Showing papers on "Airfoil published in 1984"

Combined riblet and lebu drag reduction system

Abstract: The invention is a system of flow control devices which result in reduced skin friction on aerodynamic and hydrodynamic surfaces The devices cause a breakup of large-scale disturbances in the boundary layer of the flow field Referring to FIGS 1 and 2, the riblet device 10 acts to reduce disturbances near the boundary layer wall by the use of longitudinal striations forming vee-shaped grooves These grooves are dimensional on the order of the wall vortices and turbulent burst dimensions 31 depicted in FIG 3 The large-eddy breakup device 41, depicted in FIGS 4 and 5, is a small strip or airfoil which is suspended in the upper region of the boundary layer Various physical mechanisms cause a disruption of the large-scale vortices The combination of the devices of this invention result in a substantial reduction in skin friction drag

The Sears problem for a lifting airfoil revisited - new results

Abstract: It is shown that for a thin airfoil with small camber and small angle of attack moving in a periodic gust pattern, the unsteady lift caused by the gust can be constructed by linear superposition to the Sears lift of three independent components accounting separately for the effects of airfoil thickness, airfoil camber and non-zero angle of attack to the mean flow. This is true in spite of the nonlinear dependence of the unsteady flow on the mean potential flow of the airfoil. Specific lift formulas are derived and analysed to assess the importance of mean flow angle of attack and airfoil camber on the gust response.

Ice shapes and the resulting drag increase for a NACA 0012 airfoil

Abstract: Experimental measurements of the ice shapes and resulting drag increases were measured in the NASA-Lewis Icing Research Tunnel. The measurements were made over a large range of conditions (e.g., airspeed and temperature, drop size and liquid water content of the cloud, and the angle of attack of the airfoil). The measured drag increase did not agree with the existing correlation. Additional results were given which are helpful in understanding the ice structure and the way it forms, and in improving the ice accretion modeling theories. There are data on the ice surface roughness, on the effect of the ice shape on the local droplet catch, and on the relative importance of various parts of the ice shape on the drag increase. Experimental repeatability is also discussed.

Synchro-vane vertical axis wind powered generator

Abstract: A vertical axis wind powered generator apparatus has airfoil shaped vanes, a positive and synchronous vane orientation system which is controlled by a mechanism located exterior to its rotor, two innovations for improving its aerodynamic efficiency and for increasing the rotary force and horsepower developed by a tall wind generator apparatus used to power a driven machine, and a system for operational control of the device.

Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections

Abstract: Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.

Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio

Abstract: Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.

Flutter Analysis Using Nonlinear Aerodynamic Forces

Abstract: The nonlinear effects of transonic aerodynamic forces on the flutter boundary of a typical section airfoil are studied. The flutter speed dependence on amplitude is obtained by utilizing a novel variation of the describing function method which takes into account the first fundamental harmonic of the nonlinear oscillatory motion. By using an aerodynamic describing function, traditional flutter analysis methods may still be used while including the effects of aerodynamic nonlinearities. Results from such a flutter analysis are compared with those of brute force time-marching solutions. The aerodynamic forces are computed by the LTRAN2 aerodynamic code for an NACA 64A006 airfoil at Mx = 0.86.

Application of the ONERA Model of Dynamic Stall

Abstract: A semiempirical model to predict the unsteady loads on an airfoil that is experiencing dynamic stall, is investigated. The mathematical model is described from an engineering point of view, demonstrates the procedure for obtaining various empirical parameters, and compares the loads predicted by the model with those obtained in the experiment. It is found that the procedure is straightforward, and the final calculations are in qualitative agreement with the experimental results. Comparisons between calculations and measurements also indicate that a decrease in accuracy results when the values of both the reduced frequency and the amplitude of oscillation are large. Potential quantitative improvements in the accuracy of the calculations are discussed for accounting of both the hysteresis in the static data and the effects of stall delay in the governing equations.

Doublet-Point Method for Supersonic Unsteady Lifting Surfaces.

Abstract: A method to predict unsteady aerodynamic forces on lifting surfaces in supersonic flow is presented. The wing is divided into small segments in which the lift force is expressed by a single-point doublet of the acceleration potential. This is the same concept as the doublet-point method developed by the authors for subsonic flows. In order to avoid sensitiveness to the Mach number, the upwash due to the point doublet is calculated by averaging over small areas. The integration is done analyticaly so that it requires no numerical quadrature. Pressure distributions are directly obtained as the unknowns of the algebraic equation. The results are compared with those obtained by other methods for various wing geometries, including the AGARD wing-tail configuration.

Cooling arrangement for airfoil stator vane trailing edge

Abstract: A hollow, airfoil shaped stator vane 10, for a combustion turbine, is provided in its trailing edge portion with staggered rows of pin fins 20, 22 and 24 and with longitudinal ribs 26 which provide the required stiffness for the thin wall design of the vane, the ribs being provided with protuberances 36 in those locations where the turbulence inducing pin fins are displaced by the presence of the ribs.

