Showing papers on "Airfoil published in 1985"

The Kutta Condition in Unsteady Flow

Abstract: In several papers published in the first decade of this century, Kutta and Joukowsky independently proposed that the lift on an airfoil at incidence in a steady un separated flow is given by potential-flow theory with the unique value of the circulation that removes the inverse-sQuare-root velocity singularity at the trailing edge. This proposal-tantamount to saying (cf. Batchelor 1967) that in the unsteady start-up phase the action of viscosity is such that, in the ultimate steady motion, viscosity can be explicitly ignored but implicitly incorporated in a single edge condition-is known as the Kutta-Joukowsky hypothesis. Subsequently the name "Kutta condition" (no doubt largely for brevity) has come to be used to connote the removal of a velocity singularity at some distinguished point on a body in unsteady flow. 1 The condition has recently been applied to unsteadiness in a variety of mean configurations. These include trailing-edge flows with the same and with different flows on the two sides of the body upstream of the edge, attached leading-edge flows, and grossly separated flows past bluff bodies. Imposition of a Kutta condition on unsteady perturbations to one of these mean flows has a variety of physical ramifications. It represents the mechanism by which both the lift is changed and the amplitude and directivity of a sound field are modified. It is the analytical step that in many cases describes the conversion-almost total-of acoustic energy in an incident sound wave to energy of vortical motion on a shear layer ; on

Loss Reduction in Axial-Flow Compressors Through Low-Speed Model Testing

Abstract: A systematic procedure for reducing losses in axial-flow compressors is presented. In this procedure, a large, low-speed, aerodynamic model of a high-speed core compressor is designed and fabricated based on aerodynamic similarity principles. This model is then tested at low speed where high-loss regions associated with three-dimensional endwall boundary layers flow separation, leakage, and secondary flows can be located, detailed measurements made, and loss mechanisms determined with much greater accuracy and much lower cost and risk than is possible in small, high-speed compressors. Design modifications are made by using custom-tailored airfoils and vector diagrams, airfoil endbends, and modified wall geometries in the high-loss regions. The design improvements resulting in reduced loss or increased stall margin are then scaled to high speed. This paper describes the procedure and presents experimental results to show that in some cases endwall loss has been reduced by as much as 10 percent, flow separation has been reduced or eliminated, and stall margin has been substantially improved by using these techniques.

Flight Mechanics of a Dragonfly

Abstract: SUMMARY The steady slow climbing flight of a dragonfly, Sympetrum frequens, was filmed and analysed. By using the observed data, the mechanical characteristics of the beating wings were carefully analysed by a simple method based on the momentum theory and the blade element theory, and with a numerical method modified from the local circulation method (LCM), which has been developed for analysing the aerodynamic characteristics of rotary wings. The results of calculations based on the observed data show that the dragonfly performs low speed flight with ordinary airfoil characteristics, instead of adopting an abnormally large lift coefficient. The observed phase advance of the hindwing, Adi — 80° can be fully explained by the present theoretical calculation. Similarly, the spanwise variation of the airloading and the time variations of the horizontal force, vertical force, pitching moment and torque or power can be definitely estimated within a reasonable range of accuracy in comparison with the flight data. The distribution of loading between the fore and hind pairs of wings is also clarified by the calculations.

The influence of laminar separation and transition on low Reynolds number airfoil hysteresis

Abstract: An experimental study of the Lissaman 7769 and Miley MO6-13-128 airfoils at low chord Reynolds numbers is presented. Although both airfoils perform well near their design Reynolds number of about 600,000, they each produce a different type of hysteresis loop in the lift and drag forces when operated below chord Reynolds numbers of 300,000. The type of hysteresis loop was found to depend upon the relative location of laminar separation and transition. The influence of disturbance environment and experimental procedure on the low Reynolds number airfoil boundary layer behavior is also presented. The use of potential flow solutions to help predict how a given airfoil will behave at low Reynolds numbers is also discussed.

