Showing papers on "Airfoil published in 1995"

Finite state induced flow models. I - Two-dimensional thin airfoil

Abstract: A new finite state aerodynamic theory is presented for incompressible, two-dimensional flow around thin airfoils. The theory is derived directly from potential flow theory with no assumptions on the time history of airfoil motions. The aerodynamic states are the coefficients of a set of induced-flow expansions. As a result, the finite state equations are hierarchical in nature and have closed-form coefficients. Therefore, the model can be taken to as many states as are dictated by the spatial texture and frequency range of interest with no intermediate numerical analysis. The set of first-order state equations is easily coupled with structure and control equations and can be exercised in the frequency or Laplace domain as well as in the time domain. Comparisons are given with Theodorsen theory, Wagner theory, and other methods. Excellent results are found with only a few states.

NREL airfoil families for HAWTs

Abstract: The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub l,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulatedmore » turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.« less

NREL airfoil families for HAWTs

Abstract: The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub l,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulatedmore » turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.« less

Finite state induced flow models. II - Three-dimensional rotor disk

Abstract: In Part I of this two-part article, we developed a finite state induced flow model for a two-dimensional airfoil. In this second part, we develop a finite state induced flow model for the three-dimensional induced flow for a rotor. The coefficients of this model are found in a compact closed form. Although the model does not presuppose anything about the source of lift on the rotating blades, applications are given in which the Prandtl assumption is invoked. That is, the two-dimensional lift equations are used at each radial station, but with the inflow from the three-dimensional model. The results are shown to reduce (in several special cases) to Prandtl-Golds tein theory, Theodorsen theory, Loewy theory, dynamic inflow, and blade-element momentum theory. Comparisons with vortex-filament models and with experimental data in hover and forward flight also show excellent correlation.

Natural-laminar-flow airfoil for general-aviation applications

Abstract: A natural-Iamina r-flow airfoil, the NLF(1)-0115, has been recently designed for general-aviation aircraft at the NASA Langley Research Center. During the design of this airfoil, special emphasis was placed on experiences and observations gleaned from other successful general-aviation airfoils. For example, the flight lift-coefficient range is the same as that of the turbulent-flow NACA 23015 airfoil. Also, although beneficial for reducing drag and producing high lift, the NLF(1)-0115 airfoil avoids the use of aft loading, which can lead to large stick forces if utilized on portions of the wing having ailerons. Furthermore, not using aft loading eliminates the concern that the high pitching-moment coefficient generated by such airfoils can result in large trim drag if cruise flaps are not employed. The NASA NLF(1)-0115 airfoil has a thickness of 15% chord. It is designed primarily for general-aviation aircraft with wing loadings of 720-960 N/m2 (15-20 lb/ft2). Low-profile drag as a result of laminar flow is obtained over the range from c, = 0.1 and R = 9 x 106 (the cruise condition) to c, = 0.6 and R = 4 x 106 (the climb condition). While this airfoil can be used with flaps, it is designed to achieve a c,,max of 1.5 at R = 2.6 x 10 6 without flaps. The zero-lift pitching moment is held to c,H,0 = -0.055. The hinge moment for a 20% chord aileron is fixed at a value equal to that of the NACA 632-215 airfoil, CH = — 0.0022. The loss in cAmax due to leading-edge roughness at R = 2.6 x 10 6 is 11% as compared with 14% for the NACA 23015.

Wing flutter boundary prediction using unsteady Euler aerodynamic method

Abstract: Modifications to an existing three-dimensional, implicit, upwind Euler/Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the government flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 degree swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.

Method for making a double-wall airfoil

Abstract: A method for making double-wall airfoil for applications such as the blades and vanes of gas turbine engines by depositing an airfoil skin over an inner support wall which is separately formed and contains channels filled by a channel filling means. The channel filling means is removed thereby forming integral channels within the double-wall for circulating a cooling gas adjacent to the airfoil skin. The airfoil skin deposited may be a metal alloy skin or a microlaminate structure, including microlaminate composite structures.

Double-wall airfoil

Abstract: A double-wall airfoil for applications such as the blades and vanes of gas turbine engines. The double-wall comprises an outer airfoil skin and an inner support wall that are metallurgically bonded to one another. The double-wall contains integral channels for passage of cooling air adjacent to the airfoil skin. Airfoil skin may be a metal alloy skin or a microlaminate structure, including microlaminate composite structures. Microlaminate composites typically have a lower density than that of the material used for the airfoil support wall, and a simplified internal geometry which promote weight reductions in the airfoils and increases in engine operating efficiency.

