Showing papers on "Airfoil published in 1993"

Oscillatory Blowing: A Tool to Delay Boundary-Layer Separation

Abstract: The effects of oscillatory blowing as a means of delaying separation are discussed. Experiments were carried out on a follow, flapped NACA 01115 airfoil equipped with a two-dimensional slot over the hinge of the flap. The flap extended over 25% of the chord and was detected at angles as high as 40 deg. The steady blowing momentum coefficients could be varied independently of the amplitudes and frequencies of the superimposed oscillations. The modulated blowing was a major factor in improving the performance of the airfoil at much lower energy inputs than was hitherto known. Optimum benefits in performance were obtained at reduced frequencies, based on the flap chord, of an order of unity. Significant increase in lift as well as cancellation of form drag were observed

Flowfield measurements about a multi-element airfoil at high Reynolds numbers

Abstract: This paper describes experimental data obtained with a multi-element airfoil at flight Reynolds numbers and lift coefficients including Clmax. The wind tunnel test was conducted in the NASA Langley Low Turbulence Pressure Tunnel as part of a cooperative effort between McDonnell Douglas Aerospace and NASA Langley. The airfoil model is a supercritical design configured with a leading-edge slat and a single-segment trailing-edge flap. Data include surface static-pressure distributions (integrated to obtain lift), drag data obtained with wake-rake surveys, and fbwfield surveys obtained with a flat-tube and five-hole probe at nine stations on the configuration's upper surface. Effects of variations in Reynolds number and flap gap on airfoil performance and flowfield survey data are presented.

Wing flutter boundary prediction using unsteady Euler aerodynamic method

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

Unsteady Subsonic Aerodynamics for Bodies and Wings with External Stores Including Wake Effect

Abstract: An unsteady subsonic method has been developed for aerodynamic computations of any elastic or rigid aricraft with external stores. The method consists of two integral parts: a body surface panel method (SPM) and a constant-pressure lifting surface method, which is the subsonic parallel of the HGM (or the ZONA51 code) for unsteady supersonics. The body considered can be flat-based or close-ended and its geometry input is amenable to any given fuselage or store configuration. The present method is considered an advancement over the past development at least in three aspects: (1) correct unsteady boundary condition on body, (2) a new wake model to account for the body/wake effect and (3) improved accuracy in wingbody interference. Various AGARD iifting surfaces, truncated blunt and pointed bodies and a number of NLR wing-storetiptank combinations were studied for method validation. The present method has shown substantial improvement in the pressures, stability derivatives and airloads on these configurations. For all cases considered, the present results, with or without the wake model, have consistently shown closest agreement with all measured data among existing methods. Therefore, we believe that an accurate and effective method is finally at hand for subsonic aeroelastic applications.

Effect of micron-sized roughness on transition in swept-wing flows

Abstract: Boundary-layer transition-to-turbulence studies are conducted in the Arizona State University Unsteady Wind Tunnel on a 45-degree swept airfoil. The pressure gradient is designed so that the initial stability characteristics are purely crossflow-dominated. Flow visualization and hot-wire measurements show that the development of the crossflow vortices is influenced by roughness near the attachment-line. Comparisons of transition location are made between a painted surface, a machine-polished surface, and a hand-polished surface. Then, isolated 6 micron roughness elements are placed near the attachment line on the airfoil surface under conditions of the final polish (0.25 micron rms). These elements amplify a centered stationary crossflow vortex and its neighbors, resulting in localized early transition. The diameter, height, and location of these roughness elements are varied in a systematic manner. Spanwise hot-wire measurements are taken behind the roughness element to document the enhanced vortices. These scans are made at several different chord locations to examine vortex growth.

