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Sample records for airfoil section characteristics

  1. The development of cambered airfoil sections having favorable lift characteristics at supercritical Mach numbers

    NASA Technical Reports Server (NTRS)

    Graham, Donald J

    1949-01-01

    Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined from two-dimensional wind-tunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NACA 6-series airfoils.

  2. Airfoil Section Characteristics as Affected by Variations of the Reynolds Number

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Sherman, Albert

    1937-01-01

    Report presents the results of an investigation of a systematically chosen representative group of related airfoils conducted in the NACA variable-density wind tunnel over a wide range of Reynolds number extending well into the flight range. The tests were made to provide information from which the variations of airfoil section characteristics with changes in the Reynolds number could be inferred and methods of allowing for these variations in practice could be determined. This work is one phase of an extensive and general airfoil investigation being conducted in the variable-density tunnel and extends the previously published researches concerning airfoil characteristics as affected by variations in airfoil profile determined at a single value of the Reynolds number.

  3. The characteristics of 78 related airfoil sections from tests in the variable-density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Ward, Kenneth E; Pinkerton, Robert M

    1933-01-01

    An investigation of a large group of related airfoils was made in the NACA variable-density wind tunnel at a large value of the Reynolds number. The tests were made to provide data that may be directly employed for a rational choice of the most suitable airfoil section for a given application. The variation of the aerodynamic characteristics with variations in thickness and mean-line form were systematically studied. (author)

  4. Low speed aerodynamic characteristics of a 17 percent thick airfoil section designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1973-01-01

    Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1). The results were compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 million to 20.0 million. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 million to 6.0 million.

  5. Boundary-layer and stalling characteristics of two symmetrical NACA low-drag airfoil sections

    NASA Technical Reports Server (NTRS)

    Mccullough, George B; Gault, Donald E

    1947-01-01

    Two symmetrical airfoils, an NACA 633-018 and an NACA 631-012, were investigated for the purpose of determining their stalling and boundary-layer characteristics with a view toward the eventual application of this information to the problem of boundary-layer control. Force measurements, pressure distributions, tuft studies, and boundary-layer-profile measurements were made at a value of 5,800,000 Reynolds number. It was found that the 18-percent-thick airfoil stalled progressively from the trailing edge because of separation of the turbulent boundary layer. In contrast, the12-percent-thick airfoil stalled abruptly from a separation of flow near the leading edge before the turbulent boundary layer became subject to separation. From this it was concluded that if high values of lift are to be obtained with thin, high-critical-speed sections by means of boundary-layer control, the work must be directed toward delaying the separation of flow near the leading edge. It was found that the presence of a nose flap on the 12-percent-thick section caused the airfoil to stall in a manner similar to that of the 18-percent-thick section.

  6. Aerodynamic characteristics and pressure distributions for an executive-jet baseline airfoil section

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Mineck, Raymond E.

    1993-01-01

    A wind tunnel test of an executive-jet baseline airfoil model was conducted in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. The primary goal of the test was to measure airfoil aerodynamic characteristics over a wide range of flow conditions that encompass two design points. The two design Mach numbers were 0.654 and 0.735 with corresponding Reynolds numbers of 4.5 x 10(exp 6) and 8.9 x 10(exp 6) based on chord, respectively, and normal-force coefficients of 0.98 and 0.51, respectively. The tests were conducted over a Mach number range from 0.250 to 0.780 and a chord Reynolds number range from 3 x 10(exp 6) to 18 x 10(exp 6). The angle of attack was varied from -2 deg to a maximum below 10 deg with one exception in which the maximum was 14 deg for a Mach number of 0.250 at a chord Reynolds number of 4.5 x 10(exp 6). Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The adaptive-wall test section had flexible top and bottom walls and rigid sidewalls. Wall interference was minimized by the movement of the adaptive walls, and the airfoil aerodynamic characteristics were corrected for any residual top and bottom wall interference.

  7. Low-speed aerodynamic characteristics of a 13-percent-thick airfoil section designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.; Somers, D. M.

    1975-01-01

    Wind-tunnel tests were conducted to determine the low-speed section characteristics of a 13 percent-thick airfoil designed for general aviation applications. The results were compared with NACA 12 percent-thick sections and with the 17 percent-thick NASA airfoil. The tests were conducted ovar a Mach number range from 0.10 to 0.35. Chord Reynolds numbers varied from about 2,000,000 to 9,000,000.

  8. Aerodynamic Characteristics of NACA 23012 and 23021 Airfoils with 20-Percent-chord External-Airfoil Flaps of NACA 23012 Section

    NASA Technical Reports Server (NTRS)

    Platt, Robert C; Abbott, Ira H

    1937-01-01

    Report presents the results of an investigation of the general aerodynamic characteristics of the NACA 23012 and 23021 airfoils, each equipped with a 0.20c external flap of NACA 23012 section. The tests were made in the NACA 7 by 10-foot and variable-density wind tunnels and covered a range of Reynolds numbers that included values corresponding to those for landing conditions of a wide range of airplanes. Besides a determination of the variation of lift and drag characteristics with position of the flap relative to the main airfoil, complete aerodynamic characteristics of the airfoil-flap combination with a flap hinge axis selected to give small hinge moments were measured in the two tunnels. Some measurements of air loads on the flap itself in the presence of the wing were made in the 7 by 10-foot wind tunnel.

  9. Transonic Aerodynamic Characteristics of Two Wedge Airfoil Sections Including Unsteady Flow Studies

    NASA Technical Reports Server (NTRS)

    Johnston, Patrick J.

    1959-01-01

    A two-dimensional wind-tunnel investigation has been conducted on a 20-percent-thick single-wedge airfoil section. Steady-state forces and moments were determined from pressure measurements at Mach numbers from 0.70 to about 1.25. Additional information on the flows about the single wedge is provided by means of instantaneous pressure measurements at Mach numbers up to unity. Pressure distributions were also obtained on a symmetrical double-wedge or diamond-shaped profile which had the same leading-edge included angle as the single-wedge airfoil. A comparison of the data on the two profiles to provide information on the effects of the afterbody showed that with the exception of drag, the single-wedge profile proved to be aerodynamically superior to the diamond profile in all respects. The lift effectiveness of the single-wedge airfoil section far exceeded that of conventional thin airfoil sections over the speed range of the investigation. Pitching-moment irregularities, caused by negative loadings near the trailing edge, generally associated with conventional airfoils of equivalent thicknesses were not exhibited by the single-wedge profile. Moderately high pulsating pressures existing over the base of the single-wedge airfoil section were significantly reduced as the Mach number was increased beyond 0.92 and the boundaries of the dead airspace at the base of the model converged to eliminate the vortex street in the wake. Increasing the leading-edge radius from 0 to 1 percent of the chord had a minor effect on the steady-state forces and generally raised the level of pressure pulsations over the forward part of the single-wedge profile.

  10. Effect of Flap Deflection on Section Characteristics of S813 Airfoil; Period of Performance: 1993--1994

    SciTech Connect

    Somers, D. M.

    2005-01-01

    The effect of small deflections of a 30% chord, simple flap on the section characteristics of a tip airfoil, the S813, designed for 20- to 30-meter, stall-regulated, horizontal-axis wind turbines has been evaluated theoretically. The decrease in maximum lift coefficient due to leading-edge roughness increases in magnitude with increasing, positive flap deflection and with decreasing Reynolds number.

  11. The Aerodynamic Characteristics of Six Full-Scale Propellers Having Different Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Biermann, David; Hartman, Edwin P

    1939-01-01

    Wind-tunnel tests are reported of six 3-blade 10-foot propellers operated in front of a liquid-cooled engine nacelle. The propellers were identical except for blade airfoil sections, which were: Clark y, R.A.F. 6, NACA 4400, NACA 2400-34, NACA 2rsub200, and NACA 6400. The range of blade angles investigated extended for 15 degrees to 40 degrees for all propellers except the Clark y, for which it extended to 45 degrees. The results showed that the range in maximum efficiency between the highest and lowest values was about 3 percent. The highest efficiencies were for the low-camber sections.

  12. Critical Mach Numbers of Thin Airfoil Sections with Plain Flaps

    NASA Technical Reports Server (NTRS)

    Pardee, Otway O'm.; Heaslet, Max A.

    1946-01-01

    Critical Mach number as function of lift coefficient is determined for certain moderately thick NACA low-drag airfoils. Results, given graphically, included calculations on same airfoil sections with plain flaps for small flap deflections. Curves indicate optimum critical conditions for airfoils with flaps in such form that they can be compared with corresponding results for zero flap deflections. Plain flaps increase life-coefficient range for which critical Mach number is in region of high values characteristic of low-drag airfoils.

  13. Effects of forward contour modification on the aerodynamic characteristics of the NACA 641-212 airfoil section

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Mendoza, J. P.; Bandettini, A.

    1975-01-01

    Two different forward contour modifications designed to increase the maximum lift coefficient of the NACA 64 sub 1-212 airfoil section were evaluated experimentally at low speeds. One modification consisted of a slight droop of the leading edge with an increased leading-edge radius; the other modification incorporated increased thickness over the forward 35 percent of the upper surface of the profile. Both modified airfoil sections were found to provide substantially higher maximum lift coefficients than the 64 sub 1-212 section. The drooped leading-edge modification incurred a drag penalty of approximately 10 percent at low and moderate lift coefficients and exhibited a greater nosedown pitching moment than the 64 sub 1-212 profile. The upper surface modification produced about the same drag level as the 64 sub 1-212 section at low and moderate lift coefficients and less nosedown pitching moment than the 64 sub 1-212 profile. Both modified airfoil sections had lower drag coefficients than the 64 sub 1-212 section at high lift coefficients.

  14. Maximum Mean Lift Coefficient Characteristics at Low Tip Mach Numbers of a Hovering Helicopter Rotor Having an NACA 64(1)A012 Airfoil Section

    NASA Technical Reports Server (NTRS)

    Powell, Robert D., Jr.

    1959-01-01

    An investigation has been conducted on the Langley helicopter test tower to determine experimentally the maximum mean lift-coefficient characteristics at low tip Mach number and a limited amount of drag- divergence data at high tip Mach number of a helicopter rotor having an NACA 64(1)AO12 airfoil section and 8 deg of linear washout. Data are presented for blade tip Mach numbers M(t) of 0.29 to 0.74 with corresponding values 6 6 of tip Reynolds number of 2.59 x 10(exp 6) and 6.58 x 10(exp 6). Comparisons are made between the data from the present rotor with results previously obtained from two other rotors: one having NACA 0012 airfoil sections and the other having an NACA 0009 airfoil tip section. At low tip Mach numbers, the maximum mean lift coefficient for the blade having the NACA 64(1)AO12 section was about 0.08 less than that obtained with the blade having the NACA 0009 tip section and 0.21 less than the value obtained with the blade having the NACA 0012 tip section. Blade maximum mean lift coefficient values were not obtained for Mach number values greater than 0.47 because of a blade failure encountered during the tests. The effective mean lift-curve slope required for predicting rotor thrust varied from 5.8 for the tip Mach nuniber range of 0.29 to 0.55 to a value of 6.65 for a tip Mach number of 0.71. The blade pitching-moment coefficients were small and relatively unaffected by changes in thrust coefficient and Mach number. In the instances in which stall was reached, the break in the blade pitching-moment curve was in a stable direction. The efficiency of the rotor decreased with an increase in tip speed. Expressed as figure of merit, at a tip Mach number of 0.29 the maximum value was about 0.74. Similar measurements made on another rotor having an NACA 0012 airfoil and with a rotor having an NACA 0009 tip section, showed a value of 0.75. Synthesized section lift and profile-drag characteristics for the rotor-blade airfoil section are presented as an

  15. The effects of variations in Reynolds number between 3.0 x 10sub6 and 25.0 x 10sub6 upon the aerodynamic characteristics of a number of NACA 6-series airfoil sections

    NASA Technical Reports Server (NTRS)

    Loftin, Laurence K, Jr; Bursnall, William J

    1950-01-01

    Results are presented of an investigation made to determine the two-dimensional lift and drag characteristics of nine NACA 6-series airfoil section at Reynolds numbers of 15.0 x 10sub6, 20.0 x 10sub6, and 25.0 x 10sub6. Also presented are data from NACA Technical Report 824 for the same airfoils at Reynolds numbers of 3.0 x 10sub6, 6.0 x 10sub6, and 9.0 x 10sub6. The airfoils selected represent sections having variations in the airfoil thickness, thickness form, and camber. The characteristics of an airfoil with a split flap were determined in one instance, as was the effect of surface roughness. Qualitative explanations in terms of flow behavior are advanced for the observed types of scale effect.

  16. Effects of independent variation of Mach and Reynolds numbers on the low-speed aerodynamic characteristics of the NACA 0012 airfoil section

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.

    1988-01-01

    A comprehensive data base is given for the low speed aerodynamic characteristics of the NACA 0012 airfoil section. The Langley low-turbulence pressure tunnel is the facility used to obtain the data. Included in the report are the effects of Mach number and Reynolds number and transition fixing on the aerodynamic characteristics. Presented are also comparisons of some of the results with previously published data and with theoretical estimates. The Mach number varied from 0.05 to 0.36. The Reynolds number, based on model chord, varied from 3 x 10 to the 6th to 12 x 10 to the 6th power.

  17. Turbine airfoil having outboard and inboard sections

    SciTech Connect

    Mazzola, Stefan; Marra, John J

    2015-03-17

    A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.

  18. The aerodynamic characteristics of eight very thick airfoils from tests in the variable density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1932-01-01

    Report presents the results of wind tunnel tests on a group of eight very thick airfoils having sections of the same thickness as those used near the roots of tapered airfoils. The tests were made to study certain discontinuities in the characteristic curves that have been obtained from previous tests of these airfoils, and to compare the characteristics of the different sections at values of the Reynolds number comparable with those attained in flight. The discontinuities were found to disappear as the Reynolds number was increased. The results obtained from the large-scale airfoil, a symmetrical airfoil having a thickness ratio of 21 per cent, has the best general characteristics.

  19. An experimental study of dynamic stall on advanced airfoil section. Volume 2: Pressure and force data

    NASA Technical Reports Server (NTRS)

    Mcalister, K. W.; Pucci, S. L.; Mccroskey, W. J.; Carr, L. W.

    1982-01-01

    Experimentally derived force and moment data are presented for eight airfoil sections that were tested at fixed and varying incidence in a subsonic two dimensional stream. Airfoil incidence was varied through sinusoidal oscillations in pitch over a wide range of amplitude and frequency. The surface pressure distribution, as well as the lift, drag, and pitching moment derived therefrom, are displayed in a uniform fashion to delineate the static and dynamic characteristics of each airfoil both in and out of stall.

  20. The Aerodynamic Characteristics of Full-Scale Propellers Having 2, 3, and 4 Blades of Clark Y and R.A.F. 6 Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Hartman, Edwin P; Biermann, David

    1938-01-01

    Aerodynamic tests were made of seven full-scale 10-foot-diameter propellers of recent design comprising three groups. The first group was composed of three propellers having Clark y airfoil sections and the second group was composed of three propellers having R.A.F. 6 airfoil sections, the propellers of each group having 2, 3, and 4 blades. The third group was composed of two propellers, the 2-blade propeller taken from the second group and another propeller having the same airfoil section and number of blades but with the width and thickness 50 percent greater. The tests of these propellers reveal the effect of changes in solidity resulting either from increasing the number of blades or from increasing the blade width propeller design charts and methods of computing propeller thrust are included.

  1. Aerodynamic characteristics of seven symmetrical airfoil sections through 180-degree angle of attack for use in aerodynamic analysis of vertical axis wind turbines

    SciTech Connect

    Sheldahl, R E; Klimas, P C

    1981-03-01

    When work began on the Darrieus vertical axis wind turbine (VAWT) program at Sandia National Laboratories, it was recognized that there was a paucity of symmetrical airfoil data needed to describe the aerodynamics of turbine blades. Curved-bladed Darrieus turbines operate at local Reynolds numbers (Re) and angles of attack (..cap alpha..) seldom encountered in aeronautical applications. This report describes (1) a wind tunnel test series conducted at moderate values of Re in which 0 less than or equal to ..cap alpha.. less than or equal to 180/sup 0/ force and moment data were obtained for four symmetrical blade-candidate airfoil sections (NACA-0009, -0012, -0012H, and -0015), and (2) how an airfoil property synthesizer code can be used to extend the measured properties to arbitrary values of Re (10/sup 4/ less than or equal to Re less than or equal to 10/sup 7/) and to certain other section profiles (NACA-0018, -0021, -0025).

  2. A study of test section configuration for shock tube testing of transonic airfoils

    NASA Technical Reports Server (NTRS)

    Cook, W. J.

    1978-01-01

    Two methods are investigated for alleviating wall interference effects in a shock tube test section intended for testing two-dimensional transonic airfoils. The first method involves contouring the test section walls to match approximate streamlines in the flow. Contours are matched to each airfoil tested to produce results close to those obtained in a conventional wind tunnel. Data from a previous study and the present study for two different airfoils demonstrate that useful results are obtained in a shock tube using a test section with contoured walls. The second method involves use of a fixed-geometry slotted-wall test section to provide automatic flow compensation for various airfoils. The slotted-wall test section developed exhibited the desired performance characteristics in the approximate Mach number range 0.82 to 0.89, as evidenced by good agreement obtained between shock tube and wind tunnel results for several airfoil flows.

  3. Verification of U-21 Cloud Parameter Measurement Equipment and Comparison of Natural and Artifical Ice Accretion Characteristics on Rotor Blade Airfoil Sections

    DTIC Science & Technology

    1987-05-01

    for airworthiness evaluation of the JU-21A modifications. TEST OBJECTIVES 3. The objectives of this program were: a. Obtain comparative measurements... aircraft showed reasonable agreement between measured natural cloud parameters. Airworthiness test flights of the Airfoil Section Array (ASA) icing...modification of the aircraft confl- uration. Airworthiness was established by a combination of engineering analysis and flight tests. Flight loads

  4. Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils

    DTIC Science & Technology

    2003-12-01

    Development, and Engineering Command Ames Research Center Moffett Field, California December 2003 National Aeronautics and Space Administration Ames...60A ROTOR BLADE AND AIRFOILS ................................................................................... 2 EVALUATION OF SECTION CHARACTERISTICS...Characteristics of SC1095 and SC1094 R8 Airfoils WILLIAM G. BOUSMAN Aeroflightdynamics Directorate U.S. Army Research, Development, and Engineering Command Ames

  5. Theoretical and experimental study of a new method for prediction of profile drag of airfoil sections

    NASA Technical Reports Server (NTRS)

    Goradia, S. H.; Lilley, D. E.

    1975-01-01

    Theoretical and experimental studies are described which were conducted for the purpose of developing a new generalized method for the prediction of profile drag of single component airfoil sections with sharp trailing edges. This method aims at solution for the flow in the wake from the airfoil trailing edge to the large distance in the downstream direction; the profile drag of the given airfoil section can then easily be obtained from the momentum balance once the shape of velocity profile at a large distance from the airfoil trailing edge has been computed. Computer program subroutines have been developed for the computation of the profile drag and flow in the airfoil wake on CDC6600 computer. The required inputs to the computer program consist of free stream conditions and the characteristics of the boundary layers at the airfoil trailing edge or at the point of incipient separation in the neighborhood of airfoil trailing edge. The method described is quite generalized and hence can be extended to the solution of the profile drag for multi-component airfoil sections.

  6. Airfoil

    SciTech Connect

    Ristau, Neil; Siden, Gunnar Leif

    2015-07-21

    An airfoil includes a leading edge, a trailing edge downstream from the leading edge, a pressure surface between the leading and trailing edges, and a suction surface between the leading and trailing edges and opposite the pressure surface. A first convex section on the suction surface decreases in curvature downstream from the leading edge, and a throat on the suction surface is downstream from the first convex section. A second convex section is on the suction surface downstream from the throat, and a first convex segment of the second convex section increases in curvature.

  7. Characteristics of an Airfoil as Affected by Fabric Sag

    NASA Technical Reports Server (NTRS)

    Ward, Kenneth E

    1932-01-01

    This report presents the results of tests made at a high value of the Reynolds Number in the N.A.C.A. variable-density wind tunnel to determine the aerodynamic characteristics of an airfoil as affected by fabric sag. Tests were made of two Gottingen 387 airfoils, one having the usual smooth surface and the other having a surface modified to simulate two types of fabric sag. The results of these tests indicate that the usual sagging of the wind covering between ribs has a very small effect on the aerodynamic characteristics of an airfoil.

  8. Aerodynamic characteristics of an improved 10-percent-thick NASA supercritical airfoil. [Langley 8 foot transonic tunnel tests

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1974-01-01

    Refinements in a 10 percent thick supercritical airfoil produced improvements in the overall drag characteristics at normal force coefficients from about 0.30 to 0.65 compared with earlier supercritical airfoils which were developed for a normal force coefficient of 0.7. The drag divergence Mach number of the improved supercritical airfoil (airfoil 26a) varied from approximately 0.82 at a normal force coefficient to of 0.30, to 0.78 at a normal force coefficient of 0.80 with no drag creep evident. Integrated section force and moment data, surface pressure distributions, and typical wake survey profiles are presented.

  9. Closed-form equations for the lift, drag, and pitching-moment coefficients of airfoil sections in subsonic flow

    NASA Technical Reports Server (NTRS)

    Smith, R. L.

    1978-01-01

    Closed-form equations for the lift, drag, and pitching moment coefficients of two dimensional airfoil sections in steady subsonic flow were obtained from published theoretical and experimental results. A turbulent boundary layer was assumed to exist on the airfoil surfaces. The effects of section angle of attack, Mach number, Reynolds number, and the specific airfoil type were considered. The equations were applicable through an angle of attack range of -180 deg to +180 deg; however, above about + or - 20 deg, the section characteristics were assumed to be functions only of angle of attack. A computer program is presented which evaluates the equations for a range of Mach numbers and angles of attack. Calculated results for the NACA 23012 airfoil section were compared with experimental data.

  10. Stiffness characteristics of airfoils under pulse loading

    NASA Astrophysics Data System (ADS)

    Turner, Kevin Eugene

    The turbomachinery industry continually struggles with the adverse effects of contact rubs between airfoils and casings. The key parameter controlling the severity of a given rub event is the contact load produced when the airfoil tips incur into the casing. These highly non-linear and transient forces are difficult to calculate and their effects on the static and rotating components are not well understood. To help provide this insight, experimental and analytical capabilities have been established and exercised through an alliance between GE Aviation and The Ohio State University Gas Turbine Laboratory. One of the early findings of the program is the influence of blade flexibility on the physics of rub events. The core focus of the work presented in this dissertation is to quantify the influence of airfoil flexibility through a novel modeling approach that is based on the relationship between applied force duration and maximum tip deflection. This relationship is initially established using a series of forward, non-linear and transient analyses in which simulated impulse rub loads are applied. This procedure, although effective, is highly inefficient and costly to conduct by requiring numerous explicit simulations. To alleviate this issue, a simplified model, named the pulse magnification model, is developed that only requires a modal analysis and a static analyses to fully describe how the airfoil stiffness changes with respect to load duration. Results from the pulse magnification model are compared to results from the full transient simulation method and to experimental results, providing sound verification for the use of the modeling approach. Furthermore, a unique and highly efficient method to model airfoil geometries was developed and is outlined in this dissertation. This method produces quality Finite Element airfoil definitions directly from a fully parameterized mathematical model. The effectiveness of this approach is demonstrated by comparing modal

  11. Characteristics of two sharp-nosed airfoils having reduced spinning tendencies

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1932-01-01

    According to Mr. L.D. Bell, of the Consolidated Aircraft Corporation, certain undesirable spinning characteristics of a commercial airplane were eliminated by the addition of a filler to the forward part of the wing to give it a sharp leading edge. To ascertain what aerodynamic effects result from such a change of section, two airfoils having sharp leading edges were tested in the variable-density wind tunnel. Both sections were derived by modifying the Gott. 398. The tests, which were made at a large value of the Reynolds Number, were carried to very large angles of attack to provide data for application to flight at angles of attack well beyond the stall. The characteristics of the sharp-nosed airfoils are compared with those of the normal Gott. 398 airfoil. Both of the sharp-nosed airfoils, which differ in the angle between the upper and lower surfaces at the leading edge, have about the same characteristics. As compared with the normal airfoil, the maximum lift is reduced by approximately 26 per cent, but the objectionable rapidly decreasing lift with angle of attack beyond the stall is eliminated; the profile drag of the section is slightly reduced in the range of the lift coefficient between 0.2 and 0.85, but at higher and lower lift coefficients the drag is increased.

