Showing papers on "Airfoil published in 1980"

An Experimental Study of Endwall and Airfoil Surface Heat Transfer in a Large Scale Turbine Blade Cascade

Abstract: Local rates of heat transfer on the endwall, suction, and pressure surfaces of a large scale turbine blade cascade were measured for two inlet boundary layer thicknesses and for a Reynolds number typical of gas turbine engine operation. The accuracy and spatial resolution of the measurements were sufficient to reveal local variations of heat transfer associated with distinct flow regimes and with regions of strong three-dimensional flow. Pertinent results of surface flow visualization and pressure measurements are included. The dominant role of the passage vortex, which develops from the singular separation of the inlet boundary layer, in determining heat transfer at the endwall and at certain regions of the airfoil surface is illustrated. Heat transfer on the passage surfaces is discussed and measurements at airfoil midspan are compared with current finite difference prediction methods.

Low-speed single-element airfoil synthesis

Abstract: Large quantities of experimental data exist on the characteristics of airfoils operating in the Reynolds number range between one and ten million, typical of conventional atmospheric wind tunnel operating conditions. Beyond either end of this range, however, good experienental data becomes scarce. Designers of model airplanes, hang gliders, ultralarge energy efficient transport aircraft, and bio-aerodynamicists attempting to evaluate the performance of natural flying devices, are hard pressed to make the kinds of quality performance/design estimates taken for granted by sailplane and general aviation aerodynamicists. Even within the usual range of wind tunnel Reynolds number, much of the data is for "smooth" models which give little indication of how a section will perform on a wing of practical construction. The purpose of this paper is to demonstrate the use of recently developed airfoil analysis/design computational tools to clarify, enrich and extend the existing experimental data base on 1ow-speed, single element airfoils, and then proceed to a discussion of the problem of tailoring an airfoil for a specific application at its appropriate Reynolds number. This latter problem is approached by use of inverse (or "synthesis") techniques, wherein a desirable set of boundary layer characteristics, performance objectives, and constraints are specified, which then leads to derivation of a corresponding viscous flow pressure distribution. In this plocedure, the airfoil shape required to produce the desired flow characteristics is only extracted towards the end of the design cycle. This synthesis process is contrasted with the traditional "analysis" (either experimental or computational) approach in which an initial profile shape is selected which then yields a pressure distribution and boundary layer characteristic, and finally some perfonnance level. The final configuration which provides the required performance is derived by cut-and-try adjustments to the shape. These two approaches are shown disgramatically.

Flow Curvature Effects on Darrieus Turbine Blade Aerodynamics

Abstract: The effects of curvilinear flow on Darrieus turbine blade aerodynamics are described. Analysis shows that these effects can have a sizeable impact on performance for blades of large chord. Experimental data are presented which verify this forecast. Unusually large boundary-layer radial pressure gradients and virtually altered camber and incidence are identified as causal phenomena. Conformaf mapping techniques are used to transform geometric airfoils in curved flow to their virtual equivalents in rectilinear flow. It is argued that flow curvature is an important determinant of Darrieus turbine blade aerodynamic efficiency and that its proper consideration will yield performance improvements, even for blades of small chord. uncovered blade aerodynamic complexities which were beyond initial expectations. Most noteworthy are the peculiar aerodynamic phenomena associated with the orbital motion of the blades. In essence, these blades are subjected to a curvilinear flow and behave very differently than if they were immersed in a rectilinear flow. Furthermore, centrifugal forces alter the boundary layer of the airfoils rotating in this fashion. This situation presents problems in the design and analysis of all cross-flow wind turbines, because virtually all published airfoil data are derived from tests in rectilinear flow. Recent studiesl show that modest improvements in Cp yield desirable reductions in the cost of energy. Since these Cp improvements can be achieved by increasing blade aerodynamic efficiency, there exists ample incentive for considering the aerodynamic idiosyncrasies of rotating blades. In the material which follows, boundary-layer centrifugal effects will be discussed first. Treatment of this subject is brief; its significance has only recently been appreciated and extensive studies of the phenomenon have not been con- ducted. Flow curvature effects are treated next. A simple kinematic analysis demonstrates that the turbine blade relative inflow velocity and angle of attack are unique everywhere on the chord. It is then shown how conformal mapping techniques, which transform airfoils in the curved flow field to their virtual equivalents in rectilinear flow, may be used in the aerodynamic analysis of the turbine blades. The method indicates that flow curvature effects are strongly dependent upon the blade chord to turbine radius C/R. Experimental data are introduced for two sets of blades, both of NACA 0015 airfoil section. The first set of blades had C/R = 0.114 and the second set had C/R = 0.260.