Vortex generators used to control laminar separation bubbles

Abstract: This study examines the effect of various forms of vortex generators on the laminar separation bubble of a two-dimensional low Reynolds number Liebeck LA2573A airfoil. The objective of this research was to determine the effects that different generator sizes and spacings have upon the separation bubble and the drag. Windtunnel measurements were made on several generator configurations at Reynolds numbers ranging from 200,000 to 600,000 at angles of attack less than the stall angle where the separation bubble can provide a significant contribution to the airfoil drag. The vortex generators used were constructed small enough to be contained completely within the laminar boundary layer. Wind-tunnel data included airfoil drag and mean and fluctuating velocity measurements in the laminar and turbulent boundary layers. Results have shown that the use of vortex generators provides a measurable decrease in airfoil drag at the lower range of Reynolds numbers tested. At the airfoil's design condition and Reynolds number of 235,000, the submerged vortex generators were shown to decrease the airfoil drag by a maximum of 38% at C/ = 0.572.

Foreign object damage resistant composite blade and manufacture

Abstract: The present invention provides a composite airfoil particularly useful as a fan blade, having a high degree of twist, in a large high bypass ratio turbofan engine. The composite airfoil of the present invention has a progression of filament reinforced airfoil laminations of varying span, arranged in order by span, and interrupted by at least one filament reinforced airfoil lamination having a span out of height order to form what is called a confused or broken shear plane where radially outer tips of the laminations end.

Wind turbine rotor blade

Abstract: Horizontal axis, free yaw, self-regulated wind turbines which have strong, lightweight, fatigue resistant, fixed pitch, wood/GRE blades and exhibit superior performance a range of wind speeds. The blades are designed by employing defined inboard, midspan, and outboard airfoil profiles and interpolating the profiles between the defined profiles and from the latter to the root and the tip of the blades.

Transonic shock oscillations calculated with a new interactive boundary layer coupling method

Abstract: A new viscous-inviscid interactive coupling method is described with the aim of allowing time-accurate computation of unsteady transonic flows involving separation and reattachment. A lag-entrainment integral boundary layer method is used in conjunction with a transonic small disturbance potential code. The solutions are coupled with a novel variable gain, integral control method for the boundary layer displacement thickness. Efficient and robust computations of steady and unsteady separated flows, including steady separation bubbles and self-excited shock-induced oscillations, are presented. The buffet onset boundary for the NACA 0012 airfoil is accurately predicted and shown computationally to be a Hopf bifurcation. Shock-induced oscillations are also presented for the 18 percent thick circular arc airfoil. The oscillation onset boundaries and frequencies are accurately predicted, as is the experimentally observed hysteresis of the oscillations with Mach number; this latter stability boundary is identified as a jump phenomenon.

Progress in high-lift aerodynamic calculations

Abstract: The current work presents progress in the effort to numerically simulate the flow over high-lift aerodynamic components, namely, multi-element airfoils and wings in either a take-off or a landing configuration. The computational approach utilizes an incompressible flow solver and an overlaid chimera grid approach. A detailed grid resolution study is presented for flow over a three-element airfoil. Two turbulence models, a one-equation Baldwin-Barth model and a two equation k-omega model are compared. Excellent agreement with experiment is obtained for the lift coefficient at all angles of attack, including the prediction of maximum lift when using the two-equation model. Results for two other flap riggings are shown. Three-dimensional results are presented for a wing with a square wing-tip as a validation case. Grid generation and topology is discussed for computing the flow over a T-39 Sabreliner wing with flap deployed and the initial calculations for this geometry are presented.

Impingement cooled airfoil with bonding foil insert

Abstract: An impinqement cooled airfoil is fabricated by diffusion bonding a pair of airfoil half-sections together using an insert which is prefabricated from diffusion bonding foil. The insert is perforated so as to act as an impingement baffle. Axially-extending ribs may be formed on the internal walls of the airfoil half-sections or on the insert to support and accurately space the insert member from the internal walls so as to optimize impingement cooling.

Coolable airfoil structure

Abstract: A coolable airfoil having internal cooling passages (82, 84, 74) is disclosed. Various construction details are developed to increase cooling in critical locations of the airfoil. In one embodiment, a pair of side-by-side spanwisely extending passages (82, 84) flow charting air to the tip (flag) passage (74).