  12. Predicted Aerodynamic Characteristics of a NACA 0015 Airfoil Having a 25% Integral-Type Trailing Edge Flap

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed

    1999-01-01

    Using the two-dimensional ARC2D Navier-Stokes flow solver analyses were conducted to predict the sectional aerodynamic characteristics of the flapped NACA-0015 airfoil section. To facilitate the analyses and the generation of the computational grids, the airfoil with the deflected trailing edge flap was treated as a single element airfoil with no allowance for a gap between the flap's leading edge and the base of the forward portion of the airfoil. Generation of the O-type computational grids was accomplished using the HYGRID hyperbolic grid generation program. Results were obtained for a wide range of Mach numbers, angles of attack and flap deflections. The predicted sectional lift, drag and pitching moment values for the airfoil were then cast in tabular format (C81) to be used in lifting-line helicopter rotor aerodynamic performance calculations. Similar were also generated for the flap. Mathematical expressions providing the variation of the sectional lift and pitching moment coefficients for the airfoil and for the flap as a function of flap chord length and flap deflection angle were derived within the context of thin airfoil theory. The airfoil's sectional drag coefficient were derived using the ARC2D drag predictions for equivalent two dimensional flow conditions.

  13. A recontoured, upper surface designed to increase the maximum lift coefficient of a modified NACA 65 (0.82) (9.9) airfoil section

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.

    1984-01-01

    A recontoured upper surface was designed to increase the maximum lift coefficient of a modified NACA 65 (0.82)(9.9) airfoil section which was tested at Mach numbers of 0.3 and 0.4 and Reynolds numbers of 2.3x10(6) and 4.3x10(6). The original 6-series section was tested for comparison with the recontoured section. The recontoured profile was found to have a higher maximum lift coefficient at all test conditions than the original airfoil. The recontoured airfoil showed less drag and nearly the same pitching moment characteristics as the original 6-series airfoil at all test conditions. The improvements found for the recontoured airfoil of the present study are similar to those found during previous investigations of recontoured 6-series airfoils with less camber.

  14. Calculation of the chordwise load distribution over airfoil sections with plain, split, or serially hinged trailing-edge flaps

    NASA Technical Reports Server (NTRS)

    Allen, H Julian

    1938-01-01

    A method is presented for the rapid calculation of the incremental chordwise normal-force distribution over an airfoil section due to the deflection of a plain flap or tab, a split flap, or a serially hinged flap. This report is intended as a supplement to NACA Report no. 631, wherein a method is presented for the calculation of the chordwise normal-force distribution over an airfoil without a flap or, as it may be considered, an airfoil with flap (or flaps) neutral. The method enables the determination of the form and magnitude of the incremental normal-force distribution to be made for an airfoil-flap combination for which the section characteristics have been determined. A method is included for the calculation of the flap normal-force and hinge-moment coefficients without necessitating a determination of the normal-force distribution.

  15. Airfoil

    NASA Technical Reports Server (NTRS)

    Derkacs, Thomas (Inventor); Fetheroff, Charles W. (Inventor); Matay, Istvan M. (Inventor); Toth, Istvan J. (Inventor)

    1983-01-01

    Although the method and apparatus of the present invention can be utilized to apply either a uniform or a nonuniform covering of material over many different workpieces, the apparatus (20) is advantageously utilized to apply a thermal barrier covering (64) to an airfoil (22) which is used in a turbine engine. The airfoil is held by a gripper assembly (86) while a spray gun (24) is effective to apply the covering over the airfoil. When a portion of the covering has been applied, a sensor (28) is utilized to detect the thickness of the covering. A control apparatus (32) compares the thickness of the covering of material which has been applied with the desired thickness and is subsequently effective to regulate the operation of the spray gun to adaptively apply a covering of a desired thickness with an accuracy of at least plus or minus 0.0015 of an inch (1.5 mils) despite unanticipated process variations.

  16. Assessment of dual-point drag reduction for an executive-jet modified airfoil section

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Mineck, Raymond E.

    1996-01-01

    This paper presents aerodynamic characteristics and pressure distributions for an executive-jet modified airfoil and discusses drag reduction relative to a baseline airfoil for two cruise design points. A modified airfoil was tested in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) for Mach numbers ranging from 0.250 to 0.780 and chord Reynolds numbers ranging from 3.0 x 10(exp 6) to 18.0 x 10(exp 6). The angle of attack was varied from minus 2 degrees to almost 10 degrees. Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The two design Mach numbers were 0.654 and 0.735, chord Reynolds numbers were 4.5 x 10(exp 6) and 8.9 x 10(exp 6), and normal-force coefficients were 0.98 and 0.51. Test data are presented graphically as integrated force and moment coefficients and chordwise pressure distributions. The maximum normal-force coefficient decreases with increasing Mach number. At a constant normal-force coefficient in the linear region, as Mach number increases an increase occurs in the slope of normal-force coefficient versus angle of attack, negative pitching-moment coefficient, and drag coefficient. With increasing Reynolds number at a constant normal-force coefficient, the pitching-moment coefficient becomes more negative and the drag coefficient decreases. The pressure distributions reveal that when present, separation begins at the trailing edge as angle of attack is increased. The modified airfoil, which is designed with pitching moment and geometric constraints relative to the baseline airfoil, achieved drag reductions for both design points (12 and 22 counts). The drag reductions are associated with stronger suction pressures in the first 10 percent of the upper surface and weakened shock waves.

  17. Effects of Compressibility on the Maximum Lift Characteristics and Spanwise Load Distribution of a 12-Foot-Span Fighter-Type Wing of NACA 230-Series Airfoil Sections

    NASA Technical Reports Server (NTRS)

    West, F E

    1945-01-01

    Lift characteristics and pressure distribution for a NACA 230 wing were investigated for an angle of attack range of from -10 to +24 degrees and Mach range of from 0.2 to 0.7. Maximum lift coefficient increased up to a Mach number of 0.3, decreased rapidly to a Mach number of 0.55, and then decreased moderately. At high speeds, maximum lift coefficient was reached at from 10 to 12 degrees beyond the stalling angle. In high-speed stalls, resultant load underwent a moderate shift outward.

  18. Improvement of aerodynamic characteristics of a thick airfoil with a vortex cell in sub- and transonic flow

    NASA Astrophysics Data System (ADS)

    Isaev, Sergey; Baranov, Paul; Popov, Igor; Sudakov, Alexander; Usachov, Alexander

    2017-03-01

    The modified SST model (2005) is verified using Rodi- Leschziner-Isaev's approach and the multiblock computational technologies are validated in the VP2/3 code on different-structure overlapping grids by comparing the numerical predictions with the experimental data on transonic flow around an NACA0012 airfoil at an angle of attack of 4o for M=0.7 and Re=4×106. It is proved that the aerodynamic characteristics of a thick (20% of the chord) MQ airfoil mounted at an angle of attack of 2o for Re=107 and over the Mach number range 0.3-0.55 are significantly improved because an almost circular small-size (0.12) vortex cell with a defined volumetric flow rate coefficient of 0.007 during slot suction has been located on the upper airfoil section and an intense trapped vortex has been formed in it. A detailed analysis of buffeting within the self-oscillatory regime of flow around the MQ airfoil with a vortex cell has demonstrated the periodic changes in local and integral characteristics; the lift and the aerodynamic efficiency remain quite high, but inferior to the similar characteristics at M=0.55. It is found that the vortex cell at M=0.7 is inactive, and the aerodynamic characteristics of the MQ airfoil with a vortex cell are close to those of a smooth airfoil without a cell.

  19. Horizontal axis wind turbine post stall airfoil characteristics synthesization

    NASA Technical Reports Server (NTRS)

    Tangler, James L.; Ostowari, Cyrus

    1995-01-01

    Blade-element/momentum performance prediction codes are routinely used for wind turbine design and analysis. A weakness of these codes is their inability to consistently predict peak power upon which the machine structural design and cost are strongly dependent. The purpose of this study was to compare post-stall airfoil characteristics synthesization theory to a systematically acquired wind tunnel data set in which the effects of aspect ratio, airfoil thickness, and Reynolds number were investigated. The results of this comparison identified discrepancies between current theory and the wind tunnel data which could not be resolved. Other factors not previously investigated may account for these discrepancies and have a significant effect on peak power prediction.

  20. Predicting aerodynamic characteristic of typical wind turbine airfoils using CFD

    SciTech Connect

    Wolfe, W.P.; Ochs, S.S.

    1997-09-01

    An investigation was conducted into the capabilities and accuracy of a representative computational fluid dynamics code to predict the flow field and aerodynamic characteristics of typical wind-turbine airfoils. Comparisons of the computed pressure and aerodynamic coefficients were made with wind tunnel data. This work highlights two areas in CFD that require further investigation and development in order to enable accurate numerical simulations of flow about current generation wind-turbine airfoils: transition prediction and turbulence modeling. The results show that the laminar-to turbulent transition point must be modeled correctly to get accurate simulations for attached flow. Calculations also show that the standard turbulence model used in most commercial CFD codes, the k-e model, is not appropriate at angles of attack with flow separation. 14 refs., 28 figs., 4 tabs.

  1. Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

    NASA Technical Reports Server (NTRS)

    Zaman, Khairul; Fagan, Amy; Mankbadi, Mina

    2016-01-01

    An experimental investigation of tip vortex flow from a NACA0012 airfoil, pitched periodically at various frequencies, is conducted in a low-speed wind tunnel. Initially, data for stationary airfoil held fixed at various angles-of-attack are gathered. Flow visualization pictures as well as detailed cross-sectional properties areobtained at various streamwise locations using hot-wire anemometry. Data include mean velocity, streamwise vorticity as well as various turbulent stresses. Preliminary data are also acquired for periodically pitched airfoil. These results are briefly presented in this extended abstract.

  2. Tests of N-85, N-86 and N-87 airfoil sections in the 11-inch high speed wind tunnel

    NASA Technical Reports Server (NTRS)

    Stack, John; Lindsey, W F

    1938-01-01

    Three airfoils, the N-85, the N-86, and the N-87, were tested at the request of the Bureau of Aeronautics, Navy Department, to determine the suitability of these sections for use as propeller-blade sections. Further tests of the NACA 0009-64 airfoil were also made to measure the aerodynamic effect of thickening the trailing edge in accordance with current propeller practice. The N-86 and the N-87 airfoils appear to be nearly equivalent aerodynamically and both are superior to the N-85 airfoil. Comparison of those airfoils with the previously developed NACA 2409-34 airfoils indicate that the NACA 2409-34 is superior, particularly at high speeds. Thickening the trailing edge appears to have a detrimental effect, although the effect may be small if the trailing-edge radius is less than 0.5 percent of the cord. The N-86 and the N-87 airfoils appear to be nearly equivalent.

  3. Large-scale aerodynamic characteristics of airfoils as tested in the variable density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Anderson, Raymond F

    1931-01-01

    In order to give the large-scale characteristics of a variety of airfoils in a form which will be of maximum value, both for airplane design and for the study of airfoil characteristics, a collection has been made of the results of airfoil tests made at full-scale values of the reynolds number in the variable density wind tunnel of the National Advisory Committee for Aeronautics. They have been corrected for tunnel wall interference and are presented not only in the conventional form but also in a form which facilitates the comparison of airfoils and from which corrections may be easily made to any aspect ratio. An example showing the method of correcting the results to a desired aspect ratio has been given for the convenience of designers. In addition, the data have been analyzed with a view to finding the variation of the aerodynamic characteristics of airfoils with their thickness and camber.

  4. Comparisons of Theoretical Methods for Predicting Airfoil Aerodynamic Characteristics

    DTIC Science & Technology

    2010-08-01

    Airfoil ,” Airfoils , U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-107, August 2010. [2] Somers, D.M. and...Maughmer, M.D., “Design and Experimental Results for the S407 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D...S414 Airfoil ,” U.S. Army Aviation Research, Development and Engineering Command, RDECOM TR 10-D-112, August 2010. [5] Somers, D.M. and Maughmer

  5. NASA low- and medium-speed airfoil development

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.; Whitcomb, R. T.

    1979-01-01

    The status of NASA low and medium speed airfoil research is discussed. Effects of airfoil thickness-chord ratios varying from 9 percent to 21 percent on the section characteristics for a design lift coefficient of 0.40 are presented for the initial low speed family of airfoils. Also, modifications to the 17-percent low-speed airfoil to reduce the pitching-moment coefficient and to the 21-percent low speed airfoil results are shown for two new medium speed airfoils with thickness ratios of 13 percent and 17 percent and design-lift coefficients of 0.30. Applications of NASA-developed airfoils to general aviation aircraft are summarized.

  6. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

    Stivers, Louis S.; Abbott, Ira H.; von Doenhoff, Albert E.

    1945-01-01

    Recent airfoil data for both flight and wind-tunnel tests have been collected and correlated insofar as possible. The flight data consist largely of drag measurements made by the wake-survey method. Most of the data on airfoil section characteristics were obtained in the Langley two-dimensional low-turbulence pressure tunnel. Detail data necessary for the application of NACA 6-serles airfoils to wing design are presented in supplementary figures, together with recent data for the NACA 24-, 44-, and 230-series airfoils. The general methods used to derive the basic thickness forms for NACA 6- and 7-series airfoils and their corresponding pressure distributions are presented. Data and methods are given for rapidly obtaining the approximate pressure distributions for NACA four-digit, five-digit, 6-, and 7-series airfoils. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed. The data indicate that the effects of surface condition on the lift and drag characteristics are at least as large as the effects of the airfoil shape and must be considered in airfoil selection and the prediction of wing characteristics. Airfoils permitting extensive laminar flow, such as the NACA 6-series airfoils, have much lower drag coefficients at high speed and cruising lift coefficients than earlier types-of airfoils if, and only if, the wing surfaces are sufficiently smooth and fair. The NACA 6-series airfoils also have favorable critical-speed characteristics and do not appear to present unusual problems associated with the application of high-lift and lateral-control devices. Much of the data given in the NACA Advance Confidential Report entitled "Preliminary Low-Drag-Airfoil and Flap Data from

  7. Wind-Tunnel Investigation of the Lift Characteristics of an NACA 27-212 Airfoil Equipped with Two Types of Flap, Special Report

    NASA Technical Reports Server (NTRS)

    Swanson, Robert S.; Schuldenfrei, Marvin J.

    1940-01-01

    An investigation has been made in the NACA 7- by 10-foot wind tunnel of a large chord NACA 27-212 airfoil with a 20% chord split flap and with two arrangements of a 25.66% chord slotted flap to determine the section lift characteristics as affected by flap deflection for the split flap and as affected by flap deflection, flap position, and slot shape for the slotted flap. For the two arrangements of the slotted flap, the flap positions for maximum section lift are given. Comparable data on the NACA 23012 airfoil equipped with similar flaps are also given. On the basis of maximum section lift coefficient, the slotted flap with an easy slot entry was slightly better than either the split flap or the slotted flap with a sharp slot entry. With both types of flap the decrease in the angle of attack, for maximum section lift coefficient, with flap deflection is large for the NACA 27-212 airfoil as compared with the NACA 23012 airfoil. Also with both flaps, the maximum section lift coefficient obtained with flaps is much lower for the NACA 27-212 airfoil than for the NACA 23012 airfoil.

  8. Tests on a CAST 7 two-dimensional airfoil in a streamlining test section

    NASA Technical Reports Server (NTRS)

    Goodyear, M. J.

    1984-01-01

    A unique opportunity has arisen to test one and the same airfoil model of CAST-7 section in two wind tunnels having adaptive walled test sections. The tunnels are very similar in terms of size and the available range of test conditions, but differ principally in their wall setting algorithms. Detailed data from the tests of the model in the Southampton tunnel, are included with comparisons between various sources of data indicating that both adaptive walled test sections provide low interference test conditions.

  9. Scale Effect on Clark Y Airfoil Characteristics from NACA Full-Scale Wind-Tunnel Tests

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe

    1935-01-01

    This report presents the results of wind tunnel tests conducted to determine the aerodynamic characteristics of the Clark Y airfoil over a large range of Reynolds numbers. Three airfoils of aspect ratio 6 and with 4, 6, and 8 foot chords were tested at velocities between 25 and 118 miles per hour, and the characteristics were obtained for Reynolds numbers (based on the airfoil chord) in the range between 1,000,000 and 9,000,000 at the low angles of attack, and between 1,000,000 and 6,000,000 at maximum lift. With increasing Reynolds number the airfoil characteristics are affected in the following manner: the drag at zero lift decreases, the maximum lift increases, the slope of the lift curve increases, the angle of zero lift occurs at smaller negative angles, and the pitching moment at zero lift does not change appreciably.

  10. Robust, optimal subsonic airfoil shapes

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan (Inventor)

    2008-01-01

    Method system, and product from application of the method, for design of a subsonic airfoil shape, beginning with an arbitrary initial airfoil shape and incorporating one or more constraints on the airfoil geometric parameters and flow characteristics. The resulting design is robust against variations in airfoil dimensions and local airfoil shape introduced in the airfoil manufacturing process. A perturbation procedure provides a class of airfoil shapes, beginning with an initial airfoil shape.

  11. Cylinder wake influence on the tonal noise and aerodynamic characteristics of a NACA0018 airfoil

    NASA Astrophysics Data System (ADS)

    Takagi, Y.; Fujisawa, N.; Nakano, T.; Nashimoto, A.

    2006-11-01

    The influence of cylinder wake on discrete tonal noise and aerodynamic characteristics of a NACA0018 airfoil is studied experimentally in a uniform flow at a moderate Reynolds number. The experiments are carried out by measuring sound pressure levels and spectrum, separation and the reattachment points, pressure distribution, fluid forces, mean-flow and turbulence characteristics around the airfoil with and without the cylinder wake. Present results indicate that the tonal noise from the airfoil is suppressed by the influence of the cylinder wake and the aerodynamic characteristics are improved in comparison with the case without the cylinder wake. These are mainly due to the separation control of boundary layers over the airfoil caused by the wake-induced transition, which is observed by surface flow visualization with liquid- crystal coating. The PIV measurements of the flow field around the airfoil confirm that highly turbulent velocity fluctuation of the cylinder wake induces the transition of the boundary layers and produces an attached boundary layer over the airfoil. Then, the vortex shedding phenomenon near the trailing edge of pressure surface is removed by the influence of the wake and results in the suppression of tonal noise.

  12. Evaluation of Airfoil Dynamic Stall Characteristics for Maneuverability

    NASA Technical Reports Server (NTRS)

    Bousman, William G.; Aiken, Edwin W. (Technical Monitor)

    2000-01-01

    In severe maneuvers, out of necessity for a military aircraft or inadvertently for a civil aircraft, a helicopter airfoil will stall in a dynamic manner and provide lift beyond what would be calculated based on static airfoil tests. The augmented lift that occurs in dynamic stall is related to a vortex that is shed near the leading edge of the airfoil. However, directly related to the augmented lift that results from the dynamic stall vortex are significant penalties in pitching moment and drag. An understanding of the relationship between the augmented lift in dynamic stall and the associated moment and drag penalties is the purpose of this paper. This relationship is characterized using data obtained in two-dimensional wind tunnel tests and related to the problem of helicopter maneuverability.

  13. Dynamic Stall Characteristics of Drooped Leading Edge Airfoils

    NASA Technical Reports Server (NTRS)

    Sankar, Lakshmi N.; Sahin, Mehmet; Gopal, Naveen

    2000-01-01

    Helicopters in high-speed forward flight usually experience large regions of dynamic stall over the retreating side of the rotor disk. The rapid variations in the lift and pitching moments associated with the stall process can result in vibratory loads, and can cause fatigue and failure of pitch links. In some instances, the large time lag between the aerodynamic forces and the blade motion can trigger stall flutter. A number of techniques for the alleviation of dynamic stall have been proposed and studied by researchers. Passive and active control techniques have both been explored. Passive techniques include the use of high solidity rotors that reduce the lift coefficients of individual blades, leading edge slots and leading edge slats. Active control techniques include steady and unsteady blowing, and dynamically deformable leading edge (DDLE) airfoils. Considerable amount of experimental and numerical data has been collected on the effectiveness of these concepts. One concept that has not received as much attention is the drooped-leading edge airfoil idea. It has been observed in wind tunnel studies and flight tests that drooped leading edge airfoils can have a milder dynamic stall, with a significantly milder load hysteresis. Drooped leading edge airfoils may not, however, be suitable at other conditions, e.g. in hover, or in transonic flow. Work needs to be done on the analysis and design of drooped leading edge airfoils for efficient operation in a variety of flight regimes (hover, dynamic stall, and transonic flow). One concept that is worthy of investigation is the dynamically drooping airfoil, where the leading edge shape is changed roughly once-per-rev to mitigate the dynamic stall.

  14. An Experimental Study of Dynamic Stall on Advanced Airfoil Sections. Volume 1. Summary of the Experiment.

    DTIC Science & Technology

    1982-07-01

    Aeronautics and United States Army Space Administration Aviation Research and Ames Remrch Cente Development Command Moffett Field. California 94035 St...appear to be more important than airfoil shape in determining the dynamic- stall airloads. 1. INTRODUCTION Retreating- blade stall limits the high-speed...12.2% Thick R.A.E. Aerofoil Section. RAE Technical Report 68303, Royal Aircraft Establishment, Farnborough Hants, England, Jan. 1969. 14. Fromme, J. A

  15. Characteristics of NACA 4400R Series Rectangular and Tapered Airfoils, Including the Effect of Split Flaps

    NASA Technical Reports Server (NTRS)

    Greenberg, Harry

    1941-01-01

    At the request of the Bureau of Aeronautics, Navy Department, tests were made in the variable-density wind tunnel of a tapered wing of 3-10-18 plan form and based on the NACA 4400R series sections. The wing was also tested with 0.2 chord spit flaps, deflected 60 deg span ratios of 0.3, 0.5, 0.7 and 1.0 respectively. In order to get data from which to calculate the characteristics of the flapped wing, the investigation was extended to include tests of the four rectangular airfoils of the NACA 4400R series (4409R, 4412R, 4415R, and 4418R) with full-span 0.2 chord, trailing edge split flaps deflected 60 deg.

  16. A flight investigation of blade-section aerodynamics for a helicopter main rotor having RC-SC2 airfoil sections

    NASA Technical Reports Server (NTRS)

    Morris, C. E. K., Jr.

    1982-01-01

    Pressure data at 90 percent blade radius for a helicopter main rotor with RC-SC2 blade sections was obtained. Concurrent measurements were made of vehicle flight state, performance and some rotor loads. The test envelope included hover, level flight from about 65 to 144 knots, climb and descent, and collective fixed maneuvers. Airfoil pressure distributions obtained in flight agree with those theoretical calculations for two dimensional, steady flow.

  17. Aerodynamic characteristics of a rotorcraft airfoil designed for the tip region of a main rotor blade

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W.

    1991-01-01

    A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of a new rotorcraft airfoil designed for application to the tip region (stations outboard of 85 pct. radius) of a helicopter main rotor blade. The new airfoil, the RC(6)-08, and a baseline airfoil, the RC(3)-08, were investigated in the Langley 6- by 28-inch transonic tunnel at Mach numbers from 0.37 to 0.90. The Reynolds number varied from 5.2 x 10(exp 6) at the lowest Mach number to 9.6 x 10(exp 6) at the highest Mach number. Some comparisons were made of the experimental data for the new airfoil and the predictions of a transonic, viscous analysis code. The results of the investigation indicate that the RC(6)-08 airfoil met the design goals of attaining higher maximum lift coefficients than the baseline airfoil while maintaining drag divergence characteristics at low lift and pitching moment characteristics nearly the same as those of the baseline airfoil. The maximum lift coefficients of the RC(6)-08 varied from 1.07 at M=0.37 to 0.94 at M=0.52 while those of the RC(3)-08 varied from 0.91 to 0.85 over the same Mach number range. At lift coefficients of -0.1 and 0, the drag divergence Mach number of both the RC(6)-08 and the RC(3)-08 was 0.86. The pitching moment coefficients of the RC(6)-08 were less negative than those of the RC(3)-08 for Mach numbers and lift coefficients typical of those that would occur on a main rotor blade tip at high forward speeds on the advancing side of the rotor disk.

  18. CFD aerodynamic analysis of non-conventional airfoil sections for very large rotor blades

    NASA Astrophysics Data System (ADS)

    Papadakis, G.; Voutsinas, S.; Sieros, G.; Chaviaropoulos, T.