Laminar Separation, Transition, and Turbulent Reattachment near the Leading Edge of Airfoils

Abstract: The laminar separation, transition, and turbulent reattachment near the leading edge of a cylindrical noseconstant thickness airfoil model were investigated using a low-turbulence, low-speed smoke wind tunnel. The locations of separation, transition, and reattachment were obtained from smoke flow photographs and surface oil flow techniques for chord Reynolds numbers from about 150,000 to 470,000. These visual data combined with static pressure distributions delineate the effects of angle of attack, flap deflection angle, and chord Reynolds number on the separation bubble characteristics. The data concerning the length of the laminar and turbulent portions of the bubble agree with the empirical prediction methods for short bubbles.

Joined wing aircraft

Abstract: An aircraft having a fuselage and a pair of first airfoils in the form of wings extending outwardly from the vertical tail and a pair of second airfoils in the form of wings extending outwardly from the forward portion of the fuselage at a lower elevation than the first airfoils. The second wings extend rearwardly having a positive dihedral so that the tip ends of the second airfoil are located in close proximity to and may overlap the tip ends of the first wings. The pairs of wings along with the fuselage present a double triangle or diamond shape in both front elevational view and top plan view. A winglet structurally connects the tip ends of the corresponding first wings and second wings, and these winglets have airfoil surfaces which extend vertically substantially beyond the tip ends of the first and second wings in order to minimize the effects of induced drag and also to augment directional stability of the aircraft. In addition, a unique wing structure is disclosed where the average thickness varies along the chord of the wing to enhance resistance to the component of lift acting normal to the spanwise plane containing the centroids of the airfoils.

Influence of Propeller Design Parameters on Far-Field Harmonic Noise in Forward Flight

Abstract: A theory for harmonic noise radiation is studied for general guidance to the designer and is applied to some propeller noise problems of current interest. Only the linear sources are studied in detail. The frequency domain results clarify the role of acoustic noncompactness (noise cancellation due to finite chord and span effects). Nondimensional parameters arising from the analysis give design guidance by showing the potential for noise reduction due to changes in airfoil section and blade sweep, twist, and taper as functions of operating conditions. Conventional propellers are shown to be relatively insensitive to variations in blade design. However, advanced turbopropellers (prop fans) currently under development are decidedly noncompact because of their high solidity and speed. Examples of chord wise and span wise cancellation are given illustrating substantial benefits of sweep.

Unsteady laminar separation - An experimental study

Abstract: : The design of most aerodynamic surfaces, as for example the helicopter rotor, is based essentially on quasi-steady theories. However the dynamics of a rotating blade introduce unexpected fluctuations and overshoots of properties like lift, drag, etc. The phenomenon of unsteady stall is intimately connected with the development of an oscillating boundary layer and separation. Experimental investigation of such flows was undertaken by a method of visualization developed especially for the study of laminar or turbulent boundary layers and separation. The method captures the instantaneous 2-D flow field, including regions of separated flow and provides accurate quantitative information. Laser doppler anemometer measurements complement the optically received data. Results reveal that separation responds with time-lag to external disturbances, in agreement with unsteady stall data. Oscillating outer flows result in displacement of the point of separation and under certain conditions, the Despard and Miller criterion was found to hold. Earlier theoretical models of separation are confirmed qualitatively and for the early stages of the transient phenomena. The findings provide physical insight and quantitative data that may help understand the phenomenon of unsteady stall and unsteady separation. (Author)