Airfoil Shape Optimization Using Sensitivity Analysis on Viscous Flow Equations

Abstract: An aerodynamic shape optimization method has previously been developed by the authors using the Euler equations and has been applied to supersonic-hypersonic nozzle designs. This method has also included a flowfield extrapolation (or flow prediction) method based on the Taylor series expansion of an existing CFD solution. The present paper reports on the extension of this method to the thin-layer Navier-Stokes equations in order to account for the viscous effects. Also, to test the method under highly nonlinear conditions, it has been applied to the transonic flows. Initially, the success of the flow prediction method is tested. Then, the overall method is demonstrated by optimizing the shapes of two supercritical transonic airfoils at zero angle of attack. The first one is shape optimized to achieve a minimum drag while obtaining a lift above a specified value. Whereas, the second one is shape optimized for a maximum lift while attaining a drag below a specified value. The results of these two cases indicate that the present method can produce successfully optimized aerodynamic shapes.

Acoustic radiation from a thin airfoil in nonuniform subsonic flows

Abstract: Noise resulting from the interaction of a three dimensional gust is modeled directly from unsteady aerodynamic codes. This paper provides bench mark results for the case of a thin airfoil for comparison with the more complex case of a loaded airfoil. It is shown that the acoustic pressure pattern strongly depends on the value of a certain reduced frequency. The effects of an oblique gust are examined and shown to significantly reduce the acoustic power for certain gust parameters range. The Acoustic power versus the gust frequency exhibits a maximum. High frequency and non-compact source effects are also investigated.

Aerodynamic shape optimization using preconditioned conjugate gradient methods

Abstract: In an effort to further improve upon the latest advancements made in aerodynamic shape optimization procedures, a systematic study is performed to examine several current solution methodologies as applied to various aspects of the optimization procedure. It is demonstrated that preconditioned conjugate gradient-like methodologies dramatically decrease the computational efforts required for such procedures. The design problem investigated is the shape optimization of the upper and lower surfaces of an initially symmetric (NACA-012) airfoil in inviscid transonic flow and at zero degree angle-of-attack. The complete surface shape is represented using a Bezier-Bernstein polynomial. The present optimization method then automatically obtains supercritical airfoil shapes over a variety of freestream Mach numbers. Furthermore, the best optimization strategy examined resulted in a factor of 8 decrease in computational time as well as a factor of 4 decrease in memory over the most efficient strategies in current use.

Feasibility study to build a smart rotor: trailing edge flap actuation

Abstract: This paper presents an experimental study on the development of a Froude scale helicopter rotor blade with trailing edge flap as a vibration reduction device actuated by piezoelectric crystals. A fixed wing model with NACA 0012 airfoil, 3.0 inch blade chord and 20% trailing edge flap is fabricated and tested in the open-jet tunnel to determine the dynamic flap response at various blade angles of attack and excitation frequencies. The results show that for a given velocity, flap response does not change appreciably with the excitation frequency and the blade angle of attack.© (1993) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Interferometric investigations of compressible dynamic stall over a transiently pitching airfoil

Abstract: The compressible dynamic stall flowfield over a NACA 0012 airfoil transiently pitching from 0 to 60 deg at a constant rate under compressible flow conditions has been studied using real-time interferometry. A quantitative description of the overall flowfield, including the finer details of dynamic stall vortex formation, growth, and the concomitant changes in the airfoil pressure distribution, has been provided by analyzing the interferograms. For Mach numbers above 0.4, small multiple shocks appear near the leading edge and are present through the initial stages of dynamic stall. Dynamic stall was found to occur coincidentally with the bursting of the separation bubble over the airfoil. Compressibility was found to confine the dynamic stall vortical structure closer to the airfoil surface. The measurements show that the peak suction pressure coefficient drops with increasing freestream Mach number, and also it lags the steady flow values at any given angle of attack. As the dynamic stall vortex is shed, an anti-clockwise vortex is induced near the trailing edge, which actively interacts with the post-stall flow.