    2014-12-01

    The aerodynamic performance of flat-back and elliptically shaped airfoils is analyzed on the basis of CFD simulations. Incompressible and low-Mach preconditioned compressible unsteady simulations have been carried out using the k-w SST and the Spalart Allmaras turbulence models. Time averaged lift and drag coefficients are compared to wind tunnel data for the FB 3500-1750 flat back airfoil while amplitudes and frequencies are also recorded. Prior to separation averaged lift is well predicted while drag is overestimated keeping however the trend in the tests. The CFD models considered, predict separation with a 5° delay which is reflected on the load results. Similar results are provided for a modified NACA0035 with a rounded (elliptically shaped) trailing edge. Finally as regards the dynamic characteristics in the load signals, there is fair agreement in terms of Str number but significant differences in terms of lift and drag amplitudes.

  19. Experimental investigation of a 10-percent-thick helicopter rotor airfoil section designed with a viscous transonic analysis code

    NASA Technical Reports Server (NTRS)

    Noonan, K. W.

    1981-01-01

    An investigation was conducted in the Langley 6- by 28-Inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics of a 10-percent-thick helicopter rotor airfoil at Mach numbers from 0.33 to 0.87 and respective Reynolds numbers from 4.9 x 10 to the 6th to 9.8 x 10 to the 6th. This airfoil, designated the RC-10(N)-1, was also investigated at Reynolds numbers from 3.0 x 10 to the 6th to 7.3 x 10 to the 6th at respective Mach numbers of 0.33 to 0.83 for comparison wit the SC 1095 (with tab) airfoil. The RC-10(N)-1 airfoil was designed by the use of a viscous transonic analysis code. The results of the investigation indicate that the RC-10(N)-1 airfoil met all the design goals. At a Reynolds number of about 9.4 x 10 to the 6th the drag divergence Mach number at zero normal-force coefficient was 0.815 with a corresponding pitching-moment coefficient of zero. The drag divergence Mach number at a normal-force coefficient of 0.9 and a Reynolds number of about 8.0 x 10 to the 6th was 0.61. The drag divergence Mach number of this new airfoil was higher than that of the SC 1095 airfoil at normal-force coefficients above 0.3. Measurements in the same wind tunnel at comparable Reynolds numbers indicated that the maximum normal-force coefficient of the RC-10(N)-1 airfoil was higher than that of the NACA 0012 airfoil for Mach numbers above about 0.35 and was about the same as that of the SC 1095 airfoil for Mach numbers up to 0.5.

  20. Characteristics of merging shear layers and turbulent wakes of a multi-element airfoil

    NASA Technical Reports Server (NTRS)

    Adair, Desmond; Horne, W. Clifton

    1988-01-01

    Flow characteristics in the vicinity of the trailing edge of a single-slotted airfoil flap are presented and analyzed. The experimental arrangement consisted of a NACA 4412 airfoil equipped with a NACA 4415 flap whose angle of deflection was 21.8 deg. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The flow was complicated by the presence of a strong, initially inviscid jet emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer. Downstream of the flap, the airfoil and flap wakes fully merged to form an asymmetrical curved wake. The airfoil configuration was tested at an angle of attack of 8.2 deg, at a Mach number of 0.09, and a chord based Reynolds number of 1.8 x 10 to the 6th power in the Ames Research Center 7- by 10-Foot Wind Tunnel. Surface pressure measurements were made on the airfoil and flap and on the wind tunnel roof and floor. It was estimated that the wall interference increased the C sub L by 7 percent and decreased the C sub M by 4.5 percent. Velocity characteristics were quantified using hot-wire anemometry in regions of flow with preferred direction and low turbulence intensity. A 3-D laser velocimeter was used in regions of flow recirculation and relatively high turbulence intensity.

  1. Effect of advanced rotorcraft airfoil sections on the hover performance of a small-scale rotor model

    NASA Technical Reports Server (NTRS)

    Althoff, Susan L.

    1988-01-01

    A hover test was conducted on a small scale rotor model for two sets of tapered rotor blades. The baseline rotor blade set used a NACA 0012 airfoil section, whereas the second rotor blade set had advanced rotorcraft airfoils distributed along the radius. The experiment was conducted for a range of thrust coefficients and tip speeds, and the data were compared to the predictions of three analytical methods. The data show the advantage of the advanced airfoils at the higher rotor thrust levels; two of the analyses predicted the correct data trends.

  2. The Aerodynamic Characteristics of a Slotted Clark Y Wing as Affected by the Auxiliary Airfoil Position

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Shortal, Joseph A

    1932-01-01

    Aerodynamic force tests on a slotted Clark Y wing were conducted in a vertical wind tunnel to determine the best position for a given auxiliary airfoil with respect to the main wing. A systematic series of 100 changes in location of the auxiliary airfoil were made to cover all the probable useful ranges of slot gap, slot width, and slot depth. The results of the investigation may be applied to the design of automatic or controlled slots on wings with geometric characteristics similar to the wing tested. The best positions of the auxiliary airfoil were covered by the range of the tests, and the position for desired aerodynamic characteristics may easily be obtained from charts prepared especially for the purpose.

  3. Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partial-Span Leading-Edge Slat for Various Modes of Thermal Ice Protection

    NASA Technical Reports Server (NTRS)

    VonGlahn, Uwe H.; Gray, Vernon H.

    1954-01-01

    The effects of primary and runback ice formations on the section drag of a 36 deg swept NACA 63A-009 airfoil section with a partial-span leading-edge slat were studied over a range of angles of attack from 2 to 8 deg and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.39 to 1.23 grams per cubic meter and datum air temperatures from 10 to 25 F. The results with slat retracted showed that glaze-ice formations caused large and rapid increases in section drag coefficient and that the rate of change in section drag coefficient for the swept 63A-009 airfoil was about 2-1 times that for an unswept 651-212 airfoil. Removal of the primary ice formations by cyclic de-icing caused the drag to return almost to the bare-airfoil drag value. A comprehensive study of the slat icing and de-icing characteristics was prevented by limitations of the heating system and wake interference caused by the slat tracks and hot-gas supply duct to the slat. In general, the studies showed that icing on a thin swept airfoil will result in more detrimental aerodynamic characteristics than on a thick unswept airfoil.

  4. Wall interference tests of a CAST 10-2/DOA 2 airfoil in an adaptive-wall test section

    NASA Technical Reports Server (NTRS)

    Mineck, Raymond E.

    1987-01-01

    A wind-tunnel investigation of a CAST 10-2/DOA 2 airfoil model has been conducted in the adaptive-wall test section of the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) and in the National Aeronautical Establishment High Reynolds Number Two-Dimensional Test Facility. The primary goal of the tests was to assess two different wall-interference correction techniques: adaptive test-section walls and classical analytical corrections. Tests were conducted over a Mach number range from 0.3 to 0.8 and over a chord Reynolds number range from 6 million to 70 million. The airfoil aerodynamic characteristics from the tests in the 0.3-m TCT have been corrected for wall interference by the movement of the adaptive walls. No additional corrections for any residual interference have been applied to the data, to allow comparison with the classically corrected data from the same model in the conventional National Aeronautical Establishment facility. The data are presented graphically in this report as integrated force-and-moment coefficients and chordwise pressure distributions.

  5. Wind-tunnel investigation of effects of trailing-edge geometry on a NASA supercritical airfoil section

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1971-01-01

    Wind-tunnel tests have been conducted at Mach numbers from 0.60 to 0.81 to determine the effects of trailing-edge geometry on the aerodynamic characteristics of a NASA supercritical airfoil shape. Variations in trailing-edge thicknesses from 0 to 1.5 percent of the chord and a cavity in the trailing edge were investigated with airfoils with maximum thicknesses of 10 and 11 percent of the chord.

  6. Adjustment of the k-ω SST turbulence model for prediction of airfoil characteristics near stall

    NASA Astrophysics Data System (ADS)

    Matyushenko, A. A.; Garbaruk, A. V.

    2016-11-01

    A version of k-ra SST turbulence model adjusted for flow around airfoils at high Reynolds numbers is presented. The modified version decreases eddy viscosity and significantly improves the accuracy of prediction of aerodynamic characteristics in a wide range of angles of attack. However, considered reduction of eddy viscosity destroys calibration of the model, which leads to decreasing accuracy of skin-friction coefficient prediction even for relatively simple wall-bounded turbulent flows. Therefore, the area of applicability of the suggested modification is limited to flows around airfoils.

  7. Effects of Leading Edge Defect on the Aerodynamic and Flow Characteristics of an S809 Airfoil

    PubMed Central

    Wang, Yan; Zheng, Xiaojing; Hu, Ruifeng; Wang, Ping

    2016-01-01

    Background Unexpected performance degradation occurs in wind turbine blades due to leading edge defect when suffering from continuous impacts with rain drops, hails, insects, or solid particles during its operation life. To assess this issue, this paper numerically investigates the steady and dynamic stall characteristics of an S809 airfoil with various leading edge defects. More leading edge defect sizes and much closer to practical parameters are investigated in the paper. Methodology Numerical computation is conducted using the SST k-ω turbulence model, and the method has been validated by comparison with existed published data. In order to ensure the calculation convergence, the residuals for the continuity equation are set to be less than 10−7 and 10−6 in steady state and dynamic stall cases. The simulations are conducted with the software ANSYS Fluent 13.0. Results It is found that the characteristics of aerodynamic coefficients and flow fields are sensitive to leading edge defect both in steady and dynamic conditions. For airfoils with the defect thickness of 6%tc, leading edge defect has a relative small influence on the aerodynamics of S809 airfoil. For other investigated defect thicknesses, leading edge defect has much greater influence on the flow field structures, pressure coefficients and aerodynamic characteristics of airfoil at relative small defect lengths. For example, the lift coefficients decrease and drag coefficients increase sharply after the appearance of leading edge defect. However, the aerodynamic characteristics could reach a constant value when the defect length is large enough. The flow field, pressure coefficient distribution and aerodynamic coefficients do not change a lot when the defect lengths reach to 0.5%c,1%c, 2%c and 3%c with defect thicknesses of 6%tc, 12%tc,18%tc and 25%tc, respectively. In addition, the results also show that the critical defect length/thickness ratio is 0.5, beyond which the aerodynamic characteristics

  8. Flatback airfoil wind tunnel experiment.

    SciTech Connect

    Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.

    2008-04-01

    A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.

  9. General theory of airfoil sections having arbitrary shape or pressure distribution

    NASA Technical Reports Server (NTRS)

    Allen, H Julian

    1945-01-01

    In this report a theory of thin airfoils of small camber is developed which permits either the velocity distribution corresponding to a given airfoil shape, or the airfoil shape corresponding to a given velocity distribution to be calculated. The procedures to be employed in these calculations are outlined and illustrated with suitable examples.

  10. Aerodynamic characteristics of two rotorcraft airfoils designed for application to the inboard region of a main rotor blade

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W.

    1990-01-01

    A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of two new rotorcraft airfoils designed especially for application to the inboard region of a helicopter main rotor blade. The two new airfoils, the RC(4)-10 and RC(5)-10, and a baseline airfoil, the VR-7, were all studied in the Langley Transonic Tunnel at Mach nos. from about 0.34 to 0.84 and at Reynolds nos. from about 4.7 to 9.3 x 10 (exp 6). The VR-7 airfoil had a trailing edge tab which is deflected upwards 4.6 degs. In addition, the RC(4)-10 airfoil was studied in the Langley Low Turbulence Pressure Tunnel at Mach nos. from 0.10 to 0.44 and at Reynolds nos. from 1.4 to 5.4 x 10 (exp 6) respectively. Some comparisons were made of the experimental data for the new airfoils and the predictions of two different theories. The results of this study indicates that both of the new airfoils offer advantages over the baseline airfoil. These advantages are discussed.

  11. A flight investigation of blade section aerodynamics for a helicopter main rotor having NLR-1T airfoil sections

    NASA Technical Reports Server (NTRS)

    Morris, C. E. K., Jr.; Stevens, D. D.; Tomaine, R. L.

    1980-01-01

    A flight investigation was conducted using a teetering-rotor AH-1G helicopter to obtain data on the aerodynamic behavior of main-rotor blades with the NLR-1T blade section. The data system recorded blade-section aerodynamic pressures at 90 percent rotor radius as well as vehicle flight state, performance, and loads. The test envelope included hover, forward flight, and collective-fixed maneuvers. Data were obtained on apparent blade-vortex interactions, negative lift on the advancing blade in high-speed flight and wake interactions in hover. In many cases, good agreement was achieved between chordwise pressure distributions predicted by airfoil theory and flight data with no apparent indications of blade-vortex interactions.

  12. The method of complex characteristics for transonic airfoil design, with an application to compressors

    NASA Technical Reports Server (NTRS)

    Bledsoe, M.; Garabedian, P.

    1985-01-01

    The use of mathematical models to study physical problems of current interest to aeronautical engineers has been made possible by the development of numerical techniques to compute solutions of the differential equations of transonic aerodynamics. These advances have encouraged the improvement of supercritical wing technology. A method to determined steady, shockless flow of an inviscid, compressible fluid past a cascade of airfoils in the (x,y)-plane is considered, taking into account also the case of an isolated airfoil. The method of complex characteristics solves the equations in the hodograph plane by extending all variables into the complex domain, where the notion of type is no longer significant. Attention is given to the mathematical background, the method of complex characteristics, and numerical calculations.

  13. Wind Tunnel Aerodynamic Characteristics of a Transport-type Airfoil in a Simulated Heavy Rain Environment

    NASA Technical Reports Server (NTRS)

    Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Gentry, Garl L., Jr.; Melson, W. Edward, Jr.

    1992-01-01

    The effects of simulated heavy rain on the aerodynamic characteristics of an NACA 64-210 airfoil section equipped with leading-and trailing-edge high-lift devices were investigated in the Langley 14- by 22-Foot Subsonic Tunnel. The model had a chord of 2.5 ft, a span of 8 ft, and was mounted on the tunnel centerline between two large endplates. Aerodynamic measurements in and out of the simulated rain environment were obtained for dynamic pressures of 30 and 50 psf and an angle-of-attack range of 0 to 20 degrees for the cruise configuration. The rain intensity was varied to produce liquid water contents ranging from 16 to 46 gm/cu m. The results obtained for various rain intensity levels and tunnel speeds showed significant losses in maximum lift capability and increases in drag for a given lift as the liquid water content was increased. The results obtained on the landing configuration also indicate a progressive decrease in the angle of attack at which maximum lift occurred and an increase in the slope of the pitching-moment curve as the liquid water content was increased. The sensitivity of test results to the effects of the water surface tension was also investigated. A chemical was introduced into the rain environment that reduced the surface tension of water by a factor of 2. The reduction in the surface tension of water did not significantly alter the level of performance losses for the landing configuration.

  14. The role of airfoil geometry in minimizing the effect of insect contamination of laminar flow sections

    NASA Technical Reports Server (NTRS)

    Maresh, J. L.; Bragg, M. B.

    1984-01-01

    A method has been developed to predict the contamination of an airfoil by insects and the resultant performance penalty. Insect aerodynamics have been modeled and the impingement of insects on an airfoil are solved by calculating their trajectories. Upon impact, insect rupture and the resulting height of the debris is determined based on experimental data. A boundary layer analysis is performed to determine which insects cause boundary layer transition and the resultant drag penalty. A contaminated airfoil figure of merit is presented to be used to compare airfoil susceptibility. Results show that the insect contamination effects depend on accretion conditions, airfoil angle of attack and Reynolds number. The importance of the stagnation region to designing airfoils for minimum drag penalties is discussed.

  15. The Propeller and Cooling-Air-Flow Characteristics of a Twin-Engine Airplane Model Equipped with NACA D sub s -Type Cowlings and with Propellers of NACA 16-Series Airfoil Sections

    DTIC Science & Technology

    1944-09-01

    relatively low Reynolds number of an isolated nacelle fitted with an NAOA C-type cowling and did not include sufficient measurements of the internal flow to...through the nscellr was simulated by a baffle inside the cowling. The conductivity of the cowling was deter- mined from measurements of the quantity...of the various cowling flaps end variable-lenpth cowling skirts on the ratio of eowlinr-ex.it area to nacelle crocs -sectional area la shown in

  16. Wind-Tunnel Investigation of Control-Surface Characteristics. 15 - Various Contour Modifications of a 0.30-Airfoil-Chord Plain Flap on an NACA 66(215)-014 Airfoil

    DTIC Science & Technology

    1943-12-01

    a plain flap on a low —drag airfoil were not • • .•ü;.V;.:-’ ;•»**;’•.••••<«**. •’ .•• V-:.--^i* -I’-••»•’*;w .-•; ’.••.<• % •v — i f...thick low —drag airfoil and on 9— and 15—percent- thick conventional airfoils. Other modifications have included the use of a...airplanes require the use of airfoil sections with low peak pressures, such as low —drag sec- tions, for tail surfaces to

  17. Numerical and Experimental Study on Aerodynamic Characteristics of Basic Airfoils at Low Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Hirata, Katsuya; Kawakita, Masatoshi; Iijima, Takayoshi; Koga, Mitsuhiro; Kihira, Mitsuhiko; Funaki, Jiro

    The aerodynamic characteristics of airfoils have been researched in higher Reynolds-number ranges more than 106, in a historic context closely related with the developments of airplanes and fluid machineries in the last century. However, our knowledge is not enough at low and middle Reynolds-number ranges. So, in the present study, we investigate such basic airfoils as a NACA0015, a flat plate and the flat plates with modified fore-face and after-face geometries at Reynolds number Re < 1.0×105, using two- and three-dimensional computations together with wind-tunnel and water-tank experiments. As a result, we have revealed the effect of the Reynolds number Re upon the minimum drag coefficient CDmin. Besides, we have shown the effects of attack angle α upon various aerodynamic characteristics such as the lift coefficient CL, the drag coefficient CD and the lift-to-drag ratio CL/CD at Re = 1.0×102, discussing those effects on the basis of both near-flow-field information and surface-pressure profiles. Such results suggest the importance of sharp leading edges, which implies the possibility of an inversed NACA0015. Furthermore, concerning the flat-plate airfoil, we investigate the influences of fore-face and after-face geometries upon such effects.

  18. Low-speed aerodynamic characteristics of a 17-percent-thick medium speed airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beaseley, W. D.

    1980-01-01

    Wind tunnel tests were conducted to determine the low speed two dimensional aerodynamic characteristics of a 17 percent thick medium speed airfoil (MS(1)-0317) designed for general aviation applications. The results were compared with data for the 17 percent thick low speed airfoil (LS(1)-0417) and the 13 percent thick medium speed airfoil (MS(1)-0313). Theoretical predictions of the drag rise characteristics of this airfoil are also provided. The tests were conducted in the Langley low turbulence pressure tunnel over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2 million to 12 million, and an angle of attack range from about -8 to 20 deg.

  19. Effects of thickness on the aerodynamic characteristics of an initial low-speed family of airfoils for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1976-01-01

    Wind tunnel tests were conducted to determine the effects of airfoil thickness-ratio on the low speed aerodynamic characteristics of an initial family of airfoils. The results were compared with theoretical predictions obtained from a subsonic viscous method. The tests were conducted over a Mach number range from 0.10 to 0.28. Chord Reynolds numbers varied from about 2.0 x 1 million to 9.0 x 1 million.

  20. Investigation of the Kline-Fogleman airfoil section for rotor blade applications

    NASA Technical Reports Server (NTRS)

    Lumsdaine, E.; Johnson, W. S.; Fletcher, L. M.; Peach, J. E.

    1974-01-01

    Wind tunnel tests of a wedgeshaped airfoil with sharp leading edge and a spanwise step were conducted. The airfoil was tested with variations of the following parameters: (1) Reynolds number, (2) step location, (3) step shape, (4) apex angle, and (5) with the step on either the upper or lower surface. The results are compared with a flat plate and with wedge airfoils without a step having the same aspect ratio. Water table tests were conducted for flow visualization and it was determined that the flow separates from the upper surface at low angles of attack. The wind tunnel tests show that the lift/drag ratio of the airfoil is lower than for a flat plate and the pressure data show that the airfoil derives its lift in the same manner as a flat plate.

  1. Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient

    NASA Technical Reports Server (NTRS)

    Bingham, G. J.; Chen, A. W.

    1972-01-01

    An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.

  2. Viscous Thin Airfoil Theory

    DTIC Science & Technology

    1980-02-01

    the elliptic cross section is considered to be more representative of the NACA 64A010 airfoil with boundary layer displacement thickness added on than...section and the flat plate airfoil with Kutta condition. The experimental results are for the NACA 64A010 airfoil at M = 0.5 and Reynolds number between...practice for actual airfoils. The experimental data shown in Fig. 3.5 are for the NACA 4 and 5 digit series airfoils (Ref. 17). The lift curve slope is

  3. Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

    NASA Technical Reports Server (NTRS)

    Noonan, K. W.; Bingham, G. J.

    1977-01-01

    An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

  4. Effect of Ground Proximity on the Aerodynamic Characteristics of Aspect-Ratio-1 Airfoils With and Without End Plates

    NASA Technical Reports Server (NTRS)

    Carter, Arthur W.

    1961-01-01

    An investigation has been made to determine the effect of ground proximity on the aerodynamic characteristics of aspect-ratio-1 airfoils. The investigation was made with the model moving over the water in a towing tank in order to eliminate the effects of wind-tunnel walls and of boundary layer on ground boards at small ground clearances. The results indicated that, as the ground was approached, the airfoils experienced an increase in lift-curve slope and a reduction in induced drag; thus, lift-drag ratio was increased. As the ground was approached, the profile drag remained essentially constant for each airfoil. Near the ground, the addition of end plates to the airfoil resulted in a large increase in lift-drag ratio. The lift characteristics of the airfoils indicated stability of height at positive angles of attack and instability of height at negative angles; therefore, the operating range of angles of attack would be limited to positive values. At positive angles of attack, the static longitudinal stability was increased as the height above the ground was reduced. Comparison of the experimental data with Wieselsberger's ground-effect theory (NACA Technical Memorandum 77) indicated generally good agreement between experiment and theory for the airfoils without end plates.

  5. Airfoil Dynamic Stall and Rotorcraft Maneuverability

    NASA Technical Reports Server (NTRS)

    Bousman, William G.

    2000-01-01

    The loading of an airfoil during dynamic stall is examined in terms of the augmented lift and the associated penalties in pitching moment and drag. It is shown that once stall occurs and a leading-edge vortex is shed from the airfoil there is a unique relationship between the augmented lift, the negative pitching moment, and the increase in drag. This relationship, referred to here as the dynamic stall function, shows limited sensitivity to effects such as the airfoil section profile and Mach number, and appears to be independent of such parameters as Reynolds number, reduced frequency, and blade sweep. For single-element airfoils there is little that can be done to improve rotorcraft maneuverability except to provide good static C(l(max)) characteristics and the chord or blade number that is required to provide the necessary rotor thrust. However, multi-element airfoils or airfoils with variable geometry features can provide augmented lift in some cases that exceeds that available from a single-element airfoil. The dynamic stall function is shown to be a useful tool for the evaluation of both measured and calculated dynamic stall characteristics of single element, multi-element, and variable geometry airfoils.

  6. Aerodynamic characteristics of NACA 4412 airfoil sction with flap

    NASA Astrophysics Data System (ADS)

    Ockfen, Alex E.; Matveev, Konstantin I.

    2009-09-01

    Wing-in-Ground vehicles and aerodynamically assisted boats take advantage of increased lift and reduced drag of wing sections in the ground proximity. At relatively low speeds or heavy payloads of these craft, a flap at the wing trailing-ground-effect flow id numerically investigated in this study. The computational method consists of a steady-state, incompressible, finite volume method utilizing the Spalart-Allmaras turbulence model. Grid generation and solution of the Navier-Stokes equations are completed flow with a flap, as well as ground-effect motion without a flap. Aerodynamic forces are plain flap. Changes in the flow introduced with the flap addition are also discussed. Overall, the use of a flap on wings with small attack angles is found to be beneficial for small flap deflections up to 5% of the chord, where the contribution of lift augmentation exceeds the drag increase, yielding an augmented lift-to-drag ratio

  7. Numerical prediction of turbulent flow over airfoil sections with a new nonequilibrium turbulence model

    NASA Technical Reports Server (NTRS)

    Ahmed, S.; Tannehill, J. C.