Aerodynamic shape optimization of wing and wing-body configurations using control theory

Abstract: This paper describes the implementation of optimization techniques based on control theory for wing and wing-body design. In previous studies it was shown that control theory could be used to devise an effective optimization procedure for airfoils and wings in which the shape and the surrounding body-fitted mesh are both generated analytically, and the control is the mapping function. Recently, the method has been implemented for both potential flows and flows governed by the Euler equations using an alternative formulation which employs numerically generated grids, so that it can more easily be extended to treat general configurations. Here results are presented both for the optimization of a swept wing using an analytic mapping, and for the optimization of wing and wing-body configurations using a general mesh.

ONERA Three-Dimensional Icing Model

Abstract: A three-dimensional icing model has been developed at ONERA to calculate ice accretion shapes for aerodynamic components that can not be predicted using conventional two-dimensional codes. It is described, emphasizing the original parts with respect to the two-dimensional existing models. The model includes Euler inviscid flow calculation. Droplet trajectories are calculated in a three-dimensional grid. The remesh on the leading edge is adapted to follow aerodynamics singularities. The boundary layer is calculated using a mixing length formulation to model the wall roughness influence on convective heat transfer. Runback paths are integrated. The heat balance is calculated in a grid created along the runback paths. The domain of validity of the three-dimensional icing code is described; compared with the two-dimensional model this domain is wider, especially for high speeds. The three-dimensional model is shown to simulate well a uniform ice deposit on a three-dimensional rotor blade tip. Then, a comparison of the three- and two-dimensional codes on an infinite swept wing shows that the corrected two-dimensional code predicts the catch efficiency but not the ice shape. Finally, it is shown that the continuum flux hypothesis prevents the three-dimensional model from simulating correctly the "lobster tail" ice shape (nonuniform ice deposit).

Investigation of the unsteady flow velocity field above an airfoil pitching under deep dynamic stall conditions

Abstract: The unsteady flow field above a NACA 0012 airfoil pitching under deep dynamic stall conditions has been investigated in a low-speed wind tunnel by means of particle image velocimetry. The measurements of the instantaneous flow velocity field show the characteristic features of the dynamic stall process: formation and development of an organized vortex structure for increasing incidences and the subsequent separation. Vorticity and divergence estimated from the measured data give a good insight into the complex flow behaviour during the downstroke motion. Furthermore, small-scale structures could be observed in the separated flow field and even within the dynamic stall vortex.

Influence of incidence angle on sound generation by airfoils interacting with high-frequency gusts

Abstract: A theoretical model is developed for the sound generated when a convected vortical or entropic gust encounters an airfoil at non-zero angle of attack. The theory is based on a linearization of the Euler equations about the steady subsonic flow past the airfoil. High-frequency gusts, whose wavelengths are short compared to the airfoil chord, but long compared to the displacement of the mean-flow stagnation point from the leading edge, are considered. The analysis utilizes singular-perturbation techniques and involves four asymptotic regions. Local regions, which scale on the gust wavelength, are present at the airfoil leading and trailing edges. Behind the airfoil a ‘transition’ region, which is similar to the transition zone between illuminated and shadow zones in optical problems, is present. In the outer region, far away from the airfoil edges and wake, the solution has a geometric-acoustics form. The primary sound generation is found to be concentrated in the local leading-edge region. The trailing edge plays a secondary role as a scatterer of the sound generated in the leading-edge region. Parametric calculations are presented which illustrate that moderate levels of airfoil steady loading can significantly affect the sound field produced by airfoil-gust interactions.

Boundary Layer Development in Axial Compressors and Turbines: Part 1 of 4 — Composite Picture

Abstract: Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading.Part 1 of this paper draws a composite picture of boundary layer development in turbomachinery based upon a synthesis of all of our experimental findings for the compressor and turbine. Parts 2 and 3 present the experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data.For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien-Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment.Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.Copyright © 1995 by ASME