Static and dynamic pressure measurements on a NACA 0012 airfoil in the Ames High Reynolds Number Facility

Abstract: The supercritical flows at high subsonic speeds over a NACA 0012 airfoil were studied to acquire aerodynamic data suitable for evaluating numerical-flow codes. The measurements consisted primarily of static and dynamic pressures on the airfoil and test-channel walls. Shadowgraphs were also taken of the flow field near the airfoil. The tests were performed at free-stream Mach numbers from approximately 0.7 to 0.8, at angles of attack sufficient to include the onset of buffet, and at Reynolds numbers from 1 million to 14 million. A test action was designed specifically to obtain two-dimensional airfoil data with a minimum of wall interference effects. Boundary-layer suction panels were used to minimize sidewall interference effects. Flexible upper and lower walls allow test-channel area-ruling to nullify Mach number changes induced by the mass removal, to correct for longitudinal boundary-layer growth, and to provide contouring compatible with the streamlines of the model in free air.

An Experimental Investigation of an Airfoil Undergoing Large-Amplitude Pitching Motions

Abstract: FLOW visualization and near-surface hot-wire experiments were performed in the U.S.A.F Academy Aeronautics Laboratory subsonic wind tunnel on an airfoil undergoing large-amplitude pitching motions about its quarter chord. The experiments were conducted using a NACA 0015 airfoil at an airfoil Reynolds number of 45,000. Two cases are presented in which the angular pitching rate a is maintained constant during the motion. These two cases represent two different nondimensional pitching rates a+, where ot+ is equal to 6; nondimensionalized by the chord c and the freestream velocity U^ (a + ^ac/U^). Data for the two cases where values of a+ are equal to 0.2 and 0.6 show the dramatic effect of pitch rate on flow structure. Largescale vortical structures are seen in both cases at high angles of attack but appear much later and are of a different form for the case with the larger a+ value. These structures are very energetic, producing reverse flow velocities near the airfoil surface of 1.0-2.1 times the freestream velocity.

Flutter and divergence aeroelastic characteristics for composite forward swept cantilevered wing

Abstract: An analytical investigation was conducted to determine the aeroelastic flutter and divergence behavior of a cantilevered, composite, forward swept rectangular wing. The influence due to the variation in the bendingtorsion stiffness coupling of the tailored wing on the flutter and divergence critical dynamic pressure is analyzed. The analytical approach utilizes the incompressible two-dimensional unsteady aerodynamic strip theory. Flutter and divergence velocities were obtained by using an optimization procedure that solves exactly the coupled bending-torsion equations for a cantilevered swept wing. The results indicate that the flutter and divergent of a fixed-root wing involve a compromise, since the bending-torsion stiffness that maximizes the flutter velocity tends to minimize the divergent speed and vice versa.

Low Reynolds Number Vehicles

Abstract: : Recent interest in a wide variety of low Reynolds number configurations has focused attention on the design and evaluation of efficient airfoil sections at chord Reynolds numbers from about 100,000 to about 1,000, 000. These configurations include remotely piloted vehicles (RPV's) at high altitudes, sailplanes, ultra-light man-carrying/man-powered aircraft, mini-RPVs at low altitudes and wind turbines/propellers. A study is presented of the present status and future possibility of airfoil design and evaluation at subcritical speeds to meet the needs for these applications. Although the design and evaluation techniques for airfoil sections above chord Reynolds numbers of 500,000 is reasonably well developed, serious problems related to boundary layer separations and transition have been encountered below RC = 500,000. Presently available design and analysis methods need to improve their criteria for laminar separation, transition, and turbulent separation. Improved mathematical models of these complex phenomena require additional, very careful experimental studies. Because of the sensitivity of the low Reynolds number airfoil boundary layer to free stream and surface-generated disturbances, definitive experiments are very difficult. Also the physical quantities measured (i.e., pressure difference and drag forces etc.) are very small and the accuracy of such measurements depends on the method used. The results from numerous experimental studies are presented to illustrate the type of difficulties encountered.