    1990-01-01

    A new nonequilibrium turbulence closure model has been developed for computing wall bounded two-dimensional turbulent flows. This two-layer eddy viscosity model was motivated by the success of the Johnson-King model in separated flow regions. The influence of history effects are described by an ordinary differential equation developed from the turbulent kinetic energy equation. The performance of the present model has been evaluated by solving the flow around three airfoils using the Reynolds time-averaged Navier-Stokes equations. Excellent results were obtained for both attached and separated turbulent flows about the NACA 0012 airfoil, the RAE 2822 airfoil, and the Integrated Technology A 153W airfoil. Based on the comparison of the numerical solutions with the available experimental data, it is concluded that the new nonequilibrium turbulence model accurately captures the history effects of convection and diffusion on turbulence.

  8. Nonlinear power flow feedback control for improved stability and performance of airfoil sections

    DOEpatents

    Wilson, David G.; Robinett, III, Rush D.

    2013-09-03

    A computer-implemented method of determining the pitch stability of an airfoil system, comprising using a computer to numerically integrate a differential equation of motion that includes terms describing PID controller action. In one model, the differential equation characterizes the time-dependent response of the airfoil's pitch angle, .alpha.. The computer model calculates limit-cycles of the model, which represent the stability boundaries of the airfoil system. Once the stability boundary is known, feedback control can be implemented, by using, for example, a PID controller to control a feedback actuator. The method allows the PID controller gain constants, K.sub.I, K.sub.p, and K.sub.d, to be optimized. This permits operation closer to the stability boundaries, while preventing the physical apparatus from unintentionally crossing the stability boundaries. Operating closer to the stability boundaries permits greater power efficiencies to be extracted from the airfoil system.

  9. Aeroelastic dynamic response and control of an airfoil section with control surface nonlinearities

    NASA Astrophysics Data System (ADS)

    Li, Daochun; Guo, Shijun; Xiang, Jinwu

    2010-10-01

    Nonlinearities in aircraft mechanisms are inevitable, especially in the control system. It is necessary to investigate the effects of them on the dynamic response and control performance of aeroelastic system. In this paper, based on the state-dependent Riccati equation method, a state feedback suboptimal control law is derived for aeroelastic response and flutter suppression of a three degree-of-freedom typical airfoil section. With the control law designed, nonlinear effects of freeplay in the control surface and time delay between the control input and actuator are investigated by numerical approach. A cubic nonlinearity in pitch degree is adopted to prevent the aeroelastic responses from divergence when the flow velocity exceeds the critical flutter speed. For the system with a freeplay, the responses of both open- and closed-loop systems are determined with Runge-Kutta algorithm in conjunction with Henon's method. This method is used to locate the switching points accurately and efficiently as the system moves from one subdomain into another. The simulation results show that the freeplay leads to a forward phase response and a slight increase of flutter speed of the closed-loop system. The effect of freeplay on the aeroelastic response decreases as the flow velocity increases. The time delay between the control input and actuator may impair control performance and cause high-frequency motion and quasi-periodic vibration.

  10. The effect of wall interference upon the aerodynamic characteristics of an airfoil spanning a closed-throat circular wind tunnel

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G; Graham, Donald J

    1946-01-01

    The results of a theoretical and experimental investigation of wall interference for an airfoil spanning a closed-throat circular wind tunnel are presented. Analytical equations are derived which relate the characteristics of an airfoil in the tunnel at subsonic speeds with the characteristics in free air. The analysis takes into consideration the effect of fluid compressibility and is based upon the assumption that the chord of the airfoil is small as compared with the diameter of the tunnel. The development is restricted to an untwisted, constant-chord airfoil spanning the middle of the tunnel. Brief theoretical consideration is also given to the problem of choking at high speeds. Results are then presented of tests to determine the low-speed characteristics of an NACA 4412 airfoil for two chord-diameter ratios. While, on the basis of these experiments, no appraisal is possible of the accuracy of the corrections at high speeds, the data indicate that at low Mach numbers the analytical results are valid, even for relatively large values of the chord-diameter ratio.

  11. Unsteady Airloads on Airfoils in Reverse Flow

    NASA Astrophysics Data System (ADS)

    Lind, Andrew; Jones, Anya

    2014-11-01

    This work gives insight into the influence of airfoil characteristics on unsteady airloads for rotor applications where local airfoil sections may operate at high and/or reverse flow angles of attack. Two-dimensional wind tunnel experiments have been performed on four airfoil sections to investigate the effects of thickness, camber, and trailing edge shape on unsteady airloads (lift, pressure drag, and pitching moment). These model rotor blades were tested through 360 deg of incidence for 104 <=Re <=106 . Unsteady pressure transducers were mounted on the airfoil surface to measure the high frequency, dynamic pressure variations. The temporal evolution of chordwise pressure distributions and resulting airloads is quantified for each airfoil in each of the three unsteady wake regimes present in reverse flow. Specifically, the influence of the formation, growth, and shedding of vortices on the surface pressure distribution is quantified and compared between airfoils with a sharp geometric trailing edge and those with a blunt geometric trailing edge. These findings are integral to mitigation of rotor blade vibrations for applications where airfoil sections are subjected to reverse flow, such as high-speed helicopters and tidal turbines.

  12. Evaluation of CFD to Determine Two-Dimensional Airfoil Characteristics for Rotorcraft Applications

    NASA Technical Reports Server (NTRS)

    Smith, Marilyn J.; Wong, Tin-Chee; Potsdam, Mark; Baeder, James; Phanse, Sujeet

    2004-01-01

    The efficient prediction of helicopter rotor performance, vibratory loads, and aeroelastic properties still relies heavily on the use of comprehensive analysis codes by the rotorcraft industry. These comprehensive codes utilize look-up tables to provide two-dimensional aerodynamic characteristics. Typically these tables are comprised of a combination of wind tunnel data, empirical data and numerical analyses. The potential to rely more heavily on numerical computations based on Computational Fluid Dynamics (CFD) simulations has become more of a reality with the advent of faster computers and more sophisticated physical models. The ability of five different CFD codes applied independently to predict the lift, drag and pitching moments of rotor airfoils is examined for the SC1095 airfoil, which is utilized in the UH-60A main rotor. Extensive comparisons with the results of ten wind tunnel tests are performed. These CFD computations are found to be as good as experimental data in predicting many of the aerodynamic performance characteristics. Four turbulence models were examined (Baldwin-Lomax, Spalart-Allmaras, Menter SST, and k-omega).

  13. Low-speed aerodynamic characteristics of a 13 percent thick medium speed airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1979-01-01

    Wind tunnel tests were conducted to determine the low speed, two dimensional aerodynamic characteristics of a 13percent thick medium speed airfoil designed for general aviation applications. The results were compared with data for the 13 percent thick low speed airfoil. The tests were conducted over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle of attack frange from about -8 deg to 10 deg. The objective of retaining good high-lift low speed characteristics for an airfoil designed to have good medium speed cruise performance was achieved.

  14. Experimental and Analytical Investigation of the Coolant Flow Characteristics in Cooled Turbine Airfoils

    NASA Technical Reports Server (NTRS)

    Damerow, W. P.; Murtaugh, J. P.; Burggraf, F.

    1972-01-01

    The flow characteristics of turbine airfoil cooling system components were experimentally investigated. Flow models representative of leading edge impingement, impingement with crossflow (midchord cooling), pin fins, feeder supply tube, and a composite model of a complete airfoil flow system were tested. Test conditions were set by varying pressure level to cover the Mach number and Reynolds number range of interest in advanced turbine applications. Selected geometrical variations were studied on each component model to determine these effects. Results of these tests were correlated and compared with data available in the literature. Orifice flow was correlated in terms of discharge coefficients. For the leading edge model this was found to be a weak function of hole Mach number and orifice-to-impinged wall spacing. In the impingement with crossflow tests, the discharge coefficient was found to be constant and thus independent of orifice Mach number, Reynolds number, crossflow rate, and impingement geometry. Crossflow channel pressure drop showed reasonable agreement with a simple one-dimensional momentum balance. Feeder tube orifice discharge coefficients correlated as a function of orifice Mach number and the ratio of the orifice-to-approach velocity heads. Pin fin data was correlated in terms of equivalent friction factor, which was found to be a function of Reynolds number and pin spacing but independent of pin height in the range tested.

  15. Numerical Simulations of the Steady and Unsteady Aerodynamic Characteristics of a Circulation Control Wing Airfoil

    NASA Technical Reports Server (NTRS)

    Liu, Yi; Sankar, Lakshmi N.; Englar, Robert J.; Ahuja, Krishan K.

    2003-01-01

    The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.

  16. Experimental Study of Tip Vortex Flow from a Periodically Pitched Airfoil Section

    NASA Technical Reports Server (NTRS)

    Zaman, KBMQ; Fagan, A. F.; Mankbadi, M. R.

    2016-01-01

    An experimental investigation of a tip vortex from a NACA0012 airfoil is conducted in a low-speed wind tunnel at a chord Reynolds number of 4x10(exp 4). Initially, data for a stationary airfoil held at various angles-of-attack (alpha) are gathered. Detailed surveys are done for two cases: alpha=10 deg with attached flow and alpha=25 deg with massive flow separation on the upper surface. Distributions of various properties are obtained using hot-wire anemometry. Data include mean velocity, streamwise vorticity and turbulent stresses at various streamwise locations. For all cases, the vortex core is seen to involve a mean velocity deficit. The deficit apparently traces to the airfoil wake, part of which gets wrapped by the tip vortex. At small alpha, the vortex is laminar within the measurement domain. The strength of the vortex increases with increasing alpha but undergoes a sudden drop around alpha (is) greater than 16 deg. The drop in peak vorticity level is accompanied by transition and a sharp rise in turbulence within the core. Data are also acquired with the airfoil pitched sinusoidally. All oscillation cases pertain to a mean alpha=15 deg while the amplitude and frequency are varied. An example of phase-averaged data for an amplitude of +/-10 deg and a reduced frequency of k=0.2 is discussed. All results are compared with available data from the literature shedding further light on the complex dynamics of the tip vortex.

  17. An investigation of the aerodynamic characteristics of a new general aviation airfoil in flight

    NASA Technical Reports Server (NTRS)

    Gregorek, G. M.; Hoffmann, M. J.; Weislogel, G. S.

    1982-01-01

    A low speed airfoil, the GA(W)-2, - a 13% thickness to chord ratio airfoil was evaluated. The wing of a Beech Sundowner was modified at by adding balsa ribs and covered with aluminum skin, to alter the existing airfoil shape to that of the GA(W)-2 airfoil. The aircraft was flown in a flight test program that gathered wing surface pressures and wake data from which the lift drag, and pitching moment of the airfoil could be determined. After the base line performance of the airfoil was measured, the drag due to surface irregularities such as steps, rivets and surface waviness was determined. The potential reduction of drag through the use of surface coatings such as KAPTON was also investigated.

  18. The Effects of Blowing Over Various Trailing-edge Flaps on an NACA 0006 Airfoil Section, Comparisons with Various Types of Flaps on other Airfoil Sections, and an Analysis of Flow and Power Relationships for Blowing Systems

    NASA Technical Reports Server (NTRS)

    Dods, J. B., Jr.; Watson, E. C.

    1976-01-01

    The results are presented of a two-dimensional investigation conducted to determine the effect of blowing over various types of trailing-edge flaps on a wing having the NACA 0006 airfoil section and a drooped-nose flap. The position and profile of the trailing-edge flap, the nozzle height, and the location of the flap with respect to the nozzle were found to be important variables. Data from many investigations were used to make an evaluation of the effects of blowing on lift. An analysis was made of flow and power relationships for blowing systems.

  19. An Experimental Investigation of the Acoustic and Fluid Dynamic Characteristics of a Circulation-Controlled Airfoil

    DTIC Science & Technology

    2011-05-13

    sound production from a hydrofoil and identified three mechanisms: (1) low frequency curvature noise associated with interaction of a turbulent...2002). 2 Technical Approach A two-dimensional, dual-slotted, elliptic circulation control airfoil based on the hydrofoil studied by Rogers...airfoil, shown in Figure 1A, is designed based on the geometry of the hydrofoil previously studied by Rogers & Donnelly (2004). The airfoil’s profile

  20. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

    Abbott, Ira H; Von Doenhoff, Albert E; Stivers, Louis, Jr

    1945-01-01

    The historical development of NACA airfoils is briefly reviewed. New data are presented that permit the rapid calculation of the approximate pressure distributions for the older NACA four-digit and five-digit airfoils by the same methods used for the NACA 6-series airfoils. The general methods used to derive the basic thickness forms for NACA 6 and 7-series airfoils together with their corresponding pressure distributions are presented. Detail data necessary for the application of the airfoils to wing design are presented in supplementary figures placed at the end of the paper. The report includes an analysis of the lift, drag, pitching-moment, and critical-speed characteristics of the airfoils, together with a discussion of the effects of surface conditions. Available data on high-lift devices are presented. Problems associated with lateral-control devices, leading-edge air intakes, and interference are briefly discussed, together with aerodynamic problems of application. (author)

  1. Comparison of Full-Scale Propellers Having R.A.F.-6 and Clark Y Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Freeman, Hugh B

    1932-01-01

    In this report the efficiencies of two series of propellers having two types of blade sections are compared. Six full-scale propellers were used, three having R. A. F.-6 and three Clark Y airfoil sections with thickness/chord ratios of 0.06, 0.08, and 0.10. The propellers were tested at five pitch setting, which covered the range ordinarily used in practice. The propellers having the Clark Y sections gave the highest peak efficiency at the low pitch settings. At the high pitch settings, the propellers with R. A. F.-6 sections gave about the same maximum efficiency as the Clark Y propellers and were more efficient for the conditions of climb and take-off.

  2. Comparison of the experimental aerodynamic characteristics of theoretically and experimentally designed supercritical airfoils

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1974-01-01

    A lifting airfoil theoretically designed for shockless supercritical flow utilizing a complex hodograph method has been evaluated in the Langley 8-foot transonic pressure tunnel at design and off-design conditions. The experimental results are presented and compared with those of an experimentally designed supercritical airfoil which were obtained in the same tunnel.

  3. Aerodynamic Characteristics of a 14-Percent-Thick NASA Supercritical Airfoil Designed for a Normal-Force Coefficient of 0.7

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1975-01-01

    This report documents the experimental aerodynamic characteristics of a 14 percent thick supercritical airfoil based on an off design sonic pressure plateau criterion. The design normal force coefficient was 0.7. The results are compared with those of the family related 10 percent thick supercritical airfoil 33. Comparisons are also made between experimental and theoretical characteristics and composite drag rise characteristics derived for a full scale Reynolds number of 40 million.

  4. Wind tunnel tests of two airfoils for wind turbines operating at high reynolds numbers

    SciTech Connect

    Sommers, D.; Tangler, J.

    2000-06-29

    The objectives of this study were to verify the predictions of the Eppler Airfoil Design and Analysis Code for Reynolds numbers up to 6 x 106 and to acquire the section characteristics of two airfoils being considered for large, megawatt-size wind turbines. One airfoil, the S825, was designed to achieve a high maximum lift coefficient suitable for variable-speed machines. The other airfoil, the S827, was designed to achieve a low maximum lift coefficient suitable for stall-regulated machines. Both airfoils were tested in the NASA Langley Low-Turbulence Pressure Tunnel (LTPT) for smooth, fixed-transition, and rough surface conditions at Reynolds numbers of 1, 2, 3, 4, and 6 x 106. The results show the maximum lift coefficient of both airfoils is substantially underpredicted for Reynolds numbers over 3 x 106 and emphasized the difficulty of designing low-lift airfoils for high Reynolds numbers.

  5. The Effectiveness at High Speeds of a 20-Percent-chord Plain Trailing-edge Flap on the NACA 65-210 Airfoil Section

    NASA Technical Reports Server (NTRS)

    Stivers, Louis S., Jr.

    1947-01-01

    An analysis has been made of the lift-control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flag on the INCA 65-210 airfoil section.

  6. Tests of related forward-camber airfoils in the variable-density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M; Greenberg, Harry

    1937-01-01

    A recent investigation of numerous related airfoils indicated that positions of camber forward of the usual location resulted in an increase of the maximum lift. As an extension of this investigation, a series of forward-camber airfoils has been developed, the members of which show airfoil characteristics superior to those of the airfoils previously investigated. The primary object of this report is to present fully corrected results for airfoils in the useful range of shapes. With the data thus made available, an airplane designer may intelligently choose the best possible airfoil-section shape for a given application and may predict to a reasonable degree the aerodynamic characteristics to be expected in flight from the section shape chosen.

  7. Two-dimensional aerodynamic characteristics of three rotorcraft airfoils at Mach numbers from 0.35 to 0.90

    NASA Technical Reports Server (NTRS)

    Bingham, G. J.; Noonan, K. W.

    1982-01-01

    Three airfoils designed for helicopter rotor application were investigated in the Langley 6- by 28-inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics at Mach numbers from 0.34 to 0.88 and respective Reynolds numbers from about 4.4 x 10(6) power to 9.5 x 10(6) power. The airfoils have thickness-to-chord ratios of 0.08, 0.10, and 0.12. Trailing-edge reflex was applied to minimize pitching moment. The maximum normal-force coefficient of the RC(3)-12 airfoil is from 0.1 to 0.2 higher, depending on Mach number M, than that of the NACA 0012 airfoil tested in the same facility. The maximum normal-force coefficient of the RC(3)-10 is about equal to that of the NACA 0012 at Mach numbers to 0.40 and is higher than that of the NACA 0012 at Mach numbers above 0.40. The maximum normal force coefficient of the RC(3)-08 is about 0.19 lower than that of the NACA 0012 at a Mach number of 0.35 and about 0.05 lower at a Mach number of 0.54. The drag divergence Mach number of the RC(3)-08 airfoil at normal-force coefficients below 0.1 was indicated to be greater than the maximum test Mach number of 0.88. At zero lift, the drag-divergence Mach numbers of the RC(3)-12 and the RC(3)-10 are about 0.77 and 0.82, respectively.

  8. Effect of pivot location and passive heave on propulsion from a pitching airfoil

    NASA Astrophysics Data System (ADS)

    Mackowski, A. W.; Williamson, C. H. K.

    2017-01-01

    We experimentally investigate the propulsive characteristics of a pitching NACA 0012 airfoil section, with emphasis on thrust and propulsive efficiency, at a Reynolds number of 1.7 ×104 . For the sake of mechanical simplicity, we consider an airfoil restricted to a single actuator in the pitching direction. We examine the effect of changing the airfoil's axis of rotation, finding that contrary to Garrick's linear theory, there exists a pitching axis near the airfoil that maximizes propulsive efficiency. Next, we examine the effect of placing passive springs on the airfoil in the heave (transverse) direction using our Cyber-Physical Fluid Dynamics technique. This elastic heaving motion allows the airfoil to combine pitching and heaving modes while being actuated only in the pitching direction. Two sets of dynamics are considered: one case where the airfoil is weighted unevenly and pitched about its center of mass (so that the resulting heaving motion is independent of inertial forces), and another case where the airfoil's center of mass is fixed at its centroid. For pitching at an amplitude of 8∘ and a reduced frequency k of two, we find that elastic heave produces a maximum propulsive efficiency of 35%, compared to 25% without any heave motion. Further, while operating at the same efficiency as the static-pivot case, we find that passive heaving greatly increases the magnitude of the airfoil's thrust. The airfoil configurations with highest propulsive efficiency generally involve pitching near or ahead of the airfoil's leading edge.

  9. Low-speed aerodynamic characteristics of a transport configuration having a 42 deg swept supercritical airfoil wing and three tail height positions

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Sleeman, W. C., Jr.

    1974-01-01

    A low speed investigation was conducted in the Langley V/STOL tunnel to define the static stability characteristics of an advanced high subsonic speed transport aircraft model in the cruise configuration (no high lift system). The wing of the model had 42 deg sweep of the quarter chord line, an aspect ratio of 6.78, and supercritical airfoil sections. Three different horizontal tail configurations (high, mid, and low) were investigated on the complete model and for the model with the wing removed in order to assess effects of the wing flow field on the tail contributions to both longitudinal and lateral stability characteristics. All the model configurations investigated were tested over an angle of attack range from approximately -5 to 23 deg. Some model configurations were also tested over an angle of attack range from about 11 to 38 deg in order to explore the aerodynamic characteristics in the deep stall region.

  10. An Empirical Method Permitting Rapid Determination of the Area, Rate and Distribution of Water-Drop Impingement on an Airfoil of Arbitrary Section at Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Bergrun, N. R.

    1951-01-01

    An empirical method for the determination of the area, rate, and distribution of water-drop impingement on airfoils of arbitrary section is presented. The procedure represents an initial step toward the development of a method which is generally applicable in the design of thermal ice-prevention equipment for airplane wing and tail surfaces. Results given by the proposed empirical method are expected to be sufficiently accurate for the purpose of heated-wing design, and can be obtained from a few numerical computations once the velocity distribution over the airfoil has been determined. The empirical method presented for incompressible flow is based on results of extensive water-drop. trajectory computations for five airfoil cases which consisted of 15-percent-thick airfoils encompassing a moderate lift-coefficient range. The differential equations pertaining to the paths of the drops were solved by a differential analyzer. The method developed for incompressible flow is extended to the calculation of area and rate of impingement on straight wings in subsonic compressible flow to indicate the probable effects of compressibility for airfoils at low subsonic Mach numbers.

  11. Wind-tunnel investigation of NACA 23012, 23021, and 23030 airfoils equipped with 40-percent-chord double slotted flaps

    NASA Technical Reports Server (NTRS)

    Harris, Thomas A; Recant, Isidore G

    1941-01-01

    Report presents the results of an investigation conducted in the NACA 7 by 10-foot win tunnel to determine the effect of the deflection of main and auxiliary slotted flaps on the aerodynamic section characteristics of large-chord NACA 23012, 23021, 23030 airfoils equipped with 40-percent-chord double slotted flaps. The complete aerodynamic section characteristics and envelope polar curves are given for each airfoil-flap combination. The effect of airfoil thickness is shown, and comparisons are made of single slotted flaps with double slotted flaps on each of the airfoils.

  12. Two-Dimensional Wind-Tunnel Investigation of Modified NACA 65(sub 112)-111 Airfoil with 35-Percent-Chord Slotted Flap to Determine Pitching-Moment Characteristics and Effects of Roughness

    NASA Technical Reports Server (NTRS)

    Racisz, Stanley F.

    1947-01-01

    An investigation has been made in the Langley two-dimensional low-turbulence pressure tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65(sub 1120)-111 airfoil section modified by removing the trailing-edge cusp. The section pitching-moment characteristics and the effects of standard roughness on the section characteristics were determined for the flap retracted at Reynolds numbers ranging from 3.0 x 10(exp 6) to 9.0 x 10(exp 6).

  13. Computational design and analysis of flatback airfoil wind tunnel experiment.

    SciTech Connect

    Mayda, Edward A.; van Dam, C.P.; Chao, David D.; Berg, Dale E.

    2008-03-01

    A computational fluid dynamics study of thick wind turbine section shapes in the test section of the UC Davis wind tunnel at a chord Reynolds number of one million is presented. The goals of this study are to validate standard wind tunnel wall corrections for high solid blockage conditions and to reaffirm the favorable effect of a blunt trailing edge or flatback on the performance characteristics of a representative thick airfoil shape prior to building the wind tunnel models and conducting the experiment. The numerical simulations prove the standard wind tunnel corrections to be largely valid for the proposed test of 40% maximum thickness to chord ratio airfoils at a solid blockage ratio of 10%. Comparison of the computed lift characteristics of a sharp trailing edge baseline airfoil and derived flatback airfoils reaffirms the earlier observed trend of reduced sensitivity to surface contamination with increasing trailing edge thickness.

  14. Potential flow analysis of glaze ice accretions on an airfoil

    NASA Technical Reports Server (NTRS)

    Zaguli, R. J.

    1984-01-01

    The results of an analytical/experimental study of the flow fields about an airfoil with leading edge glaze ice accretion shapes are presented. Tests were conducted in the Icing Research Tunnel to measure surface pressure distributions and boundary layer separation reattachment characteristics on a general aviation wing section to which was affixed wooden ice shapes which approximated typical glaze ice accretions. Comparisons were made with predicted pressure distributions using current airfoil analysis codes as well as the Bristow mixed analysis/design airfoil panel code. The Bristow code was also used to predict the separation reattachment dividing streamline by inputting the appropriate experimental surface pressure distribution.

  15. Airfoil structure

    DOEpatents

    Frey, G.A.; Twardochleb, C.Z.

    1998-01-13

    Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally ``C`` configuration of the airfoil. The generally ``C`` configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion. 6 figs.

  16. Airfoil structure

    DOEpatents

    Frey, Gary A.; Twardochleb, Christopher Z.