Dynamic Stall Control for Advanced Rotorcraft Application

Abstract: Advanced concepts designed to improve the lift, drag, and pitching moment characteristics of rotor blades have been investigated for the purpose of enhancing rotor maneuver capability. The advantages and disadvantages of these concepts have been evaluated using both computational and experimental means. The concepts that were considered in this study included a leading-edge slat, a deformable leading-edge, and upper-surface blowing. The results show the potential of these concepts for substantially improving the performance of a rotor. HE next generation of rotorcraft will be required to operate at much higher performance levels than in the past, particularly in the areas of nap-of-the-ea rth (NOE), deep-penetration operations, and air-to-air combat. These new requirements will require highly maneuverable, agile, and survivable rotorcraft, far exceeding the capabilities of those in the current inventory. The objectives of this project include an increase in the maneuverability/agility capability of the helicopter and a reduction in the acoustic detection range. The single most important element of the rotorcraft for meeting these requirements is the rotor itself, since it is the primary source of lift, control, and speed. At the same time, the rotor is also a major source of acoustically detectable radiation. Among the many factors affecting rotorcraft performance, the aerodynamic characteristics of the rotor system are the most important and are the main subject of this paper. The maneuvering capability of a rotorcraft can be improved by re- ducing or suppressing the vibratory loads on the rotor blades caused by aerodynamic separation and stall. This would have the effect of expanding the stall-limiting boundary of the rotor and thereby increase the available load factor in all flight regimes. The con- ventional way to obtain higher lift is to increase the blade area, however, this usually results in a heavier rotor that is also less ef- ficient. With regard to compressibili ty effects and acoustic radia- tion, improvements have been obtained by sweeping, tapering, and thinning the tip region of the rotor blade. As a result, numerous families of airfoils and planform shapes have evolved that offer bet- ter advancing-blade characteristics. However, improvements on the retreating-blade side have not been as impressive. One reason for this imbalance may be that design codes are available for treating blades at low angles of attack and high Mach number (characteris- tic of the advancing side), whereas the design strategy has had to depend heavily on costly empirical studies for blades at high angles of attack and having some amount of separation (characteristic of the retreating side). Increasing the tip speed of the rotor to achieve a maneuvering ad- vantage may produce a dangerous condition with regard to acoustic detection. Rapid advancements in passive acoustic sensor arrays and advanced signal processing technologies pose a serious threat to the mission effectiveness of Army helicopters. Since the rotor blade generates acoustic radiations that can be easily detected and identified, airfoil and planform shapes must be carefully optimized to reduce the detection range of the rotorcraft. The requirements for improved maneuverability and reduced sus- ceptibility will clearly demand a substantial growth in the technolo- gies for addressing rotor aerodynamics. New control techniques must be considered, both passive and active, and these must be ac- companied by a more thorough physical understanding of these flow phenomena along with substantially improved prediction capabili- ties. To meet these requirements, computational and experimental efforts have been initiated to evaluate the effectiveness of various concepts. At present these concepts include airfoils with slats and slots, airfoils that deform, and airfoils with flow energizers. Description of Experiment and Computational Fluid Dynamics Code

Airfoils for wind turbine

Abstract: Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

Real-time prediction of unsteady aerodynamics: Application for aircraft control and manoeuvrability enhancement

Abstract: The capability to control unsteady separated flow fields could dramatically enhance aircraft agility. To enable control, however, real-time prediction of these flow fields over a broad parameter range must be realized. The present work describes real-time predictions of three-dimensional unsteady separated flow fields and aerodynamic coefficients using neural networks. Unsteady surface-pressure readings were obtained from an airfoil pitched at a constant rate through the static stall angle. All data sets were comprised of 15 simultaneously acquired pressure records and one pitch angle record. Five such records and the associated pitch angle histories were used to train the neural network using a time-series algorithm. Post-training, the input to the network was the pitch angle (/spl alpha/), the angular velocity (d/spl alpha//dt), and the initial 15 recorded surface pressures at time (t/sub 0/). Subsequently, the time (t+/spl Delta/t) network predictions, for each of the surface pressures, were fed back as the input to the network throughout the pitch history. The results indicated that the neural network accurately predicted the unsteady separated flow fields as well as the aerodynamic coefficients to within 5% of the experimental data. Consistent results were obtained both for the training set as well as for generalization to both other constant pitch rates and to sinusoidal pitch motions. The results clearly indicated that the neural-network model could predict the unsteady surface-pressure distributions and aerodynamic coefficients based solely on angle of attack information. The capability for real-time prediction of both unsteady separated flow fields and aerodynamic coefficients across a wide range of parameters in turn provides a critical step towards the development of control systems targeted at exploiting unsteady aerodynamics for aircraft manoeuvrability enhancement. >

Navier-Stokes Computations and Experimental Comparisons for Multielement Airfoil Configurations

Abstract: A two-dimensional unstructured Navier-Stokes code is utilized for computing the flow around multielement airfoil configurations. Comparisons are shown for a landing configuration with an advanced-technology flap. Grid convergence studies are conducted to assess inaccuracies caused by inadequate grid resolution. Although adequate resolution is obtained for determining the pressure distributions, further refinement is needed to sufficiently resolve the velocity profiles at high angles of attack. For the advanced flap configuration, comparisons of pressure distributions and lift are made with experimental data. Here, two flap riggings and two Reynolds numbers are considered. In general, the trends caused by variations in these quantities are well predicted by the computations, although the angle of attack for maximum lift is overpredicted.

