Showing papers on "Airfoil published in 1989"

XFOIL: An Analysis and Design System for Low Reynolds Number Airfoils

Abstract: Calculation procedures for viscous/inviscid analysis and mixed-inverse design of subcritical airfoils are presented. An inviscid linear-vorticity panel method with a Karman-Tsien compressiblity correction is developed for direct and mixed-inverse modes. Source distributions superimposed on the airfoil and wake permit modeling of viscous layer influence on the potential flow. A two-equation lagged dissipation integral method is used to represent the viscous layers. Both laminar and turbulent layers are treated, with an e 9-type amplification formulation determinining the transition point. The boundary layer and transition equations are solved simultaneously with the inviscid flowfield by a global Newton method. The procedure is especially suitable for rapid analysis of low Reynolds number airfoil flows with transitional separation bubbles. Surface pressure distributions and entire polars are calculated and compared with experimental data. Design procedure examples are also presented.

Airfoil self-noise and prediction

Abstract: A prediction method is developed for the self-generated noise of an airfoil blade encountering smooth flow. The prediction methods for the individual self-noise mechanisms are semiempirical and are based on previous theoretical studies and data obtained from tests of two- and three-dimensional airfoil blade sections. The self-noise mechanisms are due to specific boundary-layer phenomena, that is, the boundary-layer turbulence passing the trailing edge, separated-boundary-layer and stalled flow over an airfoil, vortex shedding due to laminar boundary layer instabilities, vortex shedding from blunt trailing edges, and the turbulent vortex flow existing near the tip of lifting blades. The predictions are compared successfully with published data from three self-noise studies of different airfoil shapes. An application of the prediction method is reported for a large scale-model helicopter rotor, and the predictions compared well with experimental broadband noise measurements. A computer code of the method is given.

Investigation of the flow structure around a rapidly pitching airfoil

Abstract: A numerical study is presented for unsteady laminar flow past a NACA 0015 airfoil that is pitched, at a nominally constant rate, from zero incidence to a very high angle of attack. The flowfield simulation is obtained by solving the full two-dimensional compressible Navier-Stokes equations on a moving grid employing an implicit approximate-factorization algorithm. An evaluation of the accuracy of the computed solutions is presented, and the numerical results are shown to be of sufficient quality to merit physical interpretation. The highly unsteady flowfield structure is described and is found to be in qualitative agreement with available experimental observations. A discussion is provided for the effects of pitch rate and pitch axis location on the induced vortical structures and on the airfoil aerodynamic forces.

A natural low-frequency oscillation of the flow over an airfoil near stalling conditions

Abstract: An experimental and computational study of the low-frequency oscillation observed in the flow over an airfoil at the onset of static stall is presented. Wind-tunnel results obtained with two-dimensional airfoil models show that this phenomena takes place only with a transitional state of the separating boundary layer. It is noted that the flowfield does not involve a Karman vortex street. The experimental results agree well with the results of a two-dimensional Navier-Stokes code. The present study demonstrates that the low-frequency oscillations produce intense flow fluctuations which impart much larger unsteady forces to the airfoil than experienced by bluff-body shedding and which may represent the primary aerodynamics of stall flutter of blades and wings.

Analytical study of the subsonic wing-rock phenomenon for slender delta wings

Abstract: An analytic expression describing the aerodynamic roll moment has been obtained from the numerical simulation of wing rock. This expression is used in the equation governing the rolling motion of a delta wing around its midspan axis. The result is used to construct phase planes, which reveal the general global nature of wing rock—stable limit cycles, unstable foci, saddle points, and domains of initial conditions leading to oscillatory motion and divergence. An asymptotic approximation to the solution of the governing equation is obtained; this result provides expressions for the amplitudes and frequencies of limit cycles. The present analysis provides a penetrating global view of the wing-rock phenomenon.