    1998-01-01

    Past airfoil configurations have been used to improve aerodynamic performance and engine efficiencies. The present airfoil configuration further increases component life and reduces maintenance by reducing internal stress within the airfoil itself. The airfoil includes a chord and a span. Each of the chord and the span has a bow being summed to form a generally "C" configuration of the airfoil. The generally "C" configuration includes a compound bow in which internal stresses resulting from a thermal temperature gradient are reduced. The structural configuration reduces internal stresses resulting from thermal expansion.

  17. Natural laminar flow airfoil design considerations for winglets on low-speed airplanes

    NASA Technical Reports Server (NTRS)

    Vandam, C. P.

    1984-01-01

    Winglet airfoil section characteristics which significantly influence cruise performance and handling qualities of an airplane are discussed. A good winglet design requires an airfoil section with a low cruise drag coefficient, a high maximum lift coefficient, and a gradual and steady movement of the boundary layer transition location with angle of attack. The first design requirement provides a low crossover lift coefficient of airplane drag polars with winglets off and on. The other requirements prevent nonlinear changes in airplane lateral/directional stability and control characteristics. These requirements are considered in the design of a natural laminar flow airfoil section for winglet applications and chord Reynolds number of 1 to 4 million.

  18. Analysis of a theoretically optimized transonic airfoil

    NASA Technical Reports Server (NTRS)

    Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.

    1978-01-01

    Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.

  19. Study of a new airfoil used in reversible axial fans

    NASA Technical Reports Server (NTRS)

    Li, Chaojun; Wei, Baosuo; Gu, Chuangang

    1991-01-01

    The characteristics of the reverse ventilation of axial flow are analyzed. An s shaped airfoil with a double circular arc was tested in a wind tunnel. The experimental results showed that the characteristics of this new airfoil in reverse ventilation are the same as those in normal ventilation, and that this airfoil is better than the existing airfoils used on reversible axial fans.

  20. Aerodynamic characteristics of airfoils VI : continuation of reports nos. 93, 124, 182, 244, and 286

    NASA Technical Reports Server (NTRS)

    1930-01-01

    This collection of data on airfoils has been made from the published reports of a number of the leading aerodynamic laboratories of this country and Europe. The information which was originally expressed according to the different customs of the several laboratories is here presented in a uniform series of charts and tables suitable for use of designing engineers and for purposes of general reference. The authority for the results here presented is given as the name of the laboratory at which the experiments were conducted, with the size of the model, wind velocity, and year of test.

  1. Comparison of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an analysis code

    NASA Technical Reports Server (NTRS)

    Harris, C. D.; Allison, D. O.

    1977-01-01

    Comparisons of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an advanced analysis code. Comparisons are made at near-design normal-force coefficients for Reynolds numbers from 2 to 11 million. Comments are made concerning various input parameters to the code.

  2. Additional Testing of the DHC-6 Twin Otter Tailplane Iced Airfoil Section in the Ohio State University 7x10 Low Speed Wind Tunnel. Volume 2

    NASA Technical Reports Server (NTRS)

    Gregorek, Gerald; Dresse, John J.; LaNoe, Karine; Ratvasky, Thomas (Technical Monitor)

    2000-01-01

    The need for fundamental research in Ice Contaminated Tailplane Stall (ICTS) was established through three international conferences sponsored by the FAA. A joint NASA/FAA Tailplane Icing Program was formed in 1994 with the Ohio State University playing a critical role for wind tunnel and analytical research. Two entries of a full-scale 2-dimensional tailplane airfoil model of a DHC-6 Twin Otter were made in The Ohio State University 7x10 ft wind tunnel. This report describes the second test entry that examined additional ice shapes and roughness, as well as airfoil section differences. The addition data obtained in this test fortified the original database of aerodynamic coefficients that permit a detailed analysis of flight test results with an OSU-developed analytical program. The testing encompassed a full range of angles of attack and elevator deflections at flight Reynolds number conditions. Aerodynamic coefficients, C(L), C(M), and C(He), were obtained by integrating static pressure coefficient, C(P), values obtained from surface taps. Comparisons of clean and iced airfoil results show a significant decrease in the tailplane aeroperformance (decreased C(Lmax), decreased stall angle, increased C(He)) for all ice shapes with the grit having the lease affect and the LEWICE shape having the greatest affect. All results were consistent with observed tailplane stall phenomena and constitute an effective set of data for comprehensive analysis of ICTS.

  3. Tests If a Highly Cambered Low-Drag-Airfoil Section with a Lift-Control Flap, Special Report

    NASA Technical Reports Server (NTRS)

    Abbott, Ira H.; Miller, Ralph B.

    1942-01-01

    Tests were made in the NACA two-dimensional low turbulence pressure tunnel of a highly cambered low-drag airfoil (NACA 65,3-618) with a plain flap designed for lift control. The results indicate that such a combination offers attractive possibilities for obtaining low profile-drag coefficients over a wide range of lift coefficients without large reductions of critical speed.

  4. A computer program for estimating the aerodynamic characteristics of NACA 16-series airfoils

    NASA Technical Reports Server (NTRS)

    Maksymiuk, C. M.; Watson, S. A.

    1983-01-01

    A computer program written in a table ""look-up'' format, is presented which provides a comprehensive data base on NACA 16-series airfoils. The geometry covered is limited to cambers for a design-lift coefficient from 0.0 to 0.7 and thickness ratios from 4 to 21%. The data include Mach numbers from 0.3 to 1.6, angles of attack from -4 to 8 degrees, and lift coefficients from 0.0 to 0.8. Extrapolation is used to obtain data from Mach numbers, angles of attack, and lift coefficients beyond those for which data are available. A routine to adjust the lift and drag coefficients beyond stall is included. The uses and limitations of the program are also discussed.

  5. Impact of pulsed blowing jet on aerodynamic characteristics of wind turbine airfoils

    NASA Astrophysics Data System (ADS)

    Bobonea, Andreea

    2012-11-01

    Wind turbine growth in size and weight made it impossible to control turbines passively as they were controlled in the past. Current efforts focus on increasing their aerodynamic efficiency and operational range through active flow control methods. One of the main methods of active flow control is the usage of blowing devices with constant or pulsed jets. By adding stored high-momentum air through slots into the boundary layer, they overcome adverse pressure gradients and postpone separation. Pulsed blowing sends short pulses rather than a continuous jet of fluid into the boundary layer and has been found to be more effective. Through CFD simulations over a 2D wind turbine airfoil, this research highlights the impact of different slot geometries with constant/pulsed blowing, on the effectiveness of this active flow control technique.

  6. An improved viscid/inviscid interaction procedure for transonic flow over airfoils

    NASA Technical Reports Server (NTRS)

    Melnik, R. E.; Chow, R. R.; Mead, H. R.; Jameson, A.

    1985-01-01

    A new interacting boundary layer approach for computing the viscous transonic flow over airfoils is described. The theory includes a complete treatment of viscous interaction effects induced by the wake and accounts for normal pressure gradient effects across the boundary layer near trailing edges. The method is based on systematic expansions of the full Reynolds equation of turbulent flow in the limit of Reynolds numbers, Reynolds infinity. Procedures are developed for incorporating the local trailing edge solution into the numerical solution of the coupled full potential and integral boundary layer equations. Although the theory is strictly applicable to airfoils with cusped or nearly cusped trailing edges and to turbulent boundary layers that remain fully attached to the airfoil surface, the method was successfully applied to more general airfoils and to flows with small separation zones. Comparisons of theoretical solutions with wind tunnel data indicate the present method can accurately predict the section characteristics of airfoils including the absolute levels of drag.

  7. Transonic airfoil and axial flow rotary machine

    DOEpatents

    Nagai, Naonori; Iwatani, Junji

    2015-09-01

    Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.

  8. Transonic airfoil codes

    NASA Technical Reports Server (NTRS)

    Garabedian, P. R.

    1979-01-01

    Computer codes for the design and analysis of transonic airfoils are considered. The design code relies on the method of complex characteristics in the hodograph plane to construct shockless airfoil. The analysis code uses artificial viscosity to calculate flows with weak shock waves at off-design conditions. Comparisons with experiments show that an excellent simulation of two dimensional wind tunnel tests is obtained. The codes have been widely adopted by the aircraft industry as a tool for the development of supercritical wing technology.

  9. Effect of trailing edge shape on the separated flow characteristics around an airfoil at low Reynolds number: A numerical study

    NASA Astrophysics Data System (ADS)

    Thomareis, Nikitas; Papadakis, George

    2017-01-01

    Direct numerical simulations of the flow field around a NACA 0012 airfoil at Reynolds number 50 000 and angle of attack 5° with 3 different trailing edge shapes (straight, blunt, and serrated) have been performed. Both time-averaged flow characteristics and the most dominant flow structures and their frequencies are investigated using the dynamic mode decomposition method. It is shown that for the straight trailing edge airfoil, this method can capture the fundamental as well as the subharmonic of the Kelvin-Helmholtz instability that develops naturally in the separating shear layer. The fundamental frequency matches well with relevant data in the literature. The blunt trailing edge results in periodic vortex shedding, with frequency close to the subharmonic of the natural shear layer frequency. The shedding, resulting from a global instability, has an upstream effect and forces the separating shear layer. Due to forcing, the shear layer frequency locks onto the shedding frequency while the natural frequency (and its subharmonic) is suppressed. The presence of serrations in the trailing edge creates a spanwise pressure gradient, which is responsible for the development of a secondary flow pattern in the spanwise direction. This pattern affects the mean flow in the near wake. It can explain an unexpected observation, namely, that the velocity deficit downstream of a trough is smaller than the deficit after a protrusion. Furthermore, the insertion of serrations attenuates the energy of vortex shedding by de-correlating the spanwise coherence of the vortices. This results in weaker forcing of the separating shear layer, and both the subharmonics of the natural frequency and the shedding frequency appear in the spectra.

  10. DHC-6 Twin Otter Tailplane Airfoil Section Testing in the Ohio State University 7x10 Wind Tunnel. Volume 1

    NASA Technical Reports Server (NTRS)

    Hiltner, Dale; McKee, Michael; LaNoe, Karine; Gregorek, Gerald; Ratvasky, Thomas (Technical Monitor)

    2000-01-01

    Ice contaminated tailplane stall (ICTS) has been found to be responsible for 16 accidents with 139 fatalities over the last three decades, and is suspected to have played a role in other accidents and incidents. The need for fundamental research in this area has been recognized at three international conferences sponsored by the FAA since 1991. In order to conduct such research, a joint NASA/FAA Tailplane Icing Program was formed in 1994: the Ohio State University has played an important role in this effort. The program employs icing tunnel testing, dry wind tunnel testing, flight testing, and analysis using a six-degrees-of-freedom computer code tailored to this problem. A central goal is to quantify the effect of tailplane icing on aircraft stability and control to aid in the analysis of flight test procedures to identify aircraft susceptibility to ICTS. This report contains the results ot testing of a full scale 2D model of a tailplane section of NASA's Icing Research Aircraft, with and without ice shapes, in an Ohio State University 7 x 10 Low Speed wind tunnel in 1994. The results have been integrated into a comprehensive database of aerodynamic coefficients and stability and control derivatives that will permit detailed analysis of flight test results with the analytical computer program. The testing encompassed a full range of angles of attack and elevator deflections, as well as two velocities to evaluate Reynolds number effects. Lift, drag, pitching moment, and hinge moment coefficients were obtained. In addition. instrumentation for use during flight testing was verified to be effective, all components showing acceptable fidelity. Comparison of clean and iced airfoil results show the ice shapes causing a significant decrease in the magnitude of CLmax (from -1.3 to -0.64) and associated stall angle (from -18.6 deg to -8.2 deg). Furthermore, the ice shapes caused an increase in hinge moment coefficient of approximately 0.02, the change being markedly abrupt

  11. Rotational Augmentation on a 2.3 MW Rotor Blade with Thick Flatback Airfoil Cross-Sections: Preprint

    SciTech Connect

    Schreck, S.; Fingersh, L.; Siegel, K.; Singh, M.; Medina, P.

    2013-01-01

    Rotational augmentation was analyzed for a 2.3 MW wind turbine, which was equipped with thick flatback airfoils at inboard radial locations and extensively instrumented for acquisition of time varying surface pressures. Mean aerodynamic force and surface pressure data were extracted from an extensive field test database, subject to stringent criteria for wind inflow and turbine operating conditions. Analyses of these data showed pronounced amplification of aerodynamic forces and significant enhancements to surface pressures in response to rotational influences, relative to two-dimensional, stationary conditions. Rotational augmentation occurrence and intensity in the current effort was found to be consistent with that observed in previous research. Notably, elevated airfoil thickness and flatback design did not impede rotational augmentation.

  12. Boundary Layer Control on Airfoils.

    ERIC Educational Resources Information Center

    Gerhab, George; Eastlake, Charles

    1991-01-01

    A phenomena, boundary layer control (BLC), produced when visualizing the fluidlike flow of air is described. The use of BLC in modifying aerodynamic characteristics of airfoils, race cars, and boats is discussed. (KR)

  13. Static-thrust Investigation of Full-scale PV-2 Helicopter Rotors Having NACA 0012.6 and 23012.6 Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Lipson, Stanley

    1946-01-01

    An investigation was conducted to compare the performance of two 25-ft-diam rotors which had identical dimensions and were similar in construction but different in blade airfoil-sections. Tests were conducted at indicated blade pitch angles from 3 degrees to 11.5 degrees and rotor speeds of 200, 290, and 371 rpm. The 23012.6 rotor required 2 percent less power to hover than the 0012.6. At thrust coefficients above design, the performance of the 23012.6 became better than the 0012.6 rotor.

  14. Computational Analysis of Dual Radius Circulation Control Airfoils

    NASA Technical Reports Server (NTRS)

    Lee-Rausch, E. M.; Vatsa, V. N.; Rumsey, C. L.

    2006-01-01

    The goal of the work is to use multiple codes and multiple configurations to provide an assessment of the capability of RANS solvers to predict circulation control dual radius airfoil performance and also to identify key issues associated with the computational predictions of these configurations that can result in discrepancies in the predicted solutions. Solutions were obtained for the Georgia Tech Research Institute (GTRI) dual radius circulation control airfoil and the General Aviation Circulation Control (GACC) dual radius airfoil. For the GTRI-DR airfoil, two-dimensional structured and unstructured grid computations predicted the experimental trend in sectional lift variation with blowing coefficient very well. Good code to code comparisons between the chordwise surface pressure coefficients and the solution streamtraces also indicated that the detailed flow characteristics were matched between the computations. For the GACC-DR airfoil, two-dimensional structured and unstructured grid computations predicted the sectional lift and chordwise pressure distributions accurately at the no blowing condition. However at a moderate blowing coefficient, although the code to code variation was small, the differences between the computations and experiment were significant. Computations were made to investigate the sensitivity of the sectional lift and pressure distributions to some of the experimental and computational parameters, but none of these could entirely account for the differences in the experimental and computational results. Thus, CFD may indeed be adequate as a prediction tool for dual radius CC flows, but limited and difficult to obtain two-dimensional experimental data prevents a confident assessment at this time.

  15. Thin oblique airfoils at supersonic speed

    NASA Technical Reports Server (NTRS)

    Jone, Robert T

    1946-01-01

    The well-known methods of thin-airfoil theory have been extended to oblique or sweptback airfoils of finite aspect ratio moving at supersonic speeds. The cases considered thus far are symmetrical airfoils at zero lift having plan forms bounded by straight lines. Because of the conical form of the elementary flow fields, the results are comparable in simplicity to the results of the two-dimensional thin-airfoil theory for subsonic speeds. In the case of untapered airfoils swept back behind the Mach cone the pressure distribution at the center section is similar to that given by the Ackeret theory for a straight airfoil. With increasing distance from the center section the distribution approaches the form given by the subsonic-flow theory. The pressure drag is concentrated chiefly at the center section and for long wings a slight negative drag may appear on outboard sections. (author)

  16. Low-speed aerodynamic characteristics of a 14-percent-thick NASA phase 2 supercritical airfoil designed for a lift coefficient of 0.7

    NASA Technical Reports Server (NTRS)

    Harris, C. D.; Mcghee, R. J.; Allison, D. O.

    1980-01-01

    The low speed aerodynamic characteristics of a 14 percent thick supercritical airfoil are documented. The wind tunnel test was conducted in the Low Turbulence Pressure Tunnel. The effects of varying chord Reynolds number from 2,000,000 to 18,000,000 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6,000,000 are included.

  17. Second Stage Turbine Bucket Airfoil.

    DOEpatents

    Xu, Liming; Ahmadi, Majid; Humanchuk, David John; Moretto, Nicholas; Delehanty, Richard Edward

    2003-05-06

    The second-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.

  18. Air/water two-phase flow test tunnel for airfoil studies

    NASA Astrophysics Data System (ADS)

    Ohashi, H.; Matsumoto, Y.; Ichikawa, Y.; Tsukiyama, T.

    1990-02-01

    A test tunnel for the study of airfoil performances under air/water two-phase flow condition has been designed and constructed. This facility will serve for a better understanding of the flow phenomena and characteristics of hydraulic machinery under gas/ liquid two-phase flow operating conditions. At the test section of the tunnel, a two-dimensional isolated airfoil or a cascade of airfoils is installed in a two-phase inlet flow with a uniform velocity (up to 10 m/s) and void fraction (up to 12%) distribution. The details of the tunnel structure and the measuring systems are described and the basic characteristics of the constructed tunnel are also given. As an example of the test results, void fraction distribution around a test airfoil is shown.

  19. Air/water two-phase flow test tunnel for airfoil studies

    NASA Astrophysics Data System (ADS)

    Ohashi, H.; Matsumoto, Y.; Ichikawa, Y.; Tsukiyama, T.

    1994-01-01

    A test tunnel for the study of airfoil performances under air/water two-phase flow condition has been designed and constructed. This facility will serve for a better understanding of the flow phenomena and characteristics of hydraulic machinery under gas/ liquid two-phase flow operating conditions. At the test section of the tunnel, a two-dimensional isolated airfoil or a cascade of airfoils is installed in a two-phase inlet flow with a uniform velocity (up to 10 m/s) and void fraction (up to 12%) distribution. The details of the tunnel structure and the measuring systems are described and the basic characteristics of the constructed tunnel are also given. As an example of the test results, void fraction distribution around a test airfoil is shown.

  20. Evaluation of a stalled airfoil analysis program

    NASA Technical Reports Server (NTRS)

    Rumsey, C. L.

    1985-01-01

    The Stalled Airfoil Analysis Program (SAAP) is a computer code for predicting the aerodynamic characteristics of an airfoil up to, and beyond, stall. SAAP is presently evaluated through comparisons with experiments and with two other theoretical methods over an extensive range of airfoils and Reynolds number conditions. SAAP modeled drag more accurately than either of the other methods, and at angles of attack below stall yielded a smoother lift variation with angle of attack.

  1. Airfoils for wind turbine

    DOEpatents

    Tangler, James L.; Somers, Dan M.

    1996-01-01

    Airfoils for the blade of a wind turbine wherein each airfoil is characterized by a thickness in a range from 16%-24% and a maximum lift coefficient designed to be largely insensitive to roughness effects. The airfoils include a family of airfoils for a blade 15 to 25 meters in length, a family of airfoils for a blade 1 to 5 meters in length, and a family of airfoils for a blade 5 to 10 meters in length.

  2. Airfoils for wind turbine

    DOEpatents

    Tangler, J.L.; Somers, D.M.

    1996-10-08

    Airfoils are disclosed for the blade of a wind turbine wherein each airfoil is characterized by a thickness in a range from 16%-24% and a maximum lift coefficient designed to be largely insensitive to roughness effects. The airfoils include a family of airfoils for a blade 15 to 25 meters in length, a family of airfoils for a blade 1 to 5 meters in length, and a family of airfoils for a blade 5 to 10 meters in length. 10 figs.

  3. Composite airfoil assembly

    DOEpatents

    Garcia-Crespo, Andres Jose

    2015-03-03

    A composite blade assembly for mounting on a turbine wheel includes a ceramic airfoil and an airfoil platform. The ceramic airfoil is formed with an airfoil portion, a blade shank portion and a blade dovetail tang. The metal platform includes a platform shank and a radially inner platform dovetail. The ceramic airfoil is captured within the metal platform, such that in use, the ceramic airfoil is held within the turbine wheel independent of the metal platform.

  4. The method of complex characteristics for design of transonic blade sections

    NASA Technical Reports Server (NTRS)

    Bledsoe, M. R.

    1986-01-01

    A variety of computational methods were developed to obtain shockless or near shockless flow past two-dimensional airfoils. The approach used was the method of complex characteristics, which determines smooth solutions to the transonic flow equations based on an input speed distribution. General results from fluid mechanics are presented. An account of the method of complex characteristics is given including a description of the particular spaces and coordinates, conformal transformations, and numerical procedures that are used. The operation of the computer program COMPRES is presented along with examples of blade sections designed with the code. A user manual is included with a glossary to provide additional information which may be helpful. The computer program in Fortran, including numerous comment cards is listed.

  5. Numerical study on reduction of aerodynamic noise around an airfoil with biomimetic structures

    NASA Astrophysics Data System (ADS)

    Wang, Jing; Zhang, Chengchun; Wu, Zhengyang; Wharton, James; Ren, Luquan

    2017-04-01

    A biomimetic airfoil featuring leading edge waves, trailing edge serrations and surface ridges is proposed in this study, based on flow control with each section meeting the NACA 0012 airfoil profile. Numerical simulations have been conducted to compare aerodynamic and acoustic performances between the NACA 0012 and biomimetic airfoils. These simulations utilize the large eddy simulation (LES) method and aeroacoustic analogy at an angle of attack of 0° and a Reynolds number of 1.0×105, based on using the airfoil chord as the characteristic length. The simulation results reveal the overall sound pressure levels (OASPLs) for all frequencies and at the seven observer points around the biomimetic airfoil, and a decrease of 13.1-13.9 dB is observed, whereas the drag coefficient is almost unchanged. The biomimetic structures can transform the shedding vortices in laminar mode for the NACA 0012 airfoil to regular horseshoe-type vortices in the wake, and reduce the spanwise correlation of the large-scale vortices, thereby restrain the vortex shedding noise around the biomimetic airfoil.

  6. Wind-Tunnel Investigation of the Effects of Profile Modification and Tabs on the Characteristics of Ailerons on a Low-Drag Airfoil

    NASA Technical Reports Server (NTRS)

    Crane, Robert M; Holtzclaw, Ralph W

    1944-01-01

    An investigation has been made to determine the effect of control-surface profile modifications on the aerodynamic characteristics of an NACA low-drag airfoil equipped with a 0.20-chord and a 0.15-chord aileron. Tab characteristics have been obtained for 0.20-aileron chord tabs on two of the 0.20-chord ailerons. Basic data are presented from which the effect of tabs can be calculated for specific cases. The data are sufficient for the solution of problems of fixed tabs with a differential linkage, as well as simple and spring-linked balancing tabs.

  7. Laminar-flow airfoil

    NASA Technical Reports Server (NTRS)

    Somers, Dan M. (Inventor)

    2005-01-01

    An airfoil having a fore airfoil element, an aft airfoil element, and a slot region in between them. These elements induce laminar flow over substantially all of the fore airfoil element and also provide for laminar flow in at least a portion of the slot region. The method of the invention is one for inducing natural laminar flow over an airfoil. In the method, a fore airfoil element, having a leading and trailing edge, and an aft airfoil element define a slot region. Natural laminar flow is induced over substantially all of the fore airfoil element, by inducing the pressures on both surfaces of the fore airfoil element to decrease to a location proximate the trailing edge of the fore airfoil element using pressures created by the aft airfoil element.

  8. Techniques for modifying airfoils and fairings on aircraft using foam and fiberglass

    NASA Technical Reports Server (NTRS)

    Meyer, M. B.; Jiran, F.

    1981-01-01

    The concept of using foam and fiberglass reinforced plastic to modify airfoils and fairings was applied successfully to high-speed aircraft at NASA Dryden Flight Research Center. An on-aircraft installation method was used to modify an F-15 wing glove and wing leading edge and an F-104 flap trailing edge in support of the Shuttle tile airload tests. A combination of methods, both an on-aircraft installation and an off-aircraft fabrication for installation on the aircraft, was used to modify a section of an F-111 supercritical wing with a natural laminar flow airfoil. Techniques, methods, problem areas, and recommendations are presented which indicate that using foam and fiberglass to modify airfoils and fairings on high-speed aircraft is a viable means of quickly developing airfoils and fairings with desired aerodynamic characteristics with little risk to the parent or carrier aircraft.