Efficient Simulation of Incompressible Viscous Flow over Single and Multielement Airfoils

Abstract: The incompressible, viscous, turbulent flow over single and multielement airfoils is numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The solution algorithm uses the method of pseudocompressibility with an upwind-differencing scheme for the convective fluxes, and an implicit line-relaxation scheme. The motivation for this work includes interest in studying high-lift takeoff and landing configurations of various aircraft. In particular, accurate computation of lift and drag at various angles of attack up to stall is desired. Two different turbulence models are tested in computing the flow over a NACA 4412 airfoil; an accurate prediction of stall is obtained. The approach used for multielement airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: 1) a patched system of grids, and 2) an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Generally, good agreement with experimental surface pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of 1 min of CPU time on a CRAY YMP per element in the airfoil configuration.

Light weight steam turbine blade

Abstract: A steam turbine blade is provided in which the geometry of the blade airfoil is configured to reduce the weight of the blade while maintaining adequate thermodynamic performance. The blade has a high camber angle to increase its stiffness and an airfoil portion having pressure and suction surfaces. The curvature of the portion of the pressure surface between the base and 75% blade height has a point of inflection in an area adjacent the leading edge so that the curvature switches sign. The curvature upstream of the point of inflection is concave, while the curvature downstream of the point of inflection is convex.

Pressure measurements on a pitching airfoil in a water channel

Abstract: Measurements of unsteady pressures over a symmetric NACA 0015 airfoil performing pitching maneuvers are reported. The tests were performed in an open-surface water channel specially constructed for this purpose. The design of the apparatus allowed the pressure measurements to be made to a very high degree of spatial and temperal resolution. Reynolds numbers in the range of 5.2×10 4 to 2.2×10 5 were studied. Although the results qualitatively agreed with earlier studies performed at similar Reynolds numbers, the magnitudes of pressure and aerodynamic forces measured were observed to be much larger than those measured in earlier pitchup studies

Deforming Grid Variational Principle for Unsteady Small Disturbance Flows in Cascades

Abstract: A variational method for computing unsteady subsonic flows in turbomachinery blade rows is presented. A variational principle that describes the harmonic small disturbance behavior of the full potential equations about a nonlinear mean flow is developed. Included in this variational principle is the effect of a deforming computational grid that conforms to the motion of vibrating airfoils. Bilinear isoparametric finite elements are used to discretize the variational principle, and the resulting discretized equations are solved efficiently using lower-upper decomposition. The use of a deforming computational grid dramatically improves the accuracy of the method since no error-producing extrapolation is required to apply the upwash boundary conditions or to evaluate the unsteady pressure on the airfoil surfaces

Micro-grooved heat transfer wall

Abstract: A gas turbine engine hot section component such as a turbine blade or vane having an airfoil is provided a non-film cooled portion of a heat transfer wall having a hot surface and a plurality of longitudinally extending micro-grooves disposed in the portion of the wall along the hot surface in a direction parallel to the direction of the hot gas flow. The depth of the micro-grooves is very small and on the order of magnitude of a predetermined laminar sublayer of a turbulent boundary layer. The grooves are sized so as to alter the boundary layer thickness near the leading edge of the airfoil to reduce heat transfer from the hot gas flow to the airfoil near the leading edge. In one embodiment the micro-grooves are about 0.001 inches deep and have a preferred depth range of from about 0.001 inches to 0.005 inches and which are square, rectangular, or triangular in cross-section and the micro-grooves are spaced about one width apart.

Antilift wiper blade assembly with integral flexor and airfoil

Abstract: An antilift wiper blade assembly for clearing a generally nonplanar vehicular windshield includes a pressure-distributing superstructure and a resilient wiping element. An elongated flexor interconnects the wiping element and the superstructure and has a longitudinally extending channel engaging the wiping element. A pair of laterally extending flanges are spaced above the bottom wall of the flexor and are engaged by the yoke elements, with the wiping element projecting through an elongated opening formed in the bottom wall of the flexor and into contact with the windshield. An upwardly curved airfoil extends tangentially from the bottom face of the flexor to generate an aerodynamic force which urges the wiping element into engagement with the windshield. A number of parallel slits are spaced along the length of the airfoil to partition the airfoil into a plurality of independently deformable aerodynamic force generating fins.