Further Generalization of an Equivalent Plate Representation for Aircraft Structural Analysis

Abstract: Recent developments from a continuing effort to provide an equivalent plate representation for aircraft structural analysis are described. Previous work provided an equivalent plate analysis formulation that is capable of modeling aircraft wing structures with a general planform such as cranked wing boxes. However, the modeling is restricted to representing wing boxes having symmetric cross sections. Further developments, which are described, allow modeling of wing cross sections having asymmetries that can arise from airfoil camber or from thicknesses being different in the upper and lower cover skins. An implementation of thermal loadings, which are described as temperature distributions over the planform of the cover skins, has been included. Spring supports have been added to provide for a more general set of boundary conditions. Numerical results are presented to assess the effect of wing camber on the static and dynamic response of an example wing structure under pressure and thermal loading. These results are compared with results from a finite element analysis program to indicate how well a cambered wing box can be represented with an equivalent plate formulation.

Newton solution of inviscid and viscous problems

Abstract: The application of Newton iteration to inviscid and viscous airfoil calculations is examined. Spatial discretization is performed using upwind differences with split fluxes. The system of linear equations which arises as a result of linearization in time is solved directly using either a banded matrix solver or a sparse matrix solver. In the latter case, the solver is used in conjunction with the nested dissection strategy, whose implementation for airfoil calculations is discussed. The boundary conditions are also implemented in a fully implicit manner, thus yielding quadratic convergence. Complexities such as the ordering of cell nodes and the use of a far field vortex to correct freestream for a lifting airfoil are addressed. Various methods to accelerate convergence and improve computational efficiency while using Newton iteration are discussed. Results are presented for inviscid, transonic nonlifting and lifting airfoils and also for laminar viscous cases.

On unsteady surface forces, and sound produced by the normal chopping of a rectilinear vortex

Abstract: An investigation is made of the sound produced when a rectilinear vortex is cut at right angles to its axis by a non-lifting airfoil of symmetric section. The motions are at sufficiently low Mach number that the wavelength of the sound is large relative to the chord of the airfoil. In these circumstances the airfoil experiences no fluctuating lift during the interaction, and the radiation may be ascribed to an acoustic source of dipole type whose strength is equal to the unsteady drag. It is argued that previous analyses of the related problem of ‘unsteady thickness noise’ have ignored certain terms whose inclusion greatly reduces the predicted intensity of the radiation. A general formula for the surface forces (derived in an appendix) is applied to deduce that the dipole strength is proportional to the square of the circulation of the vortex, and depends on the spanwise acceleration of the vortex induced by images in the airfoil. Numerical results are presented for typical airfoil sections, and a comparison is made with the unsteady lifting noise generated when the axis of the vortex is inclined at a small angle to the normal to the median plane of the airfoil.

Separation control on an airfoil by periodic forcing

Abstract: The introduction of transverse velocity fluctuations into a separated shear layer on an airfoil at high angles of attack is presently demonstrated to be an effective separation-control technique. Airfoil aerodynamic characteristics, including poststall lift and drag as well as maximum lift coefficient and stall angle, all exhibited improvements controlled forcing at 20 deg angle of attack led to an increased spreading of the mean velocity profile, together with increased turbulence activity; separation moved from the leading edge to about 80 percent of chord.

Unsteady Euler airfoil solutions using unstructured dynamic meshes

Abstract: Two algorithms for the solution of the time-dependent Euler equations are presented for unsteady aerodynamic analysis of oscillating airfoils. Both algorithms were developed for use on an unstructured grid made up of triangles. The first flow solver involves a Runge-Kutta time-stepping scheme with a finite-volume spatial discretization that reduces to central differencing on a rectangular mesh. The second flow solver involves a modified Euler time-integration scheme with an upwind-biased spatial discretization based on the flux-vector splitting of Van Leer. The paper presents descriptions of the Euler solvers and dynamic mesh algorithm along with results which assess the capability.

Design and experimental results for a high-altitude, long-endurance airfoil

Abstract: Currently, there is interest in the development of high-altitude, long-endurance vehicles for a number of missions including communications relaying, weather monitoring, and provision of targeting information for cruise missiles. The preliminary design of such aircraft is complicated, however, by the lack of suitable airfoils. This is due to the fact that such vehicles, unlike those for which the majority of airfoils have been developed in the past, operate at fairly high lift coefficients and at relatively low Reynolds numbers. Thus, to provide realistic airfoil performance information for preliminary design efforts, an airfoil has been designed for an aircraft with missions similar to those noted. The airfoil is unflapped and has a thickness of 15% chord. The design Reynolds number range is 7 x 10 to 2 x 10. Low drag is predicted for lift coefficients ranging from 0.4, which corresponds to a high-speed dash, to 1.5, the maximum endurance condition. Further, the airfoil is designed specifically such that the maximum lift coefficient is unaffected by surface contamination. Consequently, takeoff and landing in rain or with insect residue on the wings should present no special difficulties. The airfoil has been tested in the NASA Langley Low-Turbulence Pressure Tunnel and, with the exception of the maximum lift coefficient prediction, the results generally confirm the theoretical predictions.