  9. Airfoil shape for flight at subsonic speeds

    DOEpatents

    Whitcomb, Richard T.

    1976-01-01

    An airfoil having an upper surface shaped to control flow accelerations and pressure distribution over the upper surface and to prevent separation of the boundary layer due to shock wave formulation at high subsonic speeds well above the critical Mach number. A highly cambered trailing edge section improves overall airfoil lifting efficiency.

  10. Wake structure and aerodynamic characteristics of an auto-propelled pitching airfoil

    NASA Astrophysics Data System (ADS)

    Hanchi, S.; Benkherouf, T.; Mekadem, M.; Oualli, H.; Keirsbulck, L.; Labraga, L.

    2013-05-01

    In the present study, we investigate the wake configuration as well as the flow aerodynamic and propulsive characteristics of a system equipped with a nature-inspired propulsion system. The study focuses on the effect of a set of pitching frequency and amplitude values on the flow behavior for a symmetric foil performing pitching sinusoidal rolling oscillations. The viscous, non-stationary flow around the pitching foil is simulated using ANSYS FLUENT 13. The foil movement is reproduced using the dynamic mesh technique and an in-house developed UDF (User Define Function). Our results show the influence of the pitching frequency and the amplitude on the wake. We provide the mechanisms relating the system behavior to the applied forces. The frequency varies from 1 to 400Hz and the considered amplitudes are 18%, 24%, 30%, 37%, 53%, 82% and 114% of the foil chord.

  11. Generalized multi-point inverse airfoil design

    NASA Technical Reports Server (NTRS)

    Selig, Michael S.; Maughmer, Mark D.

    1991-01-01

    In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution or boundary-layer development, etc., then from this information determine the corresponding airfoil shape. This paper presents a method which approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed. In addition to these local desired distributions, single parameters like the airfoil thickness can be specified. The problem of finding the airfoil shape is determined by coupling an incompressible, inviscid, inverse airfoil design method with a direct integral boundary-layer analysis method and solving the resulting nonlinear equations via a multidimensional Newton iteration technique. The approach is fast and easily allows for interactive design. It is also flexible and could be adapted to solving compressible, inverse airfoil design problems.

  12. Numerical Study of Aerodynamic Characteristics of a Symmetric NACA Section with Simulated Ice Shapes

    NASA Astrophysics Data System (ADS)

    Tabatabaei, N.; Cervantes, M. J.; Trivedi, C.; Aidanpää, Jan-Olof

    2016-09-01

    To develop a numerical model of icing on wind turbine blades, a CFD simulation was conducted to investigate the effect of critical ice accretions on the aerodynamic characteristics of a 0.610 m chord NACA 0011 airfoil section. Aerodynamic performance coefficients and pressure profile were calculated and compared with the available measurements for a chord Reynolds number of 1.83x106. Ice shapes were simulated with flat plates (spoiler-ice) extending along the span of the wing. Lift, drag, and pressure coefficients were calculated in zero angle of attack through the steady state and transient simulations. Different approaches of numerical studies have been applied to investigate the icing conditions on the blades. The simulated separated flow over the sharp spoilers is challenging and can be seen as a worst test case for validation. It allows determining a reliable strategy to simulate real ice shapes [1] for which the detailed validation cannot easily be provided.

  13. Low-speed aerodynamic characteristics of a 42 deg swept high-wing model having a double-slotted flap system and a supercritical airfoil

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Goodson, K. W.

    1974-01-01

    A low-speed investigation was conducted over an angle-of-attack range from about -4 deg to 20 deg in the Langley V/STOL tunnel to determine the effects of a double-slotted flap, high-lift system on the aerodynamic characteristics of a 42 deg swept high-wing model having a supercritical airfoil. The wing had an aspect ratio of 6.78 and a taper ratio of 0.36; the double-slotted flap consisted of a 35-percent-chord flap with a 15-percent-chord vane. The model was tested with a 15-percent-chord leading-edge slat.

  14. Application of holographic interferometry for analysis of the dynamic and modal characteristics of an advanced exotic metal airfoil structure

    NASA Astrophysics Data System (ADS)

    Fein, Howard

    1999-03-01

    Holographic Interferometry has been successfully employed to characterize the materials and behavior of diverse types of structures under stress. Specialized variations of this technology have also been applied to define dynamic and vibration related structural behavior. Such applications of holographic technique offer some of the most effective methods of modal and dynamic analysis available. Real-time dynamic testing of the modal and mechanical behavior of aerodynamic control and airfoil structures for advanced aircraft has always required advanced instrumentation for data collection in either actual flight test or wind-tunnel simulations. Advanced optical holography techniques are alternate methods which result in actual full-field behavioral data on the ground in a noninvasive environment. These methods offer significant insight in both the development and subsequent operational test and modeling of advanced exotic metal control structures and their integration with total vehicle system dynamics. Structures and materials can be analyzed with very low amplitude excitation and the resultant data can be used to adjust the accuracy mathematically derived structural and behavioral models. Holographic Interferometry offers a powerful tool to aid in the developmental engineering of exotic metal structures for high stress applications. Advanced Titanium alloy is a significant example of these sorts of materials which has found continually increased use in advanced aerodynamic, undersea, and other highly mobil platforms. Aircraft applications in particular must consider environments where extremes in vibration and impulsive mechanical stress can affect both operation and structural stability. These considerations present ideal requisites for analysis using advanced holographic methods in the initial design and test of structures made with such advanced materials. Holographic techniques are nondestructive, real- time, and definitive in allowing the identification of

  15. The Low-Speed Characteristics of a 15-Percent Quasi-Elliptical Circulation Control Airfoil with Distributed Camber.

    DTIC Science & Technology

    1979-05-01

    The second difficulty is a limitation of the compressi- 6 bility factor technique. As detailed by Rogers , correction techniques including Karman- Tsien ...the proposed airfoil would accomplish this, a finite difference technique developed by 6 Rogers was employed during the design phase. This procedure...analytically 6 predicted Mcrit is shown in Figure 8 with the Mcrit obtained by applying the Karman- Tsien compressibility correction to the subsonic

  16. Multiple piece turbine engine airfoil with a structural spar

    DOEpatents

    Vance, Steven J.

    2011-10-11

    A multiple piece turbine airfoil having an outer shell with an airfoil tip that is attached to a root with an internal structural spar is disclosed. The root may be formed from first and second sections that include an internal cavity configured to receive and secure the one or more components forming the generally elongated airfoil. The internal structural spar may be attached to an airfoil tip and place the generally elongated airfoil in compression. The configuration enables each component to be formed from different materials to reduce the cost of the materials and to optimize the choice of material for each component.

  17. Supercritical Flow Past Symmetrical Airfoils.

    DTIC Science & Technology

    1980-12-01

    about quasi-elliptic airfoil sections. The method was later extended by Boerstoel [1967] to present a catalog of solutions for certain body shapes. Bauer...Lecture Notes in Economics and Mathematical Systems, Springer- Verlag, New York, 1972. Boerstoel , J. W., "A Survey of Symmetrical Transonic Potential

  18. AFSMO/AFSCL- AIRFOIL SMOOTHING AND SCALING

    NASA Technical Reports Server (NTRS)

    Morgan, H. L

    1994-01-01

    Since its early beginnings, NASA has been actively involved in the design and testing of airfoil sections for a wide variety of applications. Recently a set of programs has been developed to smooth and scale arbitrary airfoil coordinates. The smoothing program, AFSMO, utilizes both least-squares polynomial and least-squares cubic-spline techniques to iteratively smooth the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. The camber and thickness distribution of the smooth airfoil are also computed. The scaling program, AFSCL, may then be used to scale the thickness distribution generated by the smoothing program to a specified maximum thickness. Once the thickness distribution has been scaled, it is combined with the camber distribution to obtain the final scaled airfoil contour. The airfoil smoothing and scaling programs are written in FORTRAN IV for batch execution and have been implemented on a CDC CYBER 170 series computer with a central memory requirement of approximately 70K (octal) of 60 bit words. Both programs generate plotted output via CALCOMP type plotting calls. These programs were developed in 1983.

  19. Low-speed aerodynamic characteristics of a model having a 42 deg swept low wing with a supercritical airfoil, double-slotted flaps, and a T-tail

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Sleeman, W. C., Jr.

    1972-01-01

    A low speed wind tunnel test was conducted in the Langley V/STOL tunnel to determine the static longitudinal and lateral stability characteristics of a general research model which simulated an advance configuration for a commercial transport airplane with a T tail. The model had a 42 deg swept, aspect ratio 6.78 wing with a supercritical airfoil and a high lift system which consisted of a leading edge slat and a double slotted flap. Various slat and flap deflection combinations represented clean, take off, and landing configurations. Effects on the longitudinal and lateral aerodynamic characteristics were determined for two flow through, simulated engine nacelles located on the sides of the fuselage near the rear of the model.

  20. Airfoil shape for a turbine bucket

    DOEpatents

    Hyde, Susan Marie; By, Robert Romany; Tressler, Judd Dodge; Schaeffer, Jon Conrad; Sims, Calvin Levy

    2005-06-28

    Third stage turbine buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth Table I wherein X and Y values are in inches and the Z values are non-dimensional values from 0 to 0.938 convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. The X and Y distances may be scalable as a function of the same constant or number to provide a scaled up or scaled down airfoil section for the bucket. The nominal airfoil given by the X, Y and Z distances lies within an envelop of .+-.0.150 inches in directions normal to the surface of the airfoil.

  1. Section 10: Ground Water - Waste Characteristics & Targets

    EPA Pesticide Factsheets

    HRS Training. The waste characteristics factor category in the ground water pathway is made up of two components: the toxicity/mobility of the most hazardous substance associated with the site and the hazardous waste quantity at the site.

  2. Separated transonic airfoil flow calculations with a nonequilibrium turbulence model

    NASA Technical Reports Server (NTRS)

    King, L. S.; Johnson, D. A.

    1985-01-01

    Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.

  3. Leading edge embedded fan airfoil concept -- A new powered high lift technology

    NASA Astrophysics Data System (ADS)

    Phan, Nhan Huu

    A new powered-lift airfoil concept called Leading Edge Embedded Fan (LEEF) is proposed for Extremely Short Take-Off and Landing (ESTOL) and Vertical Take-Off and Landing (VTOL) applications. The LEEF airfoil concept is a powered-lift airfoil concept capable of generating thrust and very high lift-coefficient at extreme angles-of attack (AoA). It is designed to activate only at the take-off and landing phases, similar to conventional flaps or slats, allowing the aircraft to operate efficiently at cruise in its conventional configuration. The LEEF concept consists of placing a crossflow fan (CFF) along the leading-edge (LE) of the wing, and the housing is designed to alter the airfoil shape between take-off/landing and cruise configurations with ease. The unique rectangular cross section of the crossflow fan allows for its ease of integration into a conventional subsonic wing. This technology is developed for ESTOL aircraft applications and is most effectively applied to General Aviation (GA) aircraft. Another potential area of application for LEEF is tiltrotor aircraft. Unlike existing powered high-lift systems, the LEEF airfoil uses a local high-pressure air source from cross-flow fans, does not require ducting, and is able to be deployed using distributed electric power systems throughout the wing. In addition to distributed lift augmentation, the LEEF system can provide additional thrust during takeoff and landing operation to supplement the primary cruise propulsion system. Two-dimensional (2D) and three-dimensional (3D) Computational Fluid Dynamics (CFD) simulations of a conventional airfoil/wing using the NACA 63-3-418 section, commonly used in GA, and a LEEF airfoil/wing embedded into the same airfoil section were carried out to evaluate the advantages of and the costs associated with implementing the LEEF concept. Computational results show that significant lift and augmented thrust are available during LEEF operation while requiring only moderate fan power

  4. Effectiveness of spoilers on the GA(W)-1 airfoil with a high performance Fowler flap

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.

    1975-01-01

    Two-dimensional wind-tunnel tests were conducted to determine effectiveness of spoilers applied to the GA(W)-1 airfoil. Tests of several spoiler configurations show adequate control effectiveness with flap nested. It is found that providing a vent path allowing lower surface air to escape to the upper surface as the spoiler opens alleviates control reversal and hysteresis tendencies. Spoiler cross-sectional shape variations generally have a modest influence on control characteristics. A series of comparative tests of vortex generators applied to the (GA-W)-1 airfoil show that triangular planform vortex generators are superior to square planform vortex generators of the same span.

  5. Upper-surface modifications for C sub l max improvement of selected NASA 6-series airfoils

    NASA Technical Reports Server (NTRS)

    Szelazek, C. A.; Hicks, R. M.

    1979-01-01

    The thickness of the upper surface of 64 airfoils was increased from the leading edge to the position of maximum thickness. The modifications were generated using a numerical optimization routine coupled with an aerodynamic analysis code. The type of modification presented can be used for aircraft design or for the retrofit of current aircraft to improve the stall characteristics and climb performance. The coordinates of the modified airfoils are presented with plots of the forward 45% of the profiles and pressure distributions for both the modified and unmodified sections at an angle of attack of 14 degrees.

  6. Second-stage turbine bucket airfoil

    DOEpatents

    Wang, John Zhiqiang; By, Robert Romany; Sims, Calvin L.; Hyde, Susan Marie

    2002-01-01

    The second-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X and Y values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket. The second-stage wheel has sixty buckets.

  7. Third-stage turbine bucket airfoil

    DOEpatents

    Pirolla, Peter Paul; Siden, Gunnar Leif; Humanchuk, David John; Brassfield, Steven Robert; Wilson, Paul Stuart

    2002-01-01

    The third-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinates defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.

  8. Trailing edge modifications for flatback airfoils.

    SciTech Connect

    Kahn, Daniel L.; van Dam, C.P.; Berg, Dale E.

    2008-03-01

    The adoption of blunt trailing edge airfoils (also called flatback airfoils) for the inboard region of large wind turbine blades has been proposed. Blunt trailing edge airfoils would not only provide a number of structural benefits, such as increased structural volume and ease of fabrication and handling, but they have also been found to improve the lift characteristics of thick airfoils. Therefore, the incorporation of blunt trailing edge airfoils would allow blade designers to more freely address the structural demands without having to sacrifice aerodynamic performance. These airfoils do have the disadvantage of generating high levels of drag as a result of the low-pressure steady or periodic flow in the near-wake of the blunt trailing edge. Although for rotors, the drag penalty appears secondary to the lift enhancement produced by the blunt trailing edge, high drag levels are of concern in terms of the negative effect on the torque and power generated by the rotor. Hence, devices are sought that mitigate the drag of these airfoils. This report summarizes the literature on bluff body vortex shedding and bluff body drag reduction devices and proposes four devices for further study in the wind tunnel.

  9. Low-speed aerodynamic characteristics of a 16-percent-thick variable-geometry airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Barnwell, R. W.; Noonan, K. W.; Mcghee, R. J.

    1978-01-01

    Tests were conducted in the Langley low-turbulence pressure tunnel to determine the aerodynamic characteristics of climb, cruise, and landing configurations. These tests were conducted over a Mach number range from 0.10 to 0.35, a chord Reynolds number range from 2.0 x 1 million to 20.0 x 1 million, and an angle-of-attack range from -8 deg to 20 deg. Results show that the maximum section lift coefficients increased in the Reynolds number range from 2.0 x 1 million to 9.0 x 1 million and reached values of approximately 2.1, 1.8, and 1.5 for the landing, climb, and cruise configurations, respectively. Stall characteristics, although of the trailing-edge type, were abrupt. The section lift-drag ratio of the climb configuration with fixed transition near the leading edge was about 78 at a lift coefficient of 0.9, a Mach number of 0.15, and a Reynolds number of 4.0 x 1 million. Design lift coefficients of 0.9 and 0.4 for the climb and cruise configurations were obtained at the same angle of attack, about 6 deg, as intended. Good agreement was obtained between experimental results and the predictions of a viscous, attached-flow theoretical method.

  10. Simplified dragonfly airfoil aerodynamics at Reynolds numbers below 8000

    NASA Astrophysics Data System (ADS)

    Levy, David-Elie; Seifert, Avraham

    2009-07-01

    Effective aerodynamics at Reynolds numbers lower than 10 000 is of great technological interest and a fundamental scientific challenge. The current study covers a Reynolds number range of 2000-8000. At these Reynolds numbers, natural insect flight could provide inspiration for technology development. Insect wings are commonly characterized by corrugated airfoils. In particular, the airfoil of the dragonfly, which is able to glide, can be used for two-dimensional aerodynamic study of fixed rigid wings. In this study, a simplified dragonfly airfoil is numerically analyzed in a steady free-stream flow. The aerodynamic performance (such as mean and fluctuating lift and drag), are first compared to a "traditional" low Reynolds number airfoil: the Eppler-E61. The numerical results demonstrate superior performances of the corrugated airfoil. A series of low-speed wind and water tunnel experiments were performed on the corrugated airfoil, to validate the numerical results. The findings indicate quantitative agreement with the mean wake velocity profiles and shedding frequencies while validating the two dimensionality of the flow. A flow physics numerical study was performed in order to understand the underlying mechanism of corrugated airfoils at these Reynolds numbers. Airfoil shapes based on the flow field characteristics of the corrugated airfoil were built and analyzed. Their performances were compared to those of the corrugated airfoil, stressing the advantages of the latter. It was found that the flow which separates from the corrugations and forms spanwise vortices intermittently reattaches to the aft-upper arc region of the airfoil. This mechanism is responsible for the relatively low intensity of the vortices in the airfoil wake, reducing the drag and increasing the flight performances of this kind of corrugated airfoil as compared to traditional low Reynolds number airfoils such as the Eppler E-61.

  11. Low speed aerodynamic characteristics of NACA 6716 and NACA 4416 airfoils with 35 percent-chord single-slotted flaps. [low turbulence pressure tunnel tests to determine two dimensional lift and pitching moment characteristics

    NASA Technical Reports Server (NTRS)

    Bingham, G. J.; Noonan, K. W.

    1974-01-01

    An investigation was conducted in a low-turbulence pressure tunnel to determine the two-dimensional lift and pitching-moment characteristics of an NACA 6716 and an NACA 4416 airfoil with 35-percent-chord single-slotted flaps. Both models were tested with flaps deflected from 0 deg to 45 deg, at angles of attack from minus 6 deg to several degrees past stall, at Reynolds numbers from 3.0 million to 13.8 million, and primarily at a Mach number of 0.23. Tests were also made to determine the effect of several slot entry shapes on performance.

  12. Transonic airfoil analysis and design in nonuniform flow

    NASA Technical Reports Server (NTRS)

    Chang, J. F.; Lan, C. E.

    1986-01-01

    A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.

  13. Design procedure for low-drag subsonic airfoils

    NASA Technical Reports Server (NTRS)

    Peterson, J. B.; Chen, A. B.

    1975-01-01

    Airfoil has least amount of drag under given restrictions of boundary layer transition position, lift coefficient, thickness ratio, and Reynolds number based on airfoil chord. It is suitable for use as wing and propeller aircraft sections operating at subsonic speeds and for hydrofoil sections and blades for fans, compressors, turbines, and windmills.

  14. Development of drive mechanism for an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Sticht, Clifford D.

    1988-01-01

    The design and development of an in-draft wind tunnel test section which will be used to study the dynamic stall of airfoils oscillating in pitch is described. The hardware developed comprises a spanned airfoil between schleiren windows, a four bar linkage, flywheels, a drive system and a test section structure.

  15. Determination of forced convective heat transfer coefficients for subsonic flows over heated asymmetric NANA 4412 airfoil

    NASA Astrophysics Data System (ADS)

    Dag, Yusuf

    Forced convection over traditional surfaces such as flat plate, cylinder and sphere have been well researched and documented. Data on forced convection over airfoil surfaces, however, remain very scanty in literature. High altitude vehicles that employ airfoils as lifting surfaces often suffer leading edge ice accretions which have tremendous negative consequences on the lifting capabilities and stability of the vehicle. One of the ways of mitigating the effect of ice accretion involves judicious leading edge convective cooling technique which in turn depends on the accuracy of convective heat transfer coefficient used in the analysis. In this study empirical investigation of convective heat transfer measurements on asymmetric airfoil is presented at different angle of attacks ranging from 0° to 20° under subsonic flow regime. The top and bottom surface temperatures are measured at given points using Senflex hot film sensors (Tao System Inc.) and used to determine heat transfer characteristics of the airfoils. The model surfaces are subjected to constant heat fluxes using KP Kapton flexible heating pads. The monitored temperature data are then utilized to determine the heat convection coefficients modelled empirically as the Nusselt Number on the surface of the airfoil. The experimental work is conducted in an open circuit-Eiffel type wind tunnel, powered by a 37 kW electrical motor that is able to generate subsonic air velocities up to around 41 m/s in the 24 square-inch test section. The heat transfer experiments have been carried out under constant heat flux supply to the asymmetric airfoil. The convective heat transfer coefficients are determined from measured surface temperature and free stream temperature and investigated in the form of Nusselt number. The variation of Nusselt number is shown with Reynolds number at various angles of attacks. It is concluded that Nusselt number increases with increasing Reynolds number and increase in angle of attack from 0

  16. A computer program for the design and analysis of low-speed airfoils

    NASA Technical Reports Server (NTRS)

    Eppler, R.; Somers, D. M.

    1980-01-01

    A conformal mapping method for the design of airfoils with prescribed velocity distribution characteristics, a panel method for the analysis of the potential flow about given airfoils, and a boundary layer method have been combined. With this combined method, airfoils with prescribed boundary layer characteristics can be designed and airfoils with prescribed shapes can be analyzed. All three methods are described briefly. The program and its input options are described. A complete listing is given as an appendix.

  17. Pressure Distribution Over Airfoils with Fowler Flaps

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Anderson, Walter B

    1938-01-01

    Report presents the results of tests made of a Clark y airfoil with a Clark y Fowler flap and of an NACA 23012 airfoil with NACA Fowler flaps. Some of the tests were made in the 7 by 10-foot wind tunnel and others in the 5-foot vertical wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoils and on the flaps for several angles of attack with the flaps located at the maximum-lift settings. A test installation was used in which the model was mounted in the wind tunnel between large end planes so that two-dimensional flow was approximated. The data are given in the form of pressure-distribution diagrams and as plots of calculated coefficients for the airfoil-and-flap combinations and for the flaps alone.

  18. Airfoil self-noise and prediction

    NASA Technical Reports Server (NTRS)

    Brooks, Thomas F.; Pope, D. Stuart; Marcolini, Michael A.

    1989-01-01

    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.

  19. Low speed airfoil study

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.

    1977-01-01

    Airfoil geometries were developed for low speed high lift applications, such as general aviation aircraft, propellers and helicopter rotors. The primary effort was to determine the extent to which the application of turbulent boundary layer separation criteria, plus manipulation of other input parameters, specifically trailing edging velocity ratio, could be utilized to achieve high C sub Lmax airfoils with relatively low drag at C sub Lmax. Both single-element and double-element airfoils were considered. Wind tunnel testing of some airfoils was included.

  20. Design and test of a natural laminar flow/large Reynolds number airfoil with a high design cruise lift coefficient

    NASA Technical Reports Server (NTRS)

    Kolesar, C. E.

    1987-01-01

    Research activity on an airfoil designed for a large airplane capable of very long endurance times at a low Mach number of 0.22 is examined. Airplane mission objectives and design optimization resulted in requirements for a very high design lift coefficient and a large amount of laminar flow at high Reynolds number to increase the lift/drag ratio and reduce the loiter lift coefficient. Natural laminar flow was selected instead of distributed mechanical suction for the measurement technique. A design lift coefficient of 1.5 was identified as the highest which could be achieved with a large extent of laminar flow. A single element airfoil was designed using an inverse boundary layer solution and inverse airfoil design computer codes to create an airfoil section that would achieve performance goals. The design process and results, including airfoil shape, pressure distributions, and aerodynamic characteristics are presented. A two dimensional wind tunnel model was constructed and tested in a NASA Low Turbulence Pressure Tunnel which enabled testing at full scale design Reynolds number. A comparison is made between theoretical and measured results to establish accuracy and quality of the airfoil design technique.

  1. An analytical study for the design of advanced rotor airfoils

    NASA Technical Reports Server (NTRS)

    Kemp, L. D.

    1973-01-01

    A theoretical study has been conducted to design and evaluate two airfoils for helicopter rotors. The best basic shape, designed with a transonic hodograph design method, was modified to meet subsonic criteria. One airfoil had an additional constraint for low pitching-moment at the transonic design point. Airfoil characteristics were predicted. Results of a comparative analysis of helicopter performance indicate that the new airfoils will produce reduced rotor power requirements compared to the NACA 0012. The hodograph design method, written in CDC Algol, is listed and described.