Aerodynamic characteristics of the HL-20

Abstract: Wind-tunnel tests were made from subsonic to hypersonic speeds to define the aerodynamic characteristics of the HL-20 lifting-body configuration. The data have been assembled into an aerodynamic database for flight analysis of this proposed vehicle. The wind-tunnel data indicated that the model was longitudinally and laterally stable (about a center-of-gravity location of 0.54 body length) over the test range from Mach 20 to 0.3. At hypersonic speeds, the HL-20 model trimmed at a lift/drag (L/D) ratio of 1.4. This value gives the vehicle a crossrange capability similar to that of the Space Shuttle. At subsonic speeds, the HL-20 has a trimmed LID ratio of about 3.6. Replacing the flat-plate outboard fins with fins haying an airfoil shape increased the maximum subsonic trimmed LID to 4.2.

Aerodynamic Performance Effects due to Small Leading-Edge Ice (Roughness) on Wings and Tails

Abstract: Systematic experimental studies have been carried out to establish the aerodynamic impact of very small leading-edge simulated ice (roughness) formations on lifting surfaces. The geometries studied include singleelement configurations (airfoil and three-dimensional tail) as well as multielement high-lift airfoil geometries. Emphasis in these studies was placed on obtaining results at high Reynolds numbers to insure the applicability of the findings to full-scale situations. It has been found that the well-known Brumby correlation for the adverse lift impact of discrete roughness elements at the leading edge is not appropriate for cases representative of initial frost formation (i.e., distributed roughness). It has further been found that allowing initial ice formations, of a size required for removal by presently proposed de-icing systems, could lead to maximum lift losses of approximately 40% for single-element airfoils. Losses in angle-of-attack margin-to-stall are equally substantial— as high as 6 deg. Percentage losses for multielement airfoils are not as severe as for single-element configurations, but degradations of the angle of attack-to-sta ll margin are the same for both.

Static aeroelastic control of an adaptive lifting surface

Abstract: Piezoelectric actuators are embedded in an idealized laminated composite wing structure that can adapt to changing flight conditions. A feedback control system senses wing roots strains and then applies a proportional voltage to active actuator layers to change wing lift or divergence dynamic pressure

Direct Calculation of Sound Radiated From Bodies in Nonuniform Flows

Abstract: Sound radiated from a single airfoil and a cascade of airfoils in three-dimensional gusts is directly calculated. Euler’s equations are linearized about the mean flow of the airfoil or cascade. The velocity field is split into a vortical part and a potential part. The latter is governed by a single nonconstant-coefficient convective wave equation. For a single airfoil, the radiated sound is calculated using Kirchhoff’s method from the mid field of the unsteady pressure obtained through the unsteady aerodynamic solver. The results indicate the importance of the contribution of the quadrupole effects to the sound field. For a cascade of airfoils, the acoustic pressure is directly obtained by solving the partial differential equation. The results show that, as the maximum Mach number on the blade surface nears unity, there is a significant rise in the local unsteady pressure, and also a significant increase in the upstream acoustic pressure.

Prescribed-Curvature-Distribution Airfoils for the Preliminary Geometric Design of Axial-Turbomachinery Cascades

Abstract: Blade surfaces with continuous curvature and continuous slope of curvature minimize the possibility of flow separation, lead to improved blade designs, and reduce the direct and inverse blade-design iterations for the seection of isolated airfoils and gas-turbine-blade cascades. A method for generating two-dimensional blade shapes is presented. The geometry near the trailing edge is specified by an analytic polynomial, the main portion of the blade surface is mapped using as input a prescribed surface-curvature distribution, and the leading edge is specified as a thickness distribution added to a construction line. This procedure is similar for the suction and pressure surfaces, and by specification it constructs continuous slope-of-curvature surfaces that result in smooth surface-Mach-number and surface-pressure distributions