Transonic Flutter Analysis of a Rectangular Wing with Conventional Airfoil Sections

Abstract: Flutter analysts have encountered considerable analytical difficulties in the prediction of the flutter stability of aircraft operating in the transonic Mach number regime Because of the shocks and nonlinearities of transonic flow the aerodynamic unsteady forces have been difficult to determine and have prohibited accurate determination of the flutter speed The finite-difference relaxation method is used to determine the oscillatory transonic aerodynamic forces on a uniformly stiff cantilever rectangular wing in a flow field with mixed subsonic and supersonic regions together with shock waves The flutter speed is determined at two transonic Mach numbers and is compared to the flutter speed obtained using a linear aerodynamic theory

Airstream modification device for airfoils

Abstract: An airfoil modification device that helps maintain airstream attachment to either upper or lower airfoil surfaces, the device in the form of a plurality of spaced, low profile rods arranged generally end-to-end and attached to the airfoil and generally normal to the airstream to define a serrated edge that induces multi-directional airstream turbulence.

Application of the fast Fourier transform to two-dimensional wind tunnel wall interference

Abstract: Wall interference corrections are evaluated from experimental wind tunnel wall pressure distributions using the Fourier solution for the Dirichlet problem in a rectangle. The series coefficients are computed by the fast Fourier transform, making the method very efficient and suitable as a practical wall correction procedure for two-dimensional wind tunnel data. The method is applicable to arbitrary subcritical wind tunnel walls and the knowledge of their cross-flow properties is not required. A practical example is given for the BGK 1 airfoil, tested at supercritical flow conditions in the 20% perforated wall test section of the NAE high Reynolds number wind tunnel.

Numerical Study of Transonic Flutter of A Two-dimensional Airfoil

Abstract: Transonic flutter characteristics of a NAGA64A010 airfoil with two-degrees-of-freedom are investigated theoretically. For this purpose, an unsteady aerodynamic code based on the transonic small perturbation equation, which can be applied for the wide range of the reduced frequency based on semi-chord(0 less than or equal to k less than or equal to 0.5) and Mach number(from subcritical to above Mach 1), has been developed. The finite differnce scheme employed in the code is a time-marching, semi-implicit and implicit two-sweep procedure. Flutter calculations are performed for two typical binary systems, one of which simulates the vibrational characteristics of a typical streamwise section of a sweptback wing, and the other of which simulates that of an unswept wing. A sharp transonic dip of the flutter boundary has been predicted for the former case while the relatively mild dip for the latter. For the purpose of identifying the possible mechanism of the transonic dip phenomenon, examinations are made of not only the flutter modes and frequencies but also the shock wave patterns and the unsteady load distributions at each Mach numbers corresponding to the flutter boundary. As a result of these examinations, it is concluded that the mechanism of the single-degree-of-freedom flutter, which is casued by the large negative damping produced by the phase lag of the shock wave motion, is dominating the flutters at the bottom of the transonic dip when the mass ratio is relatively large.

Composite blade for turbofan engine fan

Abstract: A turbofan engine fan blade made of composite material, and a method of making the fan, are taught The fan blade essentially comprises an airfoil section having a root end and made of a plurality of bonded plies of composite material which are splayed and which are in a staggered condition at the root end; a two-piece platform section made of titanium or of aluminum, with one piece of the platform on each side of the airfoil section; and a steel outsert section which holds and secures the platform section to the airfoil section, with the outsert section having a triangular shaped cavity located at the root end of the airfoil section, between the splayed and staggered plies of the airfoil section Among other advantages, the cavity eliminates the "insert plies" (or wedge) used in the prior art and the inherent disadvantages associated with such use

Airfoil shape for arrays of airfoils

Abstract: An Airfoil Shape for Arrays of Transonic Airfoils Abstract A flow directing assembly 14 having an airfoil section or shape 28 of the type adapted for use in an axial flow gas turbine engine is disclosed. The cambered meanline MCL of the airfoil shape is formed of a front circular arc FA and a rear circular arc RA. Athickness distribution TD is applied to the meanline to form the convex suction surface 20 and the concave pressure surface 22. The airfoil section exhibits good aerodynamic performance as compared with an equivalent circular arc airfoil in a transonic flow field.method for making the airfoil shape is disclosed. The method includes the steps of: forming a cambered mean-line of two circular arcs; forming a thickness distri-bution about the conical chord line Bt; and applying thickness distribution to the cambered meanline such that a portion of the suction surface is stretched and a portion of the pressure surface is compressed. .