  2. Experiments on airfoils with trailing edge cut away

    NASA Technical Reports Server (NTRS)

    Ackeret, J

    1927-01-01

    Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords.

  3. Development and testing of airfoils for high-altitude aircraft

    NASA Technical Reports Server (NTRS)

    Drela, Mark (Principal Investigator)

    1996-01-01

    Specific tasks included airfoil design; study of airfoil constraints on pullout maneuver; selection of tail airfoils; examination of wing twist; test section instrumentation and layout; and integrated airfoil/heat-exchanger tests. In the course of designing the airfoil, specifically for the APEX test vehicle, extensive studies were made over the Mach and Reynolds number ranges of interest. It is intended to be representative of airfoils required for lightweight aircraft operating at extreme altitudes, which is the primary research objective of the APEX program. Also considered were thickness, pitching moment, and off-design behavior. The maximum ceiling parameter M(exp 2)C(sub L) value achievable by the Apex-16 airfoil was found to be a strong constraint on the pullout maneuver. The NACA 1410 and 2410 airfoils (inverted) were identified as good candidates for the tail, with predictable behavior at low Reynolds numbers and good tolerance to flap deflections. With regards to wing twist, it was decided that a simple flat wing was a reasonable compromise. The test section instrumentation consisted of surface pressure taps, wake rakes, surface-mounted microphones, and skin-friction gauges. Also, a modest wind tunnel test was performed for an integrated airfoil/heat-exchanger configuration, which is currently on Aurora's 'Theseus' aircraft. Although not directly related to the APEX tests, the aerodynamics or heat exchangers has been identified as a crucial aspect of designing high-altitude aircraft and hence is relevant to the ERAST program.

  4. Analytical studies of new airfoils for wind turbines

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.; Calhoun, J. T.

    1981-01-01

    Computer studies were conducted to analyze the potential gains associated with utilizing new airfoils for large wind turbine rotor blades. Attempts to include 3-dimensional stalling effects were inconclusive. It is recommended that blade pressure measurements be made to clarify the nature of blade stalling. It is also recommended that new laminar flow airfoils be used as rotor blade sections.

  5. Modern Airfoil Ice Accretions

    NASA Technical Reports Server (NTRS)

    Addy, Harold E., Jr.; Potapczuk, Mark G.; Sheldon, David W.

    1997-01-01

    This report presents results from the first icing tests performed in the Modem Airfoils program. Two airfoils have been subjected to icing tests in the NASA Lewis Icing Research Tunnel (IRT). Both airfoils were two dimensional airfoils; one was representative of a commercial transport airfoil while the other was representative of a business jet airfoil. The icing test conditions were selected from the FAR Appendix C envelopes. Effects on aerodynamic performance are presented including the effects of varying amounts of glaze ice as well as the effects of approximately the same amounts of glaze, mixed, and rime ice. Actual ice shapes obtained in these tests are also presented for these cases. In addition, comparisons are shown between ice shapes from the tests and ice shapes predicted by the computer code, LEWICE for similar conditions. Significant results from the tests are that relatively small amounts of ice can have nearly as much effect on airfoil lift coefficient as much greater amounts of ice and that glaze ice usually has a more detrimental effect than either rime or mixed ice. LEWICE predictions of ice shapes, in general, compared reasonably well with ice shapes obtained in the IRT, although differences in details of the ice shapes were observed.

  6. Plasma Flow Control Optimized Airfoil

    NASA Astrophysics Data System (ADS)

    Voikov, Vladimir; Patel, Mehul

    2005-11-01

    Recent advances in flow control research have demonstrated that plasma actuators can be efficient in different aerodynamic applications, particularly in providing flight control without conventional moving surfaces. The concept involves the use of a laminar airfoil design that employs a separation ramp at the trailing edge that can be manipulated by a plasma actuator to control lift, similar to trailing-edge flaps. The advantages are lower drag by a combination of the laminar flow design, and elimination of parasitic drag associated with wing-flap junctions. This work involves numerical simulations and experiments on a HSNLF(1)-0213 airfoil. The numerical results are obtained using an unsteady, compressible Navier-Stokes simulation that includes a model for the plasma actuators. The experiments are performed on a 2-D airfoil section that is mounted on a lift-drag force balance. The results demonstrate lift enhancement produced by the plasma actuator that is comparable to a plane flap. They also reveal an optimum actuator unsteady frequency that scales with the length of the separated region and local velocity, and is associated with the generation of a train of spanwise vortices. Other scaling including the effect of Reynolds number is presented.

  7. Multiple piece turbine airfoil

    DOEpatents

    Kimmel, Keith D; Wilson, Jr., Jack W.

    2010-11-02

    A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of dog bone struts each mounted within openings formed within the shell and spar to allow for relative motion between the spar and shell in the airfoil chordwise direction while also forming a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure.

  8. Two experimental supercritical laminar-flow-control swept-wing airfoils

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Dagenhart, J. Ray

    1987-01-01

    Two supercritical laminar-flow-control airfoils were designed for a large-chord swept-wing experiment in the Langley 8-Foot Transonic Pressure Tunnel where suction was provided through most of the model surface for boundary-layer control. The first airfoil was derived from an existing full-chord laminar airfoil by extending the trailing edge and making changes in the two lower-surface concave regions. The second airfoil differed from the first one in that it was designed for testing without suction in the forward concave region of the lower surface. Differences between the first airfoil and the one from which it was derived as well as between the first and second airfoils are discussed. Airfoil coordinates and predicted pressure distributions for the design normal Mach number of 0.755 and section lift coefficient of 0.55 are given for the three airfoils.

  9. Closed loop steam cooled airfoil

    DOEpatents

    Widrig, Scott M.; Rudolph, Ronald J.; Wagner, Gregg P.

    2006-04-18

    An airfoil, a method of manufacturing an airfoil, and a system for cooling an airfoil is provided. The cooling system can be used with an airfoil located in the first stages of a combustion turbine within a combined cycle power generation plant and involves flowing closed loop steam through a pin array set within an airfoil. The airfoil can comprise a cavity having a cooling chamber bounded by an interior wall and an exterior wall so that steam can enter the cavity, pass through the pin array, and then return to the cavity to thereby cool the airfoil. The method of manufacturing an airfoil can include a type of lost wax investment casting process in which a pin array is cast into an airfoil to form a cooling chamber.

  10. Effect of Compressibility on Pressure Distribution over an Airfoil with a Slotted Frise Aileron

    NASA Technical Reports Server (NTRS)

    Luoma, Avro A

    1944-01-01

    Pressure distribution measurements were made over an airfoil with slotted Frise aileron up to 0.76 Mach at various angles of attack and aileron defections. Section characteristics were determined from these pressure data. Results indicated loss of aileron rolling power for deflections ranging from -12 Degrees to -19 Degrees. High stick forces for non-differential deflections incurred at high speed, which were due to overbalancing tendency of up-moving aileron, may precipitate serious control difficulties. Detailed results are presented graphically.

  11. An efficient algorithm for numerical airfoil optimization

    NASA Technical Reports Server (NTRS)

    Vanderplaats, G. N.

    1979-01-01

    A new optimization algorithm is presented. The method is based on sequential application of a second-order Taylor's series approximation to the airfoil characteristics. Compared to previous methods, design efficiency improvements of more than a factor of 2 are demonstrated. If multiple optimizations are performed, the efficiency improvements are more dramatic due to the ability of the technique to utilize existing data. The method is demonstrated by application to subsonic and transonic airfoil design but is a general optimization technique and is not limited to a particular application or aerodynamic analysis.

  12. Transonic flow theory of airfoils and wings

    NASA Technical Reports Server (NTRS)

    Garabedian, P. R.

    1976-01-01

    There are plans to use the supercritical wing on the next generation of commercial aircraft so as to economize on fuel consumption by reducing drag. Computer codes have served well in meeting the consequent demand for new wing sections. The possibility of replacing wind tunnel tests by computational fluid dynamics is discussed. Another approach to the supercritical wing is through shockless airfoils. A novel boundary value problem in the hodograph plane is studied that enables one to design a shockless airfoil so that its pressure distribution very nearly takes on data that are prescribed.

  13. NASA supercritical airfoils: A matrix of family-related airfoils

    NASA Technical Reports Server (NTRS)

    Harris, Charles D.

    1990-01-01

    The NASA supercritical airfoil development program is summarized in a chronological fashion. Some of the airfoil design guidelines are discussed, and coordinates of a matrix of family related supercritical airfoils ranging from thicknesses of 2 to 18 percent and over a design lift coefficient range from 0 to 1.0 are presented.

  14. Tests of NACA 0009, 0012, and 0018 Airfoils in the Full-Scale Tunnel

    NASA Technical Reports Server (NTRS)

    Goett, Harry J; Bullivant, W Kenneth

    1939-01-01

    An investigation was conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the NACA 0009, 0012, and 0018 airfoils, with the ultimate purpose of providing data to be used as a basis for comparison with other wind-tunnel data, mainly in the study of scale and turbulence effects. Three symmetrical 6 by 36-foot rectangular airfoils were used. The Reynolds number range for minimum drag was form 1,800,000 to 7,000,000 and for maximum lift, from 1,700,000 to 4,500,000. The effect of rounded tips was determined for each of the airfoils. Tests were also made of the NACA 0012 airfoil equipped with a 0.20c full-span split flap hinged at 0.80c. Tuft surveys were included to show the progressive breakdown of flow near maximum lift. Momentum surveys were made in conjunction with force measurements at zero lift as an aid in converting force-test data to section coefficients.

  15. Airfoils for wind turbine

    SciTech Connect

    Tangler, J.L.; Somers, D.M.

    2000-05-30

    Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

  16. Airfoils for wind turbine

    DOEpatents

    Tangler, James L.; Somers, Dan M.

    2000-01-01

    Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

  17. Computation of turbulent near wake for asymmetric airfoils

    NASA Technical Reports Server (NTRS)

    Deiwert, G. S.

    1979-01-01

    A numerical procedure for studying the turbulent near wake of two dimensional airfoil sections is presented. The Reynolds Navier-Stokes equations were written for flow about bodies of arbitrary geometry and solved on an arbitrary nonuniform curvilinear computational mesh. Eddy viscosity and Reynolds stress turbulence transport models are considered. Specific examples are shown for airfoil section by using an algebraic viscosity model with streamwise relaxation and the interactive Reynolds stress model.

  18. Robust, Optimal Subsonic Airfoil Shapes

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan

    2014-01-01

    A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.

  19. Airfoil family design for large offshore wind turbine blades

    NASA Astrophysics Data System (ADS)

    Méndez, B.; Munduate, X.; San Miguel, U.

    2014-06-01

    , compatibility for the different airfoil family members, etc.) and with the ultimate objective that the airfoils will reduce the blade loads. In this paper the whole airfoil design process and the main characteristics of the airfoil family are described. Some force coefficients for the design Reynolds number are also presented. The new designed airfoils have been studied with computational calculations (panel method code and CFD) and also in a wind tunnel experimental campaign. Some of these results will be also presented in this paper.

  20. Iced-airfoil aerodynamics

    NASA Astrophysics Data System (ADS)

    Bragg, M. B.; Broeren, A. P.; Blumenthal, L. A.

    2005-07-01

    Past research on airfoil aerodynamics in icing are reviewed. This review emphasizes the time period after the 1978 NASA Lewis workshop that initiated the modern icing research program at NASA and the current period after the 1994 ATR accident where aerodynamics research has been more aircraft safety focused. Research pre-1978 is also briefly reviewed. Following this review, our current knowledge of iced airfoil aerodynamics is presented from a flowfield-physics perspective. This article identifies four classes of ice accretions: roughness, horn ice, streamwise ice, and spanwise-ridge ice. For each class, the key flowfield features such as flowfield separation and reattachment are discussed and how these contribute to the known aerodynamic effects of these ice shapes. Finally Reynolds number and Mach number effects on iced-airfoil aerodynamics are summarized.

  1. Multiple piece turbine airfoil

    DOEpatents

    Kimmel, Keith D

    2010-11-09

    A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of hook shaped struts each mounted within channels extending in a spanwise direction of the spar and the shell to allow for relative motion between the spar and shell in the airfoil chordwise direction while also fanning a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure. The hook struts have a hooked shaped end and a rounded shaped end in order to insert the struts into the spar.

  2. An experimental low Reynolds number comparison of a Wortmann FX67-K170 airfoil, a NACA 0012 airfoil and a NACA 64-210 airfoil in simulated heavy rain

    NASA Technical Reports Server (NTRS)

    Craig, Anthony P.; Hansman, R. John

    1987-01-01

    Wind tunnel experiments were conducted on Wortmann FX67-K170, NACA 0012, and NACA 64-210 airfoils at rain rates of 1000 mm/hr and Reynolds numbers of 310,000 to compare the aerodynamic performance degradation of the airfoils and to attempt to identify the various mechanisms which affect performance in heavy rain conditions. Lift and drag were measured in dry and wet conditions, a variety of flow visualization techniques were employed, and a computational code which predicted airfoil boundary layer behavior was used. At low angles of attack, the lift degradation in wet conditions varied significantly between the airfoils. The Wortmann section had the greatest overall lift degradation and the NACA 64-210 airfoil had the smallest. At high angles of attack, the NACA 64-210 and 0012 airfoils had improved aerodynamic performance in rain conditions due to an apparent reduction of the boundry layer separation. Performance degradation in heavy rain for all three airfoils at low angles of attack could be emulated by forced boundary layer transition near the leading edge. The secondary effect occurs at time scales consistent with top surface water runback times. The runback layer is thought to effectively alter the airfoil geometry. The severity of the performance degradation for the airfoils varied. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary layer transition.

  3. Wind tunnel test of the S814 thick root airfoil

    SciTech Connect

    Somers, D.M.; Tangler, J.L.

    1996-11-01

    The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient designed to be largely insensitive to leading-edge roughness effects. The 24 percent thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur at a high lift coefficient. To accomplish the objective, a two-dimensional wind tunnel test of the S814 thick root airfoil was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. Data were obtained with transition free and transition fixed for Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds number of 1.5 {times} 10{sup 6}, the maximum lift coefficient with transition free is 1.32, which satisfies the design specification. However, this value is significantly lower than the predicted maximum lift coefficient of almost 1.6. With transition fixed at the leading edge, the maximum lift coefficient is 1.22. The small difference in maximum lift coefficient between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low maximum-lift-coefficient tip-region airfoils for rotor blades 10 to 15 meters in length.

  4. An Approach to the Constrained Design of Natural Laminar Flow Airfoils

    NASA Technical Reports Server (NTRS)

    Green, Bradford E.

    1997-01-01

    A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integral turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the laminar flow toward the desired amount. An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

  5. An approach to the constrained design of natural laminar flow airfoils

    NASA Technical Reports Server (NTRS)

    Green, Bradford Earl

    1995-01-01

    A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

  6. Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: Preprint

    SciTech Connect

    Migliore, P.; Oerlemans, S.

    2003-12-01

    Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils.

  7. Design of a shape adaptive airfoil actuated by a Shape Memory Alloy strip for airplane tail

    NASA Astrophysics Data System (ADS)

    Shirzadeh, R.; Raissi Charmacani, K.; Tabesh, M.

    2011-04-01

    Of the factors that mainly affect the efficiency of the wing during a special flow regime, the shape of its airfoil cross section is the most significant. Airfoils are generally designed for a specific flight condition and, therefore, are not fully optimized in all flight conditions. It is very desirable to have an airfoil with the ability to change its shape based on the current regime. Shape memory alloy (SMA) actuators activate in response to changes in the temperature and can recover their original configuration after being deformed. This study presents the development of a method to control the shape of an airfoil using SMA actuators. To predict the thermomechanical behaviors of an SMA thin strip, 3D incremental formulation of the SMA constitutive model is implemented in FEA software package ABAQUS. The interactions between the airfoil structure and SMA thin strip actuator are investigated. Also, the aerodynamic performance of a standard airfoil with a plain flap is compared with an adaptive airfoil.

  8. Turbine airfoil film cooling

    NASA Technical Reports Server (NTRS)

    Hylton, Larry D.

    1986-01-01

    Emphasis is placed on developing more accurate analytical models for predicting turbine airfoil external heat transfer rates. Performance goals of new engines require highly refined, accurate design tools to meet durability requirements. In order to obtain improvements in analytical capabilities, programs are required which focus on enhancing analytical techniques through verification of new models by comparison with relevant experimental data. The objectives of the current program are to develop an analytical approach, based on boundary layer theory, for predicting the effects of airfoil film cooling on downstream heat transfer rates and to verify the resulting analytical method by comparison of predictions with hot cascade data obtained under this program.

  9. Turbine airfoil with controlled area cooling arrangement

    SciTech Connect

    Liang, George

    2010-04-27

    A gas turbine airfoil (10) includes a serpentine cooling path (32) with a plurality of channels (34,42,44) fluidly interconnected by a plurality of turns (38,40) for cooling the airfoil wall material. A splitter component (50) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel (46) passing in between the outer wall (28) and the inner wall (30) of the pressure side (24) and a suction-side channel (48) passing in between the outer wall (28) and the inner wall (30) of the suction side (26) longitudinally downstream of an intermediate height (52). The cross-sectional area of the pressure-side channel (46) and suction-side channel (48) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.

  10. Characterization of the Effect of Wing Surface Instrumentation on UAV Airfoil Performance

    NASA Technical Reports Server (NTRS)

    Ratnayake, Nalin A.

    2009-01-01

    Recently proposed flight research at NASA Dryden Flight Research Center (DFRC) has prompted study into the aerodynamic effects of modifications made to the surfaces of laminar airfoils. The research is focused on the high-aspect ratio, laminar-flow type wings commonly found on UAVs and other aircraft with a high endurance requirement. A broad range of instrumentation possibilities, such as structural, pressure, and temperature sensing devices may require the alteration of the airfoil outer mold line as part of the installation process. This study attempts to characterize the effect of installing this additiona1 instrumentation on key airfoil performance factors, such as transition location, lift and drag curves, and stall point. In particular, the general case of an airfoil that is channeled in the spanwise direction is considered, and the impact on key performance characteristics is assessed. Particular attention is focused on exploring the limits of channel depth and low-Reynolds number on performance and stall characteristics. To quantify the effect of increased skin friction due to premature transition caused by protruding or recessed instrumentation, two simplified, conservative scenarios are used to consider two potential sources of diaturbance: A) that leading edge alterations would cause linearly expanding areas (triangles) of turbulent flow on both surfaces of the wing upstream of the natural transition point, and B) that a channel or bump on the upper surface would trip turbulent flow across the whole upper surface upstream of the natural transition point. A potentially more important consideration than the skin friction drag increment is the change in overall airfoil performance due to the installation of instrumentation along most of the wingspan. To quantify this effect, 2D CFD simulations of the flow over a representative mid-span airfoil section were conducted in order to assess the change in lift and drag curves for the airfoil in the presence of

  11. Low-speed single-element airfoil synthesis

    NASA Technical Reports Server (NTRS)

    Mcmasters, J. H.; Henderson, M. L.

    1979-01-01

    The use of recently developed airfoil analysis/design computational tools to clarify, enrich and extend the existing experimental data base on low-speed, single element airfoils is demonstrated. A discussion of the problem of tailoring an airfoil for a specific application at its appropriate Reynolds number is presented. This problem is approached by use of inverse (or synthesis) techniques, wherein a desirable set of boundary layer characteristics, performance objectives, and constraints are specified, which then leads to derivation of a corresponding viscous flow pressure distribution. Examples are presented which demonstrate the synthesis approach, following presentation of some historical information and background data which motivate the basic synthesis process.

  12. Performance of two transonic airfoil wind tunnels utilizing limited ventilation

    NASA Technical Reports Server (NTRS)

    Lee, J. D.; Gregorek, G. M.

    1984-01-01

    A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.

  13. Navier-Stokes analysis of blunt trailing edge airfoils

    NASA Technical Reports Server (NTRS)

    Stanaway, Sharon; Mccroskey, W. J.; Kroo, Ilan

    1992-01-01

    The flow around blunt trailing edge airfoils was studied by solving the Reynolds-averaged Navier-Stokes equations. The solution procedure combines a grid around the airfoil with a second grid for the wake so that the time advancement over the domain is fully implicit. This is not only very efficient for the algorithm but also allows implicit solutions of a one equation turbulence model appropriate for both boundary layers and wakes. An algebraic and two one-equation turbulence models are tested for a blunt RAE 2822 airfoil section and detailed comparisons with experimental data are presented in the trailing edge region.

  14. Second-order subsonic airfoil theory including edge effects

    NASA Technical Reports Server (NTRS)

    Van Dyke, Milton D

    1956-01-01

    Several recent advances in plane subsonic flow theory are combined into a unified second-order theory for airfoil sections of arbitrary shape. The solution is reached in three steps: the incompressible result is found by integration, it is converted into the corresponding subsonic compressible result by means of the second-order compressibility rule, and it is rendered uniformly valid near stagnation points by further rules. Solutions for a number of airfoils are given and are compared with the results of other theories and of experiment. A straight-forward computing scheme is outlined for calculating the surface velocities and pressures on any airfoil at any angle of attack

  15. Multi-Element Airfoil System

    NASA Technical Reports Server (NTRS)

    Turner, Travis L. (Inventor); Khorrami, Mehdi R. (Inventor); Lockard, David P. (Inventor); McKenney, Martin J. (Inventor); Atherley, Raymond D. (Inventor); Kidd, Reggie T. (Inventor)

    2014-01-01

    A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.

  16. Airfoil Design and Rotorcraft Performance

    NASA Technical Reports Server (NTRS)

    Bousman, William G.

    2003-01-01

    The relationship between global performance of a typical helicopter and the airfoil environment, as represented by the airfoil angles of attack and Mach number, has been examined using the comprehensive analysis CAMRAD II. A general correspondence is observed between global performance parameters, such as rotor L/D, and airfoil performance parameters, such as airfoil L/D, the drag bucket boundaries, and the divergence Mach number. Effects of design parameters such as blade twist and rotor speed variation have been examined and, in most cases, improvements observed in global performance are also observed in terms of airfoil performance. The relations observed between global Performance and the airfoil environment suggests that the emphasis in airfoil design should be for good L/D, while the maximum lift coefficient performance is less important.

  17. The Effects of the Critical Ice Accretion on Airfoil and Wing Performance

    NASA Technical Reports Server (NTRS)

    Selig, Michael S.; Bragg, Michael B.; Saeed, Farooq

    1998-01-01

    In support of the NASA Lewis Modern Airfoils Ice Accretion Test Program, the University of Illinois at Urbana-Champaign provided expertise in airfoil design and aerodynamic analysis to determine the aerodynamic effect of ice accretion on modern airfoil sections. The effort has concentrated on establishing a design/testing methodology for "hybrid airfoils" or "sub-scale airfoils," that is, airfoils having a full-scale leading edge together with a specially designed and foreshortened aft section. The basic approach of using a full-scale leading edge with a foreshortened aft section was considered to a limited extent over 40 years ago. However, it was believed that the range of application of the method had not been fully exploited. Thus a systematic study was being undertaken to investigate and explore the range of application of the method so as to determine its overall potential.

  18. Characteristics-based sectional modeling of aerosol nucleation and condensation

    NASA Astrophysics Data System (ADS)

    Frederix, E. M. A.; Stanic, M.; Kuczaj, A. K.; Nordlund, M.; Geurts, B. J.

    2016-12-01

    A new numerical method for the solution of an internally mixed spatially homogeneous sectional model for aerosol nucleation and condensation is proposed. The characteristics method is used to predict droplet sizes within a discrete time step. The method is designed such that 1) a pre-specified number of moments of the droplet size distribution may be preserved, 2) there exists no time step stability restriction related to the condensation rate and section size, 3) highly skewed fixed sectional distributions may be used and 4) it is straightforward to extend to spatially inhomogeneous settings and to incorporate droplet coagulation and break-up. We derive, starting from mass conservation, a consistent internally mixed multi-species aerosol model. For certain condensational growth laws analytical solutions exist, against which the method is validated. Using two-moment and four-moment-preserving schemes, we find first order convergence of the numerical solution to the analytical result, as a function of the number of sections. As the four-moment-preserving scheme does not guarantee positivity of the solution, a hybrid scheme is proposed, which, when needed, locally reverts back to two-moment preservation, to prevent negativity. As an illustration, the method is applied to a complete multi-species homogeneous nucleation and condensation problem.