Large airfoil structure and method for its manufacture

Abstract: An airfoil particularly suited for use in rotor blades of large wind-drivenower plants is formed in three modular sections including a nose or leading section, a spar section and a trailing section, each separately formed and subsequently assembled. In the method of the invention the spar section is formed in two separate mold units which are configured to form part of the upper and lower aerodynamic profile of the airfoil and the two spar section units are trimmed along a junction plane and then joined together at the junction plane to form the spar section of the airfoil.

Analysis of transitional separation bubbles on infinite swept wings

Abstract: A previously developed two-dimensional local inviscid-viscous interaction technique for the analysis of airfoil transitional separation bubbles, ALESEP (Airfoil Leading Edge Separation), has been extended for the calculation of transitional separation bubbles over infinite swept wings. As part of this effort, Roberts' empirical correlation, which is interpreted as a separated flow empirical extension of Mack's stability theory for attached flows, has been incorporated into the ALESEP procedure for the prediction of the transition location within the separation bubble. In addition, the viscous procedure used in the ALESEP techniques has been modified to allow for wall suction. A series of two-dimensional calculations is presented as a verification of the prediction capability of the interaction techniques with the Roberts' transition model. Numerical tests have shown that this two-dimensional natural transition correlation may also be applied to transitional separation bubbles over infinite swept wings. Results of the interaction procedure are compared with Horton's detailed experimental data for separated flow over a swept plate which demonstrates the accuracy of the present technique. Wall suction has been applied to a similar interaction calculation to demonstrate its effect on the separation bubble. The principal conclusion of this paper is that the prediction of transitional separation bubbles over two-dimensional or infinite swept geometries is now possible using the present interacting boundary layer approach.

Nonlinear problems in flight dynamics involving aerodynamic bifurcations

Abstract: Aerodynamic bifurcation is defined as the replacement of an unstable equilibrium flow by a new stable equilibrium flow at a critical value of a parameter A mathematical model of the aerodynamic contribution to the aircraft's equations of motion is amended to accommodate aerodynamic bifurcations Important bifurcations such as, the onset of large-scale vortex-shedding are defined The amended mathematical model is capable of incorporating various forms of aerodynamic responses, including those associated with dynamic stall of airfoils

Characteristics of the flow around conventional and supercritical airfoils

Abstract: Measurements of the mean and fluctuating velocities have been obtained with pressure and hot-wire probes in the attached boundary layers and wakes of two airfoil models at a low Mach number. The first model is a conventional airfoil at zero incidence and the second an advanced supercritical airfoil at an angle of attack of 4°. The mean-flow and Reynolds-stress data and related quantities are presented with emphasis on the trailing-edge region. The results indicate that the flow around the conventional airfoil is a minor perturbation of a symmetric flat-plate flow with small wake curvature and weak viscous–inviscid interaction. The flow around the supercritical airfoil is in considerable contrast with strong streamwise pressure gradients, non-negligible normal pressure gradients, and large surface and streamline curvatures of the trailing-edge flow. The near wake is strongly curved and intense mixing occurs between the retarded upper-surface boundary layer and strongly accelerated lower-surface boundary layer.

Characteristics of a trailing-edge flow with turbulent boundary-layer separation

Abstract: Experimental techniques, including flying-hot-wire anemometry, have been used to determine the pressure and velocity characteristics of a flow designed to simulate the trailing-edge region of an airfoil at high angle of attack. Emphasis is placed on the region of recirculating flow and on the downstream wake. It is shown that the effect of this recirculation is large even though the details of the flow within it may be unimportant. Normal stresses and cross-stream pressure gradients are important immediately upstream and downstream of the recirculating flow and are associated with strong streamline curvature. The relative importance of the terms in the transport equations for mean momentum and turbulence energy are quantified and the implications for procedures which solve potential-flow and boundary-layer equations and for alternative calculation methods are discussed.