Porosity effect on supercritical airfoil drag reduction by shock wave/boundary layer control

Abstract: An investigation of the passive shock wave/boundary layer control for reducing the drag of 14 percent-thick supercritical airfoil was conducted in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel at transonic Mach numbers. Various porous surfaces with a cavity beneath it was positioned on the area of the airfoil, mounted on the test section bottom wall, where the shock wave occurs. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a uniform porosity surface the normal shock wave for solid surface was changed to a lambda shock wave, and the wake impact pressure data indicated an appreciable drag reduction at transonic Mach numbers. For a free stream Mach number of 0.81 the profile drag coefficient for the airfoil top surface with uniform porosity was 46 percent lower than for the solid surface airfoil.

Wing flap mechanism

Abstract: A carrier unit for chordwise extending and rotatably positioning an auxiliary airfoil mounted adjacent to the trailing edge of a relatively fixed main airfoil of an airplane. When the auxiliary airfoil is in a fully retracted and stowed position, the actuation and positioning linkage mechanism of the carrier unit is completely housed within the combined auxiliary and main airfoil envelope with no external protrusions or fairings. For auxiliary airfoil extension, to increase the overall chord plane by approximately eight percent, a first set of four-bar linkages (13, 20, 22, 24) is utilized comprising: a beam member (22) supported at a forward portion by a pair of approximately parallel links (20, 24) pivoted to wing structure (13) and swingable chordwise in a generally parallel relationship for translatably shifting the auxiliary airfoil which is pivoted to a rearward portion of the beam member. For changing the auxiliary airfoil angle-of-incidence relative to the main airfoil, one or more sets of four-bar linkage (22, 24, 28, 11) are integfatadtwith the first set for a unified rotation programming and extension driva that functionsto produes an aerodynamic slot opening when the atrxiliery airfoil is extended to takeoff and landing positions whereat it is deflected up to approximately forty degrees relative to the main airfoil reference chord plane. For an aerodynamic braking action during landing roll-out, the geometrical relationship of the sets of four-bar linkages functions to deflect the auxiliary airfoil chord plane greater than forty degrees, to approximately ninety degrees or more relative to the main airfoil reference chord plane.

Minimum induced drag of wings with curved planform

Abstract: Conclusions This study has shown that a small airfoil probe, consisting of a small canard wing mounted appropriately on an airframe and properly tapped, can serve as a viable alternative as a probe for angle-of-attack sensing on aircraft. An NACA 0012 airfoil section was used in wind tunnel tests in this study, and differential pressure coefficients greater than 3.0 at high angles of attack were achieved. These coefficients are an improvement by a factor of 2.0-3.0 over comparable coef- ficients obtained from hemispherical probes. References

A Calculation Method of Leading-Edge Separation Bubbles

Abstract: The positive pressure gradient, downstream of the suction peak at the leading edge of an airfoil at incidence, may, under some conditions, induce large perturbations on the general pattern of the flow.

Unsteady Flow Concepts for Dynamic Stall Analysis

Abstract: It is well established that there is a strong coupling between airfoil motion and boundary layer separation with attendant vortex shedding Unsteady flow mechanisms that influence this dynamic stall event have been described previously Until now sufficient information has not been available to determine the relative im portance of various unsteady flow effects, such as the time varying inviscid pressure gradient and the unsteady viscous boundary condition at the wall, the moving wall effect Recent experimental results provide the needed information, revealing how the mode of oscillation for the airfoil determines which unsteady flow effect will dominate Nomenclature chord length frequency dynamic overshoot parameter, Eqs. (5 7) section lift, coefficient c, = l/(p^ Ui/2)c Mach number section pitching moment, coefficient cm = mp c / Ka / M m q -pitch rate Re = Reynolds number based on chord length, =

Acceleration to a steady state for the Euler equations

Abstract: A multi-stage Runge-Kutta method is analyzed for solving the Euler equations exterior to an airfoil. Highly subsonic, transonic and supersonic flows are evaluated. Various techniques for accelerating the convergence to a steady state are introduced and analyzed.