Numerical investigation of dynamic stall of an oscillating wing

Abstract: The unsteady three-dimensional flowfield over an oscillating wing is investigated with the numerical solution of the compressible, time-dependent, Reynolds-averaged Navier-Stokes equations. Spatial discretization is performed with a third-order accurate, upwind-biased, vertex-based, finite volume scheme. An alternative direction implicit, iterative scheme is used for the time integration. The high Reynolds number turbulent flow behavior is modeled with a one-equation turbulence model. The effect of subiterations, time step and grid density on the accuracy of the computed solutions is investigated. It is found that scaling of the time step with the angular velocity of the motion produces accurate solutions at a reduced computational cost. The computational domain over an aspect ratio 5 wing with rounded tip and NACA-0015 airfoil sections is discretized with a single-block grid. The light stall flowfield over the wing oscillating in a subsonic freestream with a mean angle of attack of 11 deg and an amplitude of 4.2 deg is computed. The structure of the separated, unsteady flowfield is investigated and comparisons with available experimental data are performed.

Airfoil with dual source cooling

Abstract: An airfoil (42) for a gas turbine engine (10) having an airfoil portion (52) with a leading and a trailing edge (60, 62) includes dual pressure source cooling. The trailing edge (62) of the airfoil (42) includes an internal trailing edge passage (78) and the leading edge (60) includes an internal leading edge passage (64). Compressor bleed air at higher pressure is channeled through the leading edge passage (64) whereas compressor bleed air at lower pressure is channeled through the trailing edge passage (78). The higher internal pressure in the leading edge (60) ensures that the inward flow of products of combustion does not occur.

Efficiency and Accuracy of Time-Accurate Turbulent Navier-Stokes Computations

Abstract: The accuracy and efficiency of two types of subiterations in both explicit and implicit Navier-Stokes codes are explored for unsteady laminar circular-cylinder flow and unsteady turbulent flow over an 18-percent-thick circular-arc (biconvex) airfoil. Grid and time-step studies are used to assess the numerical accuracy of the methods. Nonsubiterative time-stepping schemes and schemes with physical time subiterations are subject to time-step limitations in practice that are removed by pseudo time sub-iterations. Computations for the circular-arc airfoil indicate that a one-equation turbulence model predicts the unsteady separated flow better than an algebraic turbulence model; also, the hysteresis with Mach number of the self-excited unsteadiness due to shock and boundary-layer separation is well predicted.

A new high resolution scheme for compressible viscous flows with shocks

Abstract: A new flux splitting and limiting technique which yields one-point stationary shock capturing is presented. The technique is applied to the full NavierStokes and Reynolds Averaged Navier-Stokes equations. Calculations of laminar boundary layers at subsonic and supersonic speeds are presented together with calculations of transonic flows around airfoils. The results exhibit very good agreement with theoretical solutions and existing experimental data. It is found that. the proposed scheme improves the resolution of viscous flows while maintaining excellent one-point shock capturing characteristics. d

Investigation of flow at leading and trailing edges of pitching-up airfoil

Abstract: The dynamic stall process of an NACA 0012 airfoil undergoing a constant-rate pitching-up motion is studied experimentally in a water towing tank facility. This study focuses on the detailed measurement of the unsteady separated flow in the vicinity of the leading and trailing edges of the airfoil. The measurements are carried out using the particle image velocimetry technique. This technique provides the two-dimensional velocity and associated vorticity fields, at various instants in time, in the midspan of the airfoil. Near the leading edge, large vortical structures emerge as a consequence of van Dommelen and Shen type separation and a local vorticity accumulation. The interaction of these vortices with the reversing boundary-layer vorticity initiates a secondary flow separation and the formation of a secondary vortex. The mutual induction of this counter-rotating vortex pair eventually leads to the ejection process of the dynamic stall vortex from the leading-edge region. It is found that the trailing-edge flowfield only plays a secondary role on the dynamic stall process.

Turbine airfoil with enhanced cooling

Abstract: Cooling of a turbine airfoil such as a turbine stator vane in a gas turbine engine is enhanced by the provision of additional cooling air through a secondary air inlet communicating with a medial portion of an internal, serpentine cooling air passage.