Total pressure loss in vortical solutions of the conical Euler equations

Abstract: A technique for the solution of the conically self-similar form of the Euler equations is described. Solutions for the flow past a flat-plate delta wing at angle of attack are presented. These solutions show strong leading edge vortices with large total pressure losses in the cores. A study of the effects of various computational parameters on the total pressure loss is made. An explanation for the cause of the total pressure loss is presented. It is shown to be consistent with the results for both a quasi-one-dimensional model problem and the conically self-similar flow past the flat-plate delta wing.

An Eleven Parameter Axial Turbine Airfoil Geometry Model

Abstract: The mathematical derivation, and FORTRAN code, of a comprehensive but easy to use geometry model for axial flow turbine nozzles and rotors is presented. To uniquely define an airfoil on a cylinder the aerodynamicist need only specify the number of blades, and at each radius of interest: the axial and tangential chord, throat, uncovered turning, leading and trailing edge radii, inlet and exit blade angles, and inlet wedge angle. Default values exist for six of these geometric variables, which proves useful when starting a design. Both the suction and the pressure surfaces are described entirely by analytical functions. Sample airfoils are included that demonstrate the effect of each parameter upon blade shape.Copyright © 1985 by ASME

Effects of Suction on Shock/Boundary-Layer Interaction and Shock-Induced Separation

Abstract: An experimental investigation of the effects of local boundary-layer suction on shock/boundar y-layer interaction and shock-induced separation has been conducted in the DFVLR 1 m x 1 m transonic wind tunnel utilizing an advanced transonic airfoil. Three different methods of suction were applied in the shock region. Their effectiveness in comparison to the basic closed-surface airfoil will be evaluated from surface pressure distribution, wake, and boundary-layer measurements. It will be shown that local boundary-layer suction in the shock region delays the development of shock-induced separation and considerably improves the overall aerodynamic characteristics. Moreover, two of the configurations investigated, viz., a double slot and a perforated strip with a cavity underneath showed, even without suction, a most favorable "passive" effect on shock/boundary-layer interaction and the overall flow development, thus offering a very promising means for extending the range of applicability of transonic airfoils.

Scaling of airfoil self-noise using measured flow parameters

Abstract: Data from an airfoil broadband self-noise study are reported. Attention here is restricted to two-dimensional models at zero angle of attack to the flow. The models include seven NACA 0012 airfoil sections and five flat plate sections with chordlengths ranging from 2.54 to 60.96 cm. Testing parameters include flow velocity to 71.3 m/s and boundary-layer turbulence through natural transition and by tripping. Detailed aerodynamic measurements are conducted in the near-wake of the sharp trailing edges. The noise spectra of the self-noise sources are determined by the use of a cross-spectral technique. The acoustic data are normalized using the measured aerodynamic parameters in order to evaluate a commonly used scaling law. An examination of the Reynolds number dependence of the normalized overall levels has revealed a useful scaling result. This result appears to quantify the transition between turbulent boundary-layer trailing-edge noise and laminar boundary-layer vortex shedding noise.

Local heat-transfer measurements on a large scale-model turbine blade airfoil using a composite of a heater element and liquid crystals

Abstract: Local heat transfer coefficients were experimentally mapped along the midchord of a five-time-size turbine blade airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a mylar sheet with a layer of cholesteric liquid crystals, that change color with temperature, and a heater sheet made of a carbon-impregnated paper, that produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. The local heat transfer coefficients are presented for Reynolds numbers from 2.8 x 10 to the 5th power to 7.6 x 10 to the 5th power. Comparisons are made with analytical values of heat transfer coefficients obtained from the STANS boundary layer code. Also, a leading edge separation bubble was revealed by thermal and flow visualization.

Dynamic ground effects on a two-dimensional flat plate

Abstract: The effect of time-variant vortex shedding is simulated by a sequence of discrete vortices convecting downstream in the wake of a two-dimensional flat plate whose lifting condition is modeled by means of the quasi-vortex lattice method. The boundary condition of this problem is specified in such a way that the tangency condition on the surface of the flat plate is satisfied; the boundary condition also takes into account the effect of airfoil motion relative to the ground. Significant lift changes are shown to occur, due to the dynamic ground effect, that are crucial in aircraft takeoff and landing transitions.