Post stall studies of untwisted varying aspect ratio blades with an NACA 4415 airfoil section. Part I

Abstract: Wind Turbine blades operate over a wide angle of attack range. Unlike aircraft, a wind turbine's angle of attack range extends deep into stall where the three dimensional performance characteristics of air foils are not generally known. Peak power predictions upon which wind turbine components are sized, depend on a good understanding of a blade's post stall characteristics. The results of the wind tunnel investigation of untwisted, constant chord blades having four aspect ratios, with an NACA 4415 series airfoil section, at angles of attack ranging from -10 to 100 degrees are discussed. Tests were conducted for aspect ratios of 6, 9, 12 and infinity at four Reynolds numbers ranging from one-quarter million to one million. The results on the same family of airfoil section but with varying thickness ratio are given. Results of force and pitching moment measurements over the angle of attack range for all combinations of Reynolds numbers and aspect ratios, and the effects of boundary layer tripping, are presented. Both initial and secondary stall are presented. The maximum drag coefficient is found to occur an an angle of attack of 90 degrees. The pitching moment is unstable beyond stall. The lift and post-stall drag-coefficients decrease with decreasing aspect ratio. The boundary layer tripping is observed to decrease the lift curve slope and stalling angle of attack. The drag coefficient (with tripping) is significantly affected only at low aspect ratio. Beyond secondary stall, the lift to drag ratio is independent of aspect ratio. The maximum lift to drag ratio for the infinite aspect ratio blade is roughly twice that of the blade with an aspect ratio of six. This effect is independent of Reynolds number in the range studied.

Electro-impulse de-icing system for aircraft

Abstract: What is proposed is an electro-impulse de-icing device for use in an airfoil of an aircraft that includes a system having an energy storage device connected via a triggering device to a set of coils each embodying a coil construction in the form of a spirally wound ribbon coil member whose planar dimension is pressed against an interior surface in the leading edge of the airfoil, and wherein the coil sets are arranged in a linear manner along the long axis of the airfoil and are energized simultaneously and sequentially via an interlocking pattern from the wing tips or from the fuselage of the aircraft with each coil set being triggered twice from about three to five seconds apart.

Aerodynamic Loss Penalty Produced by Film Cooling Transonic Turbine Blades

Abstract: Aerodynamic loss measurements are presented for a state-of-the-art film cooled transonic gas turbine rotor blade tested in a two-dimensional cascade. A mixture of carbon dioxide and air, which correctly simulated engine coolant-to-mainstream density ratio and blowing rate, was ejected from each of five individual cooling hole rows in the aerofoil suction surface. The temperature of the coolant was equal to the cascade inlet stagnation temperature. The dependence of blade row efficiency and turning on outlet Mach number, blowing rate, and coolant-to-mainstream density ratio was investigated. Measured surface static pressure distributions were compared with time-marching predictions for both the datum aerofoil and film cooled blades. Detailed suction surface boundary layer measurements both upstream and downstream of a cooling film were compared with available differential calculation procedures. Unexpectedly, films downstream of the throat, even at blowing rates near unity, did not generate significantly higher losses compared to prethroat suction surface films on this aerofoil.

Airfoil tip vortex formation noise

Abstract: Spectral data are presented for the noise produced due to the turbulent three-dimensional vortex flow existing near the rounded tip of lifting airfoils. The results are obtained by the comparison of sets of two- and three-dimensional test data for different airfoil model sizes, angles of attack, and tunnel flow velocities. Microphone cross-correlation and cross-spectral methods were used to determine the radiated noise. Corrections were made for tunnel shear layer and source directivity effects. Interpretation of the results are aided by a three-dimensional flow analysis developed for this study which determines open tunnel and finite aspect ratio corrections heretofore neglected in tip vortex studies. Hot wire measurements were made in the tip vortex formation region for the specification of governing flow parameters. The spectral data is normalized in a format considered most useful for subsequent quantitative prediction of this noise mechanism for practical systems such as helicopter rotors. Comparison is made to the analysis of George and Chou. A recommended prediction method is given.

Airfoil shape and thickness effects on transonic airloads and flutter

Abstract: A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small-perturbation potential equation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with a corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant. Nomenclature a = pitch axis location, referenced to midchord, in semichords b — semichord length c — chord length c/ = lift coefficient c/h = lift coefficient due to plunge c/ = lift coefficient due to pitch c^ = moment coefficient about c/4 cm — moment coefficient due to plunge cm = moment coefficient due to pitch Cp — pressure coefficient g — structural damping coefficient h — plunge displacement in semichords hj = plunge amplitude in semichord k — reduced frequency, bul V M — freestream Mach number m = airfoil mass per unit span ra — radius of gyration, referenced to pitch axis, in semichords

Numerical simulation of the interaction of a vortex with stationary airfoil in transonic flow

Abstract: A perturbation form of an implicit conservative, noniterative numerical algorithm for the two-dimensional thin layer Navier-Stokes and Euler equations is used to compute the interaction flow-field of a vortex with stationary airfoil. A Lamb-like analytical vortex having a finite core is chosen to interact with a thick (NACA 0012) and a thin (NACA 64A006) airfoil independently in transonic flow. Two different configurations of vortex interaction are studied, viz., (1) when the vortex is fixed at one location in the flowfield, and (2) when the vortex is convecting past the airfoil at freestream velocity. Parallel computations of this interacting flowfield are also done using a version of the Transonic Small Disturbance Code (ATRAN2). A special treatment of the leading edge region for thin airfoils is included in this code. With this, the three methods gave qualitatively similar results for the weaker interactions considered in this study. However, the strongest interactions considered proved to be beyond the capabilities of the small disturbance code. The results also show a far greater influence of the vortex on the airfoil flowfield when the vortex is stationary than when it is convecting with the flow.