Unsteady transonic small-disturbance theory including entropy and vorticity effects

Abstract: Modifications to unsteady transonic small-disturbance theory to include entropy and vorticity effects are presented. The modifications have been implemented in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code developed recently at the NASA Langley Research Center. The code permits the aeroelastic analysis of complete aircraft configurations in the flutter critical transonic speed range. Entropy and vorticity effects have been incorporated within the solution procedure to more accurately analyze flows with strong shock waves. The modified code includes these effects while retaining the relative simplicity and cost efficiency of the TSD formulation. The paper presents detailed descriptions of the entropy and vorticity modifications along with calculated results and comparisons which assess the modified theory. These results are in good agreement with parallel Euler calculations and with experimental data. Therefore, the present method now provides the aeroelastician with an affordable capability to analyze relatively difficult transonic flows without having to solve the computationally more expensive Euler equations.

Effect of 90 Degree Flap on the Aerodynamics of a Two-Element Airfoil

Abstract: The aerodynamic performance of a two-element airfoil with a 90-deg trailing edge flap was experimentally investigated. The 5 percent-chord long flap, significantly increased the lift of the baseline airfoil, throughout a wide range of angles of attack. The maximum lift coefficient of the flapped wing increased too, whereas the lift/drag ratio decreased.

Vortex Dynamics of Delta Wings

Abstract: The typical angle of attack for maximum lift of a delta wing is about 35°, which is much higher than for a two-dimensional airfoil. The delta wing is, therefore, suitable for highly maneuverable aircraft. In this paper, experimental results for delta wings is reviewed. The review is made from the perspective of fundamental fluid dynamic mechanisms. In particular, the balance between vorticity generation on the surface and freestream convection of it is used to understand how different parameters affect the leading edge vortices which dominate the aerodynamics of a delta wing at high angles of attack.

Airfoil measurement method

Abstract: A coordinate measuring machine is used to extract data points from an airfoil surface. The machine is controlled to move a probe to desired points by a generic program applicable to any airfoil and a nominal data file specific to the type of airfoil being measured containing coordinates of desired data points, surface normal vectors and optional machine control instructions arranged in the order of usage. The program reads the data file and predicts the location of corresponding data points on the surface from the file data and deviation information available from already-sampled data points on the blade. An analysis program accepts the data points, mathematically reconstructs portions of the airfoil and measures dimensions from those portions. The dimensions are compared to tolerances read from another file specific to the airfoil type.

New airfoil design concept

Abstract: The present airfoil design concept is based on utilizing unconventional geometry characteristics near the airfoil trailing edge which include a finite trailing edge thickness, strongly divergent trailing edge upper and lower surfaces, and high surface curvature on the lower surface at or near the lower surface trailing edge. This paper presents computational analyses of airfoils and a wing utilizing the concept, airfoil validation wind tunnel test results of several configurations, and wing-validation wind tunnel test results for a complete wing design. In addition to validating the concept, the airfoil and wing testing provided additional detailed data to better understand the aerodynamic advantage of such an unconventional trailing edge configuration. It is demonstrated that the concept represents a significant step in airfoil technology beyond that achieved with the Supercritical Airfoil. This concept provides the aerodynamicist an additional degree of design freedom and flexibility previously unrecognized.