  19. First-stage high pressure turbine bucket airfoil

    DOEpatents

    Brown, Theresa A.; Ahmadi, Majid; Clemens, Eugene; Perry, II, Jacob C.; Holiday, Allyn K.; Delehanty, Richard A.; Jacala, Ariel Caesar

    2004-05-25

    The first-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinates defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.

  20. Technology for pressure-instrumented thin airfoil models

    NASA Technical Reports Server (NTRS)

    Wigley, David A.

    1988-01-01

    A novel method of airfoil model construction was developed. This Laminated Sheet technique uses 0.8 mm thick sheets of A286 containing a network of pre-formed channels which are vacuum brazed together to form the airfoil. A 6.25 percent model of the X29A canard, which has a 5 percent thick section, was built using this technique. The model contained a total of 96 pressure orifices, 56 in three chordwise rows on the upper surface and 37 in three similar rows on the lower surface. It was tested in the NASA Langley 0.3 m Transonic Cryogenic Tunnel. Unique aerodynamic data was obtained over the full range of temperature and pressure. Part of the data was at transonic Mach numbers and flight Reynolds number. A larger two dimensional model of the NACA 64a-105 airfoil section was also fabricated. Scale up presented some problems, but a testable airfoil was fabricated.

  1. Efficiency of an auto-propelled flapping airfoil

    NASA Astrophysics Data System (ADS)

    Benkherouf, T.; Mekadem, M.; Oualli, H.; Hanchi, S.; Keirsbulck, L.; Labraga, L.

    2011-05-01

    The present study deals with an investigation of the flow aerodynamic characteristics and the propulsive velocity of a system equipped with a nature inspired propulsion system. In particular, the study is aimed at studying the effect of the flapping frequency on the flow behavior. We consider a NACA0014 airfoil undergoing a vertical sinusoidal flapping motion. In contrast to nearly all previous studies in the literature, the present work does not impose any velocity on the inlet flow. During each iteration the outer flow velocity is computed after having determined the forces exerted on the airfoil. Forward motion may only be produced by flapping motion of the airfoil. This is more consistent with the physical phenomenon. The non-stationary viscous flow around the flapping airfoil is simulated using Ansys-Fluent 12.0.7. The airfoil movement is achieved using the deformable mesh technique and an in-house developed User Define Function (UDF). Our results show the influence of flapping frequency and amplitude on both the airfoil velocity and the propulsive efficiency. The resulting motion is contrasts to the applied forces. In the present study, the frequency ranges from 0.1 to 20 Hz while the airfoil amplitude values considered are: 10%, 17.5%, 25% and 40%.

  2. Aerodynamic properties of thick airfoils II

    NASA Technical Reports Server (NTRS)

    Norton, F H; Bacon, D L

    1923-01-01

    This investigation is an extension of NACA report no. 75 for the purpose of studying the effect of various modifications in a given wing section, including changes in thickness, height of lower camber, taper in thickness, and taper in plan form with special reference to the development of thick, efficient airfoils. The method consisted in testing the wings in the NACA 5-foot wind tunnel at speeds up to 50 meters (164 feet) per second while they were being supported on a new type of wire balance. Some of the airfoils developed showed results of great promise. For example, one wing (no. 81) with a thickness in the center of 4.5 times that of the U. S. A. 16 showed both uniformly high efficiency and a higher maximum lift than this excellent section. These thick sections will be especially useful on airplanes with cantilever construction. (author)

  3. Experimental, water droplet impingement data on two-dimensional airfoils, axisymmetric inlet and Boeing 737-300 engine inlet

    NASA Technical Reports Server (NTRS)

    Papadakis, M.; Elangovan, E.; Freund, G. A., Jr.; Breer, M. D.

    1987-01-01

    An experimental method has been developed to determine the droplet impingement characteristics on two- and three-dimensional bodies. The experimental results provide the essential droplet impingement data required to validate particle trajectory codes, used in aircraft icing analyses and engine inlet particle separator analyses. A body whose water droplet impingement characteristics are required is covered at strategic locations by thin strips of moisture absorbing (blotter) paper, and then exposed to an air stream containing a dyed-water spray cloud. Water droplet impingement data are extracted from the dyed blotter strips, by measuring the optical reflectance of the dye deposit on the strips, using an automated reflectometer. Impingement efficiency data obtained for a NACA 65(2)015 airfoil section, a supercritical airfoil section, and Being 737-300 and axisymmetric inlet models are presented in this paper.

  4. Illustration of airfoil shape effect on forward-swept wing divergence

    NASA Technical Reports Server (NTRS)

    Bland, S. R.

    1980-01-01

    A static aeroelastic analysis is presented of the divergence of untapered wings with conventional and supercritical airfoil sections at sweep angles of zero and -15 deg. One bending and one torsion mode were employed for a uniform rectangular cantilevered beam with the elastic axis at midchord, and calculations were based on a two-dimensional differential equations formulation in the structural coordinate system and in simple strip theory. A minimum divergence speed in the transonic range is obtained which is associated with the rearward shift of the aerodynamic center, and a 17% difference in minimum divergence dynamic pressure is found between a supercritical and a conventional wing. It is noted that although the strip method employed allows the assessment of the sensitivity of airfoil shapes to divergence, three-dimensional transonic aerodynamic methods should be used to predict wing divergence characteristics.

  5. Improved design of subcritical and supercritical cascades using complex characteristics and boundary layer correction

    NASA Technical Reports Server (NTRS)

    Sanz, J. M.

    1983-01-01

    The method of complex characteristics and hodograph transformation for the design of shockless airfoils was extended to design supercritical cascades with high solidities and large inlet angles. This capability was achieved by introducing a conformal mapping of the hodograph domain onto an ellipse and expanding the solution in terms of Tchebycheff polynomials. A computer code was developd based on this idea. A number of airfoils designed with the code are presented. Various supercritical and subcritical compressor, turbine and propeller sections are shown. The lag-entrainment method for the calculation of a turbulent boundary layer was incorporated to the inviscid design code. The results of this calculation are shown for the airfoils described. The elliptic conformal transformation developed to map the hodograph domain onto an ellipse can be used to generate a conformal grid in the physical domain of a cascade of airfoils with open trailing edges with a single transformation. A grid generated with this transformation is shown for the Korn airfoil.

  6. Vertical axis wind turbine airfoil

    DOEpatents

    Krivcov, Vladimir; Krivospitski, Vladimir; Maksimov, Vasili; Halstead, Richard; Grahov, Jurij Vasiljevich

    2012-12-18

    A vertical axis wind turbine airfoil is described. The wind turbine airfoil can include a leading edge, a trailing edge, an upper curved surface, a lower curved surface, and a centerline running between the upper surface and the lower surface and from the leading edge to the trailing edge. The airfoil can be configured so that the distance between the centerline and the upper surface is the same as the distance between the centerline and the lower surface at all points along the length of the airfoil. A plurality of such airfoils can be included in a vertical axis wind turbine. These airfoils can be vertically disposed and can rotate about a vertical axis.

  7. Viscous effects on transonic airfoil stability and response

    NASA Technical Reports Server (NTRS)

    Berry, H. M.; Batina, J. T.; Yang, T. Y.

    1985-01-01

    Viscous effects on transonic airfoil stability and response are investigated using an integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code. Unsteady transonic airloads required for stability analyses are computed using a pulse transfer function analysis including viscous effects. The pulse analysis provides unsteady aerodynamic forces for a wide range of reduced frequency in a single flow field computation. Nonlinear time marching aeroelastic solutions are presented which show the effects of viscosity on airfoil response behavior and flutter. Effects of amplitude on time marching responses are demonstrated. A state space aeroelastic model employing Pade approximants to describe the unsteady airloads is used to study the effects of viscosity on transonic airfoil stability. State space dynamic pressure root loci are in good overall agreement with time marching damping and frequency estimates. Parallel sets of results with and without viscous effects reveal the effects of viscosity on transonic unsteady airloads and aeroelastic characteristics of airfoils.

  8. Study of the TRAC Airfoil Table Computational System

    NASA Technical Reports Server (NTRS)

    Hu, Hong

    1999-01-01

    The report documents the study of the application of the TRAC airfoil table computational package (TRACFOIL) to the prediction of 2D airfoil force and moment data over a wide range of angle of attack and Mach number. The TRACFOIL generates the standard C-81 airfoil table for input into rotorcraft comprehensive codes such as CAM- RAD. The existing TRACFOIL computer package is successfully modified to run on Digital alpha workstations and on Cray-C90 supercomputers. A step-by-step instruction for using the package on both computer platforms is provided. Application of the newer version of TRACFOIL is made for two airfoil sections. The C-81 data obtained using the TRACFOIL method are compared with those of wind-tunnel data and results are presented.

  9. Shape Changing Airfoil

    NASA Technical Reports Server (NTRS)

    Ott, Eric A.

    2005-01-01

    Scoping of shape changing airfoil concepts including both aerodynamic analysis and materials-related technology assessment effort was performed. Three general categories of potential components were considered-fan blades, booster and compressor blades, and stator airfoils. Based on perceived contributions to improving engine efficiency, the fan blade was chosen as the primary application for a more detailed assessment. A high-level aerodynamic assessment using a GE90-90B Block 4 engine cycle and fan blade geometry indicates that blade camber changes of approximately +/-4deg would be sufficient to result in fan efficiency improvements nearing 1 percent. Constraints related to flight safety and failed mode operation suggest that use of the baseline blade shape with actuation to the optimum cruise condition during a portion of the cycle would be likely required. Application of these conditions to the QAT fan blade and engine cycle was estimated to result in an overall fan efficiency gain of 0.4 percent.

  10. An experimental study of a bio-inspired corrugated airfoil for micro air vehicle applications

    NASA Astrophysics Data System (ADS)

    Murphy, Jeffery T.; Hu, Hui

    2010-08-01

    An experimental study was conducted to investigate the aerodynamic characteristics of a bio-inspired corrugated airfoil compared with a smooth-surfaced airfoil and a flat plate at the chord Reynolds number of Re C = 58,000-125,000 to explore the potential applications of such bio-inspired corrugated airfoils for micro air vehicle designs. In addition to measuring the aerodynamic lift and drag forces acting on the tested airfoils, a digital particle image velocimetry system was used to conduct detailed flowfield measurements to quantify the transient behavior of vortex and turbulent flow structures around the airfoils. The measurement result revealed clearly that the corrugated airfoil has better performance over the smooth-surfaced airfoil and the flat plate in providing higher lift and preventing large-scale flow separation and airfoil stall at low Reynolds numbers (Re C < 100,000). While aerodynamic performance of the smooth-surfaced airfoil and the flat plate would vary considerably with the changing of the chord Reynolds numbers, the aerodynamic performance of the corrugated airfoil was found to be almost insensitive to the Reynolds numbers. The detailed flow field measurements were correlated with the aerodynamic force measurement data to elucidate underlying physics to improve our understanding about how and why the corrugation feature found in dragonfly wings holds aerodynamic advantages for low Reynolds number flight applications.

  11. The S415 and S418 Airfoils

    DTIC Science & Technology

    2010-08-01

    airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 4.) This characteristic is related to the...edge with increasing (decreasing) lift coefficient. This feature results in a leading-edge shape that produces a suction peak at higher lift...should look like sketch 3. Sketch 3 1Director, Institute for Aerodynamics and Gas Dynamics, University of Stuttgart, Germany, 1974–1985.5 No suction

  12. An extended theory of thin airfoils and its application to the biplane problem

    NASA Technical Reports Server (NTRS)

    Millikan, Clark B

    1931-01-01

    The report presents a new treatment, due essentially to von Karman, of the problem of the thin airfoil. The standard formulae for the angle of zero lift and zero moment are first developed and the analysis is then extended to give the effect of disturbing or interference velocities, corresponding to an arbitrary potential flow, which are superimposed on a normal rectilinear flow over the airfoil. An approximate method is presented for obtaining the velocities induced by a 2-dimensional airfoil at a point some distance away. In certain cases this method has considerable advantage over the simple "lifting line" procedure usually adopted. The interference effects for a 2-dimensional biplane are considered in the light of the previous analysis. The results of the earlier sections are then applied to the general problem of the interference effects for a 3-dimensional biplane, and formulae and charts are given which permit the characteristics of the individual wings of an arbitrary biplane without sweepback or dihedral to be calculated. In the final section the conclusions drawn from the application of the theory to a considerable number of special cases are discussed, and curves are given illustrating certain of these conclusions and serving as examples to indicate the nature of the agreement between the theory and experiment.

  13. NREL airfoil families for HAWTs

    SciTech Connect

    Tangler, J.L.; Somers, D.M.

    1995-12-31

    The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time nine airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub 1,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.

  14. NREL airfoil families for HAWTs

    SciTech Connect

    Tangler, J L; Somers, D M

    1995-01-01

    The development of special-purpose airfoils for horizontal-axis wind turbines (HAWTs) began in 1984 as a joint effort between the National Renewable Energy Laboratory (NREL), formerly the Solar Energy Research Institute (SERI), and Airfoils, Incorporated. Since that time seven airfoil families have been designed for various size rotors using the Eppler Airfoil Design and Analysis Code. A general performance requirement of the new airfoil families is that they exhibit a maximum lift coefficient (c{sub l,max}) which is relatively insensitive to roughness effects. The airfoil families address the needs of stall-regulated, variable-pitch, and variable-rpm wind turbines. For stall-regulated rotors, better peak-power control is achieved through the design of tip airfoils that restrain the maximum lift coefficient. Restrained maximum lift coefficient allows the use of more swept disc area for a given generator size. Also, for stall-regulated rotors, tip airfoils with high thickness are used to accommodate overspeed control devices. For variable-pitch and variable-rpm rotors, tip airfoils having a high maximum lift coefficient lend themselves to lightweight blades with low solidity. Tip airfoils having low thickness result in less drag for blades having full-span pitch control. Annual energy improvements from the NREL airfoil families are projected to be 23% to 35% for stall-regulated turbines, 8% to 20% for variable-pitch turbines, and 8% to 10% for variable-rpm turbines. The improvement for stall-regulated turbines has been verified in field tests.

  15. Low-speed wind-tunnel results for symmetrical NASA LS(1)-0013 airfoil

    NASA Technical Reports Server (NTRS)

    Ferris, James C.; Mcghee, Robert J.; Barnwell, Richard W.

    1987-01-01

    A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.

  16. Inverse airfoil design procedure using a multigrid Navier-Stokes method

    NASA Technical Reports Server (NTRS)

    Malone, J. B.; Swanson, R. C.

    1991-01-01

    The Modified Garabedian McFadden (MGM) design procedure was incorporated into an existing 2-D multigrid Navier-Stokes airfoil analysis method. The resulting design method is an iterative procedure based on a residual correction algorithm and permits the automated design of airfoil sections with prescribed surface pressure distributions. The new design method, Multigrid Modified Garabedian McFadden (MG-MGM), is demonstrated for several different transonic pressure distributions obtained from both symmetric and cambered airfoil shapes. The airfoil profiles generated with the MG-MGM code are compared to the original configurations to assess the capabilities of the inverse design method.

  17. Effects of wing height on low-speed aerodynamic characteristics of a model having a 42 deg swept wing, a supercritical airfoil, double-slotted flaps, and a low tail

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.; Sleeman, W. C., Jr.

    1973-01-01

    A low speed investigation was conducted in the Langley V/STOL tunnel to determine the static longitudinal lateral stability characteristics of a general research model with the wing in a high position and a low position on the fuselage. The model had a wing with a quarter chord sweep of 42 deg, an aspect ratio of 6.78, a supercritical airfoil, and a high lift system which consisted of a leading edge slat and a double slotted flap. Various slat and flap deflections represented clean, take off, and landing configurations. A 45 deg swept horizontal tail located slightly below the fuselage center line was investigated with both the low and high wing configurations.

  18. Wind-tunnel test results of airfoil modifications for the EA-6B

    NASA Technical Reports Server (NTRS)

    Sewall, W. G.; Mcghee, R. J.; Ferris, J. C.

    1987-01-01

    Wind-tunnel tests have been conducted (to determine the effects on airfoil performance for several airfoil modifications) for the EA-6B Wing Improvement Program. The modifications consist of contour changes to the leading-edge slat and trailing-edge flap to provide a higher low-speed maximum lift with no high-speed cruise-drag penalty. Airfoil sections from the 28- and 76-percent span stations were selected as baseline shapes with the major testing devoted to the inboard airfoil section (28-percent span station). The airfoil modifications increased the low-speed maximum lift coefficient between 20 and 35 percent over test conditions of 3 to 14 million chord Reynolds number and 0.14 to 0.34 Mach number. At the high-speed test conditions of 0.4 to 0.80 Mach number and 10 million chord Reynolds number, the modified airfoils had either matched or had lower drag coefficients for all normal-force coefficients above 0.2 as compared to the baseline airfoil. At normal-force coefficients less than 0.2, the baseline (original) airfoil had lower drag coefficients than any of the modified airfoils.

  19. Lift enhancing tabs for airfoils

    NASA Technical Reports Server (NTRS)

    Ross, James C. (Inventor)

    1994-01-01

    A tab deployable from the trailing edge of a main airfoil element forces flow onto a following airfoil element, such as a flap, to keep the flow attached and thus enhance lift. For aircraft wings with high lift systems that include leading edge slats, the slats may also be provided with tabs to turn the flow onto the following main element.

  20. An Experimental Study on Active Flow Control Using Synthetic Jet Actuators over S809 Airfoil

    NASA Astrophysics Data System (ADS)

    Gul, M.; Uzol, O.; Akmandor, I. S.

    2014-06-01

    This study investigates the effect of periodic excitation from individually controlled synthetic jet actuators on the dynamics of the flow within the separation and re-attachment regions of the boundary layer over the suction surface of a 2D model wing that has S809 airfoil profile. Experiments are performed in METUWIND's C3 open-loop suction type wind tunnel that has a 1 m × 1 m cross-section test section. The synthetic jet array on the wing consists of three individually controlled actuators driven by piezoelectric diaphragms located at 28% chord location near the mid-span of the wing. In the first part of the study, surface pressure, Constant Temperature Anemometry (CTA) and Particle Image Velocimetry (PIV) measurements are performed over the suction surface of the airfoil to determine the size and characteristics of the separated shear layer and the re-attachment region, i.e. the laminar separation bubble, at 2.3x105 Reynolds number at zero angle of attack and with no flow control as a baseline case. For the controlled case, CTA measurements are carried out under the same inlet conditions at various streamwise locations along the suction surface of the airfoil to investigate the effect of the synthetic jet on the boundary layer properties. During the controlled case experiments, the synthetic jet actuators are driven with a sinusoidal frequency of 1.45 kHz and 300Vp-p. Results of this study show that periodic excitation from the synthetic jet actuators eliminates the laminar separation bubble formed over the suction surface of the airfoil at 2.3x105 Reynolds number at zero angle of attack.

  1. Full-scale Force and Pressure-distribution Tests on a Tapered U.S.A. 45 Airfoil

    NASA Technical Reports Server (NTRS)

    Parsons, John F

    1935-01-01

    This report presents the results of force and pressure-distribution tests on a 2:1 tapered USA 45 airfoil as determined in the full-scale wind tunnel. The airfoil has a constant-chord center section and rounded tips and is tapered in thickness from 18 percent at the root to 9 percent at the tip. Force tests were made throughout a Reynolds Number range of approximately 2,000,000 to 8,000,000 providing data on the scale effect in addition to the conventional characteristics. Pressure-distribution data were obtained from tests at a Reynolds Number of approximately 4,000,000. The aerodynamic characteristics given by the usual dimensionless coefficients are presented graphically.

  2. Low Reynolds number airfoil survey, volume 1

    NASA Technical Reports Server (NTRS)

    Carmichael, B. H.

    1981-01-01

    The differences in flow behavior two dimensional airfoils in the critical chordlength Reynolds number compared with lower and higher Reynolds number are discussed. The large laminar separation bubble is discussed in view of its important influence on critical Reynolds number airfoil behavior. The shortcomings of application of theoretical boundary layer computations which are successful at higher Reynolds numbers to the critical regime are discussed. The large variation in experimental aerodynamic characteristic measurement due to small changes in ambient turbulence, vibration, and sound level is illustrated. The difficulties in obtaining accurate detailed measurements in free flight and dramatic performance improvements at critical Reynolds number, achieved with various types of boundary layer tripping devices are discussed.

  3. Turbine airfoil to shround attachment

    DOEpatents

    Campbell, Christian X; Morrison, Jay A; James, Allister W; Snider, Raymond G; Eshak, Daniel M; Marra, John J; Wessell, Brian J

    2014-05-06

    A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.

  4. On the general theory of thin airfoils for nonuniform motion

    NASA Technical Reports Server (NTRS)

    Reissner, Eric

    1944-01-01

    General thin-airfoil theory for a compressible fluid is formulated as boundary problem for the velocity potential, without recourse to the theory of vortex motion. On the basis of this formulation the integral equation of lifting-surface theory for an incompressible fluid is derived with the chordwise component of the fluid velocity at the airfoil as the function to be determined. It is shown how by integration by parts this integral equation can be transformed into the Biot-Savart theorem. A clarification is gained regarding the use of principal value definitions for the integral which occur. The integral equation of lifting-surface theory is used a s the starting point for the establishment of a theory for the nonstationary airfoil which is a generalization of lifting-line theory for the stationary airfoil and which might be called "lifting-strip" theory. Explicit expressions are given for section lift and section moment in terms of the circulation function, which for any given wing deflection is to be determined from an integral equation which is of the type of the equation of lifting-line theory. The results obtained are for airfoils of uniform chord. They can be extended to tapered airfoils. One of the main uses of the results should be that they furnish a practical means for the analysis of the aerodynamic span effect in the problem of wing flutter. The range of applicability of "lifting-strip" theory is the same as that of lifting-line theory so that its results may be applied to airfoils with aspect ratios as low as three.

  5. Pressure Distribution Over Airfoils at High Speeds

    NASA Technical Reports Server (NTRS)

    Briggs, L J; Dryden, H L

    1927-01-01

    This report deals with the pressure distribution over airfoils at high speeds, and describes an extension of an investigation of the aerodynamic characteristics of certain airfoils which was presented in NACA Technical Report no. 207. The results presented in report no. 207 have been confirmed and extended to higher speeds through a more extensive and systematic series of tests. Observations were also made of the air flow near the surface of the airfoils, and the large changes in lift coefficients were shown to be associated with a sudden breaking away of the flow from the upper surface. The tests were made on models of 1-inch chord and comparison with the earlier measurements on models of 3-inch chord shows that the sudden change in the lift coefficient is due to compressibility and not to a change in the Reynolds number. The Reynolds number still has a large effect, however, on the drag coefficient. The pressure distribution observations furnish the propeller designer with data on the load distribution at high speeds, and also give a better picture of the air-flow changes.

  6. Ordered roughness effects on NACA 0026 airfoil

    NASA Astrophysics Data System (ADS)

    Harun, Z.; Abbas, A. A.; Dheyaa, R. Mohammed; Ghazali, M. I.

    2016-10-01

    The effects of highly-ordered rough surface - riblets, applied onto the surface of a NACA 0026 airfoil, are investigated experimentally using wind tunnel. The riblets are arranged in directionally converging - diverging pattern with dimensions of height, h = 1 mm, pitch or spacing, s = 1 mm, yaw angle α = 0o and 10o The airfoil with external geometry of 500 mm span, 600 mm chord and 156 mm thickness has been built using mostly woods and aluminium. Turbulence quantities are collected using hotwire anemometry. Hotwire measurements show that flows past converging and diverging pattern inherit similar patterns in the near-wall region for both mean velocity and turbulence intensities profiles. The mean velocity profiles in logarithmic regions for both flows past converging and diverging riblet pattern are lower than that with yaw angle α = 0o. Converging riblets cause the boundary layer to thicken and the flow with yaw angle α = 0o produces the thinnest boundary layer. Both the converging and diverging riblets cause pronounced outer peaks in the turbulence intensities profiles. Most importantly, flows past converging and diverging pattern experience 30% skin friction reductions. Higher order statistics show that riblet surfaces produce similar effects due to adverse pressure gradient. It is concluded that a small strip of different ordered roughness features applied at a leading edge of an airfoil can change the turbulence characteristics dramatically.

  7. Airfoil nozzle and shroud assembly

    DOEpatents

    Shaffer, J.E.; Norton, P.F.

    1997-06-03

    An airfoil and nozzle assembly are disclosed including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached. 5 figs.

  8. Airfoil nozzle and shroud assembly

    DOEpatents

    Shaffer, James E.; Norton, Paul F.

    1997-01-01

    An airfoil and nozzle assembly including an