Wind turbine blade

Abstract: A wind turbine blade of large size for a wind turbine (14) having three blades (16) and used to generate electrical power is contoured to minimize stalling and to provide unusually good efficiency. The cross section of the blade tapers from a configuration at the hub or inboard end with substantial leading and trailing edge deflection toward the wind providing high lift at low speed, to the outboard one-fourth to one-fifth which is configured as a conventional low lift airfoil since it moves through the air at comparatively high speed. The chord length of approximately the inboard one-third of the blade is chosen such that, in combination with the deflection, very little air is permitted to cross this inboard part of the blade but is forced to flow radially outwardly such that by reason of an increase in dynamic pressure it augments the flow across the faster moving part of the blade. The blade may be formed of a single long tapered spar (34) having great strength which is fastened to the hub with a plurality of ribs (a-i) attached perpendicularly to the spar to define the desired cross-sectional configurations, this assembly being covered with a suitable skin such as a glass epoxy cloth.

Flutter analysis of a NACA 64A006 airfoil in small disturbance transonic flow

Abstract: Flutter analyses are performed for a NACA 64A006 airfoil pitching and plunging in small-disturbance, unsteady transonic flow. Aerodynamic coefficients are obtained for M^ = 0.7, 0.8, 0.85, 0.8625, and 0.87, and for various values of low reduced frequencies. Two computer codes are used: 1) STRANS2 and UTRANS2 based upon the relaxation method and 2) LTRAN2 based upon the time-integration (indicial) method. Flutter results are presented as plots of flutter speed and corresponding reduced frequency vs one of the four parameters: airfoil/air mass density ratio, position of mass center, position of elastic axis, and freestream Mach number. In each figure, several sets of curves for different values of plunge/pitch frequency ratios are shown. The two sets of results based upon the two separate computer codes are, in general, in good agreement. For a special flutter analysis of a flat plate at M^ =0.7, the present methods agree well with the linear flat plate theory. The effect of each parameter on the trend of each curve of flutter speed is discussed in detail. The examples demonstrate the dip phenomenon of the curves for flutter speed in the transonic regime. Flutter results of transonic codes are compared with those obtained by linear flat plate theory.

Attachable airfoil with movable control surface

Abstract: An attachable airfoil for use on an airborne vehicle or payload having a plurality of joined, nestable sections which can expand from a collapsed, streamlined position adjacent the vehicle or payload to a fully extended airfoil configuration In addition, the airfoil incorporates therein a movable control surface which by appropriate remote actuation can be utilized to provide controlled flight of the airborne vehicle or payload

Some Features of the Unsteady Pressure Field in Transonic Airfoil Buffeting

Abstract: Comparisons are made between the unsteady transonic flowfields of two airfoils: a Whitcomb supercritical airfoil, and a conventional NACA 0012 section. Wind tunnel experiments on these airfoils included penetration into buffeting as a result of high cl and/or high M^. Fluctuating surface pressure, lift, and shock location were measured on both airfoils. Two-point pressure cross-correlations were used to determine coherence and propagation direction of pressure fluctuation patterns on the upper surface of each airfoil. Between the uppersurface shock and the trailing edge (a region with intense pressure fluctuations), pressure disturbances propagated upstream in attached flow, but traveled downstream when extensive separation existed. In the latter case, convection velocities were found to be frequency dependent. Another cross-correlation, relating surfacepressure fluctuations to unsteady lift, was employed to establish which regions of the pressure fields were of primary importance in producing buffeting forces. While many of the principal features of the pressure/lift cross-correlations were common to both airfoils, some specific differences were found. For example, the supercritical airfoil exhibited less periodicity in its cross-correlation. This result was attributed to the flattopped, aft-cambered shape of the supercritical airfoil section, which reduced the coupling between shock oscillations and lift fluctuations.

