids is not trivial. The interested reader is referred to the articles by Kallinderis3 and Mavriplis4 for details. The finite element formulations and their implementations used in the present work are well proven and have been utilized to

Method for replacing a turbine vane airfoil

Abstract: A method for refurbishing turbine engine vanes in which the airfoils are removed and replaced. The original configuration of the vanes can be modified from a single casting or welded pair to a multi-piece component assembly comprising individual airfoil segments attached to the inner and outer platforms. The component assembly allows replacement of airfoils (8, 10) and/or platforms (4, 6) with improved casings in the form of improved alloys, improved physical geometry, or both. Modifications can be made in the vanes class area without the need to modify the airfoil contours by brazing or other contour alteration processes. According to the method, the platforms (4, 6) are separated from the airfoils (8, 10), the openings in the platforms are sealed, new airfoil sockets are cut into the platforms, and the vane is reassembled.