A Theory for Wake-Induced Transition

Abstract: A theory for transition from laminar to turbulent flow as the result of unsteady, periodic passing of turbulent wakes in the free stream is developed using Emmons' transition model. Comparisons made to flat plate boundary layer measurements and airfoil heat transfer measurements confirm the theory

Aerodynamic air foil surfaces for in-flight control for projectiles

Abstract: Aerodynamical air foil surface and subsurface expressions and/or impressions of varied geometrics, angles of attack, heights and depths, comprising part of a projectile surface itself to create McClain effect molecular friction/pressure/temperature reaction flight control surfaces which automatically achieve in all fluids and velocities of flight self-stablizing spin and rotation, increased height of trajectory with corresponding enhancement of range and distance, kinetic energies, inducing smooth laminar boundary layer flows, substantially decreasing drag effects, synergistically combined to constitute a major technological improvement in performance of all projectiles.

Venturi enhanced airfoil

Abstract: A basic airfoil has its operating performance improved by incorporating one or more apertures in the airfoil adjacent its trailing edge. These apertures extend from the upper surface of the airfoil down through to the lower surface of the airfoil. The entry port and the exit port of these apertures has a greater circumference than that of the throat circumference which is intermediate thereto. This structure forms a venturi having a vertical axis. Spaced below the throat of the aperture are a plurality of air nozzles that communicate with an air plenum chamber within the airfoil. A source of pressurized air is connected to the plenum chamber. The leading edge of the airfoil causes air to flow across both the upper surface and lower surface of the airfoil. The venturi creates a strong suction on the upper surface of the airfoil to enhance the airfoil's pressure differential. When the axis of the venturi is inclined forwardly with respect to the horizontal axis of the airfoil, a vector thrust in the forward direction is created in the same direction as that of the low pressure side of the airfoil. The venturi enhanced airfoil can be utilized both in a horizontal fixed airfoil or its structure can also be incorporated into the tail rudder of an aircraft or helicopter. The venturi enhanced airfoil can be oriented in any position between the horizontal and vertical axes.

Computation of viscous transonic flow over porous airfoils

Abstract: The viscous effects on transonic flow past an airfoil which contains a shallow cavity beneath a porous surface are studied numerically. The porous region occupies a small portion of the total airfoil surface, and is located near the shock. Both an interactive boundary layer (IBL) algorithm and a thin-layer Navier-Stokes (TLNS) algorithm have been modified for use in studying the outer flow, whereas a stream-function formulation has been used to model the inner flow in the small cavity. The coupling procedure at the porous surface is based on Darcy's law and on the assumption of a constant total presusre in the cavity. In addition, a modified Baldwin-Lomax turbulence model is used to consider the transpired turbulent boundary layer in the TLNS approach, and the Cebeci-Smith turbulence model is used in the IBL approach. According to the present analysis, a porous surface can reduce the wave drag appreciably, but it can also increase viscous losses. As has been observed experimentally, the numerical results indicate that the total drag is reduced at higher Mach numbers and increased at lower Mach numbers when the angles of attack are small. Furthermore, the streamline patterns of passive-shock and boundary-layer interaction are revealed in this study.

Aerodynamic performance and flow structure studies of a low Reynoldsnumber airfoil

Abstract: Etude experimentale en soufflerie. Mesure des distributions de pression, des profils de couche limite et des proprietes aerodynamiques correspondantes pour divers angles d'attaque par anemometrie a fil chaud et visualisation au fil de fumee

The Low Frequency Oscillation in the Flow Over a NACA0012 Airfoil with an “Iced” Leading Edge

Abstract: The unusually low frequency oscillation in the wake of an airfoil, studied in [1], is explored experimentally as well as computationally for a NACA0012 airfoil with a “glaze ice accretion” at the leading edge. Experimentally, flow oscillations are observed at low frequencies that correspond to a Strouhal number of about 0.02. This occurs in the angle of attack range of 8° to 9°, near the onset of static stall for this airfoil. With a Navier-Stokes computation, “limit-cycle” oscillations in the flow and in the aerodynamic forces are also observed at low Strouhal numbers. However, the occurrence of the oscillation is found to depend on the turbulence model in use as well as the Reynolds number.