An efficient method for computing unsteady transonic aerodynamics of swept wings with control surfaces

Abstract: A transonic equivalent strip (TES) method was further developed for unsteady flow computations of arbitrary wing planforms. The TES method consists of two consecutive correction steps to a given nonlinear code such as LTRAN2; namely, the chordwise mean flow correction and the spanwise phase correction. The computation procedure requires direct pressure input from other computed or measured data. Otherwise, it does not require airfoil shape or grid generation for given planforms. To validate the computed results, four swept wings of various aspect ratios, including those with control surfaces, are selected as computational examples. Overall trends in unsteady pressures are established with those obtained by XTRAN3S codes, Isogai's full potential code and measured data by NLR and RAE. In comparison with these methods, the TES has achieved considerable saving in computer time and reasonable accuracy which suggests immediate industrial applications.

Surface wetting effects on a laminar flow airfoil in simulated heavy rain

Abstract: Wind tunnel experiments have been conducted on a natural laminar flow airfoil in a simulated heavy rain of 440 mm/h at a Reynolds number of 310,000 to assess the effect of surface wettability on the performance degradation due to rain. A significant loss of performance was observed for each of the three surfaces tested with the nonwettable, waxed surface being the most degraded (-75% reduction in maximum L/D) and the incompletely wettable epoxy gel coat being the least (-45% reduction). Accompanying the L/D loss was an effective reduction in angle of attack of up to 2 deg resulting from a downward translation of the CL polar. In photographic observations, the runback water layer was found to bead on the wax surface and sheet on the wettable surfaces. The strong dependence on surface wettability of both the airfoil performance and the water behavior indicates that the degradation due to heavy rain is primarily a result of the roughening of the airfoil surface by the runback water layer. The observed performance loss could only be partially emulated by causing premature transition from a laminar to a turbulent boundary layer. Nomenclature CL =lift coefficient, based on chord CD =drag coefficient, based on chord D = drop or bead diameter h = height of water layer L/D = lift-to-drag ratio U = flow velocity We = Weber number OL = angle of attack 0 = contact angle p = air density a = surface tension

The Gortler Instability on an Airfoil

Abstract: Gortler vortices arise in boundary layers along concave surfaces due to centrifugal effects. This paper presents some results of an experiment conducted to study the development of these vortices on an airfoil with a pressure gradient in the concave region where an attached laminar boundary layer was insured with suction through a perforated panel. The sublimating chemical technique was used to visualize Gortler vortices and the velocity field was measured by laser velocimetry. The vortex wavelength clearly varied with Gortler number as predicted by linear theory. Both flow visualization and velocity measuremetns indicated vortex damping in the convex zone. Secondary instability was observed at the higher Gortler numbers.

Airfoil design for Reynolds numbers between 50,000 and 500,000

Abstract: The design of airfoils for flows with Re of 50,000-500,000 requires consideration of laminar separation bubbles. A design approach is discussed which specifies the angle of attack at which the potential flow velocity is to be constant at each segment of the airfoil. The velocity gradient is controlled by introducing a pressure recovery function at the trailing edge. Boundary layer stability decreases with rising Re, although an upper Re value can be identified, below which the boundary layer will be stable. Adverse pressure gradients are associated with the shape parameter of the velocity profile, whose rise in value decreases stability. Transition displays similar relationships to the shape parameter. The most frequent feature of separation is the appearance of a separation bubble.

Viscous/inviscid analysis of curved sub- or supersonic wall jets

Abstract: T features of a viscous/inviscid curved wall jet model, under development for application to the analysis of circulation-controlled airfoil flowfields, are described. The model solves the surface-oriented parabolized Navier-Stokes equations using pressure-split methodology for subsonic wall jets and shock-capturing methodology for supersonic wall jets. A hybrid two-layer turbulence model is employed that combines a damped VanDriest inner layer formulation with a curvature-corrected, two-equation ke model outer region formulation. Procedures for performing the strongly interactive coupling of the wall jet solution with an external potential flow solver are discussed. Preliminary calculations for simple wall jets are shown to compare quite favorably with both mean flow and turbulent measurements.