Experimental studies of the laminar separation bubble on a two-dimensional airfoil at low Reynolds numbers

Abstract: The laminar separation, transition and turbulent reattachment near the leading edge of a two-dimensional NACA 663-018 airfoil were investigated using a low speed, smoke visualization wind tunnel. Lift and drag force measurements were made using an external strain gauge balance for a chord Reynolds number range of 40,000 to 400,000. An extensive flow visualization study was performed and correlated with the force measurements. Experiments were also conducted with distributed surface roughness at the leading edge and external acoustic excitation to influence the development of the airfoil boundary layer. This study delineates the effects of angle of attack and chord Reynolds number on the separation bubble characteristics and airfoil performance

A conservative implicit finite difference algorithm for the unsteady transonic full potential equation

Abstract: An implicit finite difference procedure is developed to solve the unsteady full potential equation in conservation law form Computational efficiency is maintained by use of approximate factorization techniques The numerical algorithm is first order in time and second order in space A circulation model and difference equations are developed for lifting airfoils in unsteady flow; however, thin airfoil body boundary conditions have been used with stretching functions to simplify the development of the numerical algorithm

Hydrodynamic interactions of ships with fixed obstacles

Abstract: The hydrodynamic interactions of a slow-moving vessel with a shoreline or an obstacle in shallow water are investigated using slender-body theory. With the assumption that the free surface is rigid, the problem is shown to be equivalent to studying the flow about a porous airfoil moving near an irregular boundary in the horizontal plane. A numerical procedure is developed based on the availability of an "obstacle Green function." Theoretical results indicate that the vessel experiences an attractive force and a bow-in moment during the approach toward a wedge-shaped obstacle, such as a pier or a breakwater, regardless of the angle of approach. For a vessel approaching a side bank at an angle, the theory yields results consistent with the so-called "bank-rejection" phenomenon.

Shock-Free Wing Design

Abstract: A simple numerical method for generating wing shapes that will be shock free at a specified supercritical Mach number is described. The method involves using a fictitious gas law for the supersonic domain to make the governing equations elliptic. Requirements on this gas law are detailed and a method for computing the real flow in the supersonic domain, given initial data on the embedded sonic surface, is described. The failure of the method to yield a shock-free flow when a limit surface occurs in the supersonic flow, and the difficulties that arise because the initial-value problem for the supersonic domain is ill-posed, are delineated. Finally, a small perturbation algorithm is used to illustrate the procedure and results are given for a simple baseline wing. NCREASED fuel efficiency, and in the case of commercial aircraft, productivity, can be achieved by operating aircraft at supercritical Mach numbers, provided that shock waves can be avoided or made acceptably weak. Two-dimensional procedures for prescribing airfoil sections that are shock free have already provided improvements in aircraft efficiency by employing these airfoils on swept wings. Three-dimensional effects have compromised such designs to some extent, and extensive wind tunnel development tests have been required to recapture the benefits of these "supercritical airfoils." Sobieczky et al. l demonstrated a method of modifying baseline configurations so that they would be shock free at a prescribed Mach number and lift coefficient. This procedure provides a special opportunity for improving aircraft per- formance through a careful selection of the baseline con- figuration in order to provide wings and wing-body com- binations that are shock free at supercritical Mach numbers, and that have acceptable off-design performance. Yu2 and Yu and Rubbert3 have also documented that this procedure is possible and demonstrated its application. As was first described by Sobieczky,4 a numerical algorithm is used to solve a fictitious set of equations for the flow past the baseline configuration. These equations are identical to the correct equations for subsonic portions of the flow, but they are modified when the flow becomes super- sonic, so that even though the flow speed is larger than the local speed of sound the equations themselves remain elliptic. This procedure generates a numerical solution that satisfies the appropriate equations where the flow is subsonic, and the appropriate boundary conditions on the configuration outside of the supersonic zone. The results of this calculation provide the flowfield at the sonic surface. This surface and flowfield define an ill-posed initial value problem for the supersonic domain that is to be solved using the correct equations. Because this problem is ill-posed in three dimensions any numerical method must, in principle, be unstable. This in- stability, however, is of no consequence for moderate to high aspect ratios. However, if the detailed definition of the spanwise modifications required to make the wing shock free