A Comparison Between Boundary Layer Measurements in a Laminar Separation Bubble Flow and Linear Stability Theory Calculations

Abstract: This research examines the details of the boundary layer flowfield from wind tunnel measurements of a two-dimensional Liebeck LA2573A airfoil over a range of Reynolds numbers from 235000 to 500000. In this range, a laminar separation bubble becomes significant in the boundary layer and provides a measurable contribution to the airfoil drag. Measurements include airfoil drag, mean and turbulent boundary layer velocity profiles, a calculation of integral parameters associated with these profiles, and energy spectra of the velocity signal inside the boundary layer. Evidence of the growth of boundary layer velocity fluctuations within a range of frequencies in the laminar separation and transition regions has been found in these spectral measurements. Results have shown that the peak frequencies measured in the velocity spectra for the instability region agree with the most amplified wave number and frequency scaling predicted by linear stability theory for these inflectional profiles. Additionally, the maximum measured growth rates at this peak frequency correlate with growth rates calculated from similarly shaped Falkner-Skan profiles at the corresponding frequency of maximum amplification. This agreement between experimental and theoretical peak frequencies and growth rates was confirmed for the range of Reynolds numbers and for airfoil incidence ranging from zero lift to stall.

The Effects of Turbulence and Stator/Rotor Interactions on Turbine Heat Transfer: Part I—Design Operating Conditions

Abstract: A combined experimental and analytical program was conducted to examine the effects of inlet turbulence, stator-rotor axial spacing, and relative circumferential spacing of first and second stators on turbine airfoil heat transfer. The experimental portion of the study was conducted in a large-scale (approximately 5X engine), ambient temperature, stage-and-a-half rotating turbine model. The data indicate that while turbine inlet turbulence can have a very strong impact on the first stator heat transfer, its impact in downstream rows is minimal. The effects on heat transfer produced by relatively large changes in stator/rotor spacing or by changing the relative row-to-row circumferential positions of stators were very small. Analytical results consist of airfoil heat transfer distributions computed with a finite-difference boundary layer code.

Effect of simulated glaze ice on a rectangular wing

Abstract: Experimental measurements of the effect of simulated glaze ice on a three-dimensional wing are presented. A semispan wing of effective aspect ratio five was mounted from a splitter plate in the OSU subsonic wind tunnel. The model has a straight, untwisted rectangular platform, and uses a NACA 0012 airfoil section. Surface pressures were measured at 5 semispan locations and a total-pressure wake-survey probe was used on the model centerline. The section lift and drag data from the model centerline compared well to earlier two-dimensional data. These data show a large drag and maximum lift penalty due to the simulated glaze ice. Three-dimensional span-load data compare well to computational results.

Fully integrated aerodynamic/dynamic optimization of helicopter rotor blades

Abstract: A fully integrated aerodynamic/dynamic optimization procedure is described for helicopter rotor blades. The procedure combines performance and dynamic analyses with a general purpose optimizer. The procedure minimizes a linear combination of power required (in hover, forward flight, and maneuver) and vibratory hub shear. The design variables include pretwist, taper initiation, taper ratio, root chord, blade stiffnesses, tuning masses, and tuning mass locations. Aerodynamic constraints consist of limits on power required in hover, forward flight and maneuvers; airfoil section stall; drag divergence Mach number; minimum tip chord; and trim. Dynamic constraints are on frequencies, minimum autorotational inertia, and maximum blade weight. The procedure is demonstrated for two cases. In the first case, the objective function involves power required (in hover, forward flight and maneuver) and dynamics. The second case involves only hover power and dynamics. The designs from the integrated procedure are compared with designs from a sequential optimization approach in which the blade is first optimized for performance and then for dynamics. In both cases, the integrated approach is superior.

Technique for the prediction of airfoil flutter characteristics in separated flow

Abstract: A solution procedure is described for determining the two-dimensional, one- or two-degree-of-freedom flutter characteristics of arbitrary airfoils at large angles of attack. The same procedure is used to predict stall flutter. This procedure requires a simultaneous integration in time of the solid and fluid equations of motion. The fluid equations of motion are the unsteady compressible Navier-Stokes equations, solved in a body-fitted moving coordinate system using an approximate factorization scheme. The solid equations of motion are integrated in time using an Euler implicit scheme. Flutter is said to occur if small disturbances imposed on the airfoil attitude lead to divergent oscillatory motions at subsequent times. Results for a number of special cases are presented to demonstrate the suitability of this scheme to predict flutter at large mean angles of attack. Some stall flutter applications are also presented.