Airfoil inspection method

Abstract: An improved inspection method determines how closely the cross sectional configuration of an airfoil corresponds to a reference cross sectional configuration. When an airfoil is to be inspected, the airfoil is measured and the mean camber line is determined. The mean camber lines of the reference and measured airfoils are shifted relative to each other through a plurality of relationships. A best fit relationship is determined in which the measured and reference airfoils overlie each other to a maximum extent possible. The best fit relationship can be determined by comparing points on the major side surfaces of the measured and reference airfoils. Alternatively, the best fit relationship can be determined by comparing points on the mean camber lines of the measured and reference airfoils.

Application of transonic codes to aeroelastic modeling of airfoils including active controls

Abstract: A study is performed using aeroelastic modeling to investigate the stability behavior of airfoils in small-disturbance transonic flow. Two conventional airfoils, NACA 64.A006 and NACA 64A010, and a supercritical airfoil, MBB A-3, are considered. Three sets of unsteady aerodynamic data are computed using three different transonic codes (LTRAN2-NLR, LTRAN2-HI, and USTS) for comparison purposes. Stability results obtained using a constant matrix, state-space, aeroelastic model are presented in a root-locus format. Use of the state-space model is demonstrated through application to flutter suppression using active controls. Aeroelastic effects due to simple, constant gain, partial feedback, control laws that utilize displacement, velocity, and acceleration sensing are studied using a variety of control gains. Calculations are also performed using linear subsonic aerodynamic theory to reveal the differences between including and not including transonic effects in the aeroelastic model. Aeroelastic stability behavior of these airfoils is physically interpreted and discussed in detail.

Tailored airfoils for Vertical Axis Wind Turbines

Abstract: The evolution of a family of airfoil sections designed to be used as blade elements of a vertical axis wind turbine (VAWT) is described. This evolution consists of extensive computer simulation, wind tunnel testing and field testing. The process reveals that significant reductions in system cost-ofenergy and increases in fatigue lifetime may be expected for VAWT systems using these blade elements.

Bifurcation Analysis of Aircraft Pitching Motions About Large Mean Angles of Attack

Abstract: Bifurcation theory is used to analyze the nonlinear dynamic stability characteristics of an aircraft'subject to single-degree-of-freedom pitching-motion perturbations about a large mean angle of attack. The requisite aerodynamic information in the equations of motion can be represented in a form equivalent to the response to finite-amplitude pitching oscillations about the mean angle of attack. It is shown how this information can be deduced from the case of infinitesimal -amplitude oscillations. The bifurcation theory analysis reveals that when the mean angle of attack is increased beyond a critical value at which the aerodynamic damping vanishes, new solutions representing finite-amplitude periodic motions bifurcate from the previously stable steady motion. The sign of a simple criterion, cast in terms of aerodynamic properties, determines whether the bifurcating solutions are stable (supercritical) or unstable (subcritical). For flat-plate airfoils flying at supersonic/hypersonic speed, the bifurcation is subcritical, implying either that exchanges of stability between steady and periodic motion are accompanied by hysteresis phenomena, or that potentially large aperiodic departures from steady motion may develop.

Trailing edge device for an airfoil

Abstract: The invention is a small wedge shaped flap (20) for attachment to or near the trailing edge (14) of a airfoil (10) which improves the coefficient of lift and reduces the coefficient of drag providing an overall increase in fuel ecomony at cruise conditions. In detail the wedge shaped flap (20) has a downward height (26) of between 0.5 percent to 1.5 percent of the chord (16) of the airfoil (10) and has an included angle (30) to the chord of between 15° and 45°. The wedge flap is preferably placed at distance (24) of between 0 to 1.0 percent of the chord (16) from the trailing edge (14).

Aerodynamic coefficients and longitudinal stability of sail aerofoils

Abstract: This paper extends previous theoretical work on inextensible sails in two ways. First, the asymptotic expansion of sail camber in terms of the small angle of attack α is continued to include terms of order α3, because only at this order can one obtain information about the longitudinal static stability of a sail aerofoil. Secondly, to describe aerofoils such as those found in pterodactyl or bat wings, we also consider pretensioned membranes which acquire camber by stretching the surface material, or bending the supporting structure. Approximate formulae are derived for the aerodynamic coefficients in the limits of large and small tension, and the effect of sail and structural flexibility on these coefficients is discussed.