Unsteady aerodynamics of conventional and supercritical airfoils

Abstract: The unsteady aerodynamics of a conventional and a supercritical airfoil are compared by examining measured chordwise unsteady pressure time-histories from four selected flow conditions. Although an oscillating supercritical airfoil excites more harmonics, the strength of the airfoil's shock wave is the more important parameter governing the complexity of the unsteady flow. Whether they are conventional or supercritical, airfoils that support weak shock waves induce unsteady loads that are qualitatively predictable with classical theories; flows with strong shock waves are sensitive to details of the shock-wave and boundary-layer interaction and cannot be adequately predicted.

Effect of Interblade Phase Angle and Incidence Angle on Cascade Pitching Stability

Abstract: A comprehensive test program was performed at low subsonic velocity on a linear cascade of airfoils oscillating in pitch about their midchords for incidence angles up to 10 deg, reduced frequencies up to 0.193, and over a range of interblade phase angles from σ = −60 deg to +60 deg. The test conditions represent significant changes in blade loading and dimensionless frequency, and the range of interblade phase angle includes those values usually encountered in actual turbomachines. The measured pressure time histories over the airfoil chord were used to calculate the stability parameters of the system including the unsteady pitching moment coefficient and the aerodynamic damping parameter. For the range of parameters tested it was found that the interblade phase angle is the most important parameter affecting the stability of oscillating cascaded airfoils. The system was unstable for most positive values of σ over the entire range of loading and frequency. This was similar in behavior (but not in magnitude) to the predictions of available potential flow cascade theories and suppports the observation that blade stall need not be present for torsional “stalled” flutter to occur. System stability for negative values of σ was more dependent on loading and frequency and, conformed more closely with the observed behavior of stalled flutter. Specifically, for σ < 0 deg stability increased with frequency and decreased with loading. A preliminary evaluation of the pressure time histories shows that a second harmonic behavior renders the 1.2 percent chord station ineffective in contributing to the blade damping. Under these circumstances it is surmised that the induced damping is associated mainly with the first harmonic component of the pressure response at the 6.2 percent chord station.

Transonic Shock-Wave/Boundary-Layer Interactions on an Oscillating Airfoil

Abstract: Unsteady aerodynamic loads were measured on an oscillating NACA 64A010 airfoil In the NASA Ames 11 by 11 ft Transonic Wind Tunnel. Data are presented to show the effect of the unsteady shock-wave/boundary-layer interaction on the fundamental frequency lift, moment, and pressure distributions. The data show that weak shock waves induce an unsteady pressure distribution that can be predicted quite well, while stronger shock waves cause complex frequency-dependent distributions due to flow separation. An experimental test of the principles of linearity and superposition showed that they hold for weak shock waves while flows with stronger shock waves cannot be superimposed.

Aerodynamic and Aeroelastic Characteristics of Oscillating Loaded Cascades at Low Mach Number—Part I: Pressure Distribution, Forces, and Moments

Abstract: A complete theory is developed in Part I of these two papers for the analysis of oscillating airfoils in cascade in uniform incompressible flows. The theory fully accounts for the geometry of the airfoils and cascade parameters. It is shown that the strong mean velocity gradient near the leading edges of the airfoils significantly affects the unsteady pressure, forces and moments acting upon the airfoils. In Part II the aeroelastic characteristics of a loaded cascade are investigated for one and two degrees of freedom oscillations.