Interrelated Effects of Pitch Rate and Pivot Point on Airfoil Dynamic Stall

Abstract: : Experimental investigations were conducted to study energetic dynamic stall vortices and the associated unsteady aerodynamics generated by a pitching NACA 0015 airfoil. The airfoil model was pitched from zero degrees to 60 degrees at constant pitch rates of 460 deg/sec, 920 deg/sec, and 1380 deg/sec about its quarter-chord, half-chord, and three quarter-chord positions. Extensive 35mm still photographs and 16mm high-speed movies, both employing smoke wire flow visualization, visually documented the initiation and development of the time dependent dynamic stall phenomena. Hot-wire anemometry measurements were also made which provided for more quantitative analysis of the unsteady separated flowfields. Pitch rate and pivot point were shown to have interrelated effects on the development of the dynamic stall flowfield. In many cases similar 'looking' flowfields were generated by different combinations of pitch rate and pivot point. However, significant differences were observed in the near-surface velocity profiles. (edc)

Numerical solution of the full-potential equation for rotors and oblique wings using a new wake model

Abstract: A three-dimensional, full-potential, quasi-steady code TFAR1 is proposed for calculating the transonic flow past a lifting helicopter rotor blade and oblique wing. The TFAR1 uses a two-dimensional nonlinear wake-model that allows a jump in velocity potential to propagate with the local fluid flow in the wake. Rotor calculations were made for a single blade at an advance ratio of 0.3, a rotational tip Mach number of 0.7, and at 0-degree incidence. A 1/7-scale model of the Cobra Operational Load Survey (OLS) rotor blade is calculated, and the pressure distributions are compared to the measurements for azimuth angles 0, 30, 60, 90, 120, and 150 degrees at the 95 percent spanwise station of the OLS blade. Furthermore, an oblique wing with Korn airfoil was calculated at the high transonic free-stream Mach number of 0.9791, zero incidence, and yaw angle of 40 degrees. The TFAR1, coupled with a helicopter performance code CAMRAD (Johnson, 1981), provides a full-potential code for calculating the entire flow field for a multiple-bladed rotor in transonic lifting forward flight.

The buffeting pressure field of a high-aspect-ratio swept wing

Abstract: An experimental study of the fluctuating pressure field of a high-aspect-ratio swept transport-type wing model in transonic buffeting flow is examined. A high degree of similarity between the buffet-related features of the wing flowfield and 2-D transonic flowfields is indicated. At several spanwise stations, unsteady lift produced by the pressure fluctuations is evaluated and found to increase in intensity as local flow separation develops. Results show that although the wing does vibrate, the measured lift fluctuations are driving, and not driven by wing vibrations. The influence of tunnel-flow unsteadiness on the data is considered.

Combination lift thrust device

Abstract: A combination lift and thrust device for increasing the performance of an aircraft by simultaneously reducing drag and augmenting the thrust of a turbojet engine carried within the device. The device comprises a wing of generally airfoil shape having numerous geometrically spaced apertures penetrating its surface, a turbojet engine carried within the wing, an elongated exhaust plenum attached to the turbojet having a number of strategically positioned exhaust nozzles, and a mixing chamber having a forward opening and a rear nozzle also carried within the wing. The mixing chamber forward opening is cooperatively associated with the exhaust nozzles to form an ejector drawing air through the apertures in the wing into the forward opening, mixing the air with the exhaust gases from the turbojet to provide thrust augmentation, and exhausting the air and gas mixture from the rear nozzle of the mixing chamber. The air drawn through the apertures in the wing surface reduces the turbulent boundary layer on the wing thus reducing the aerodynamic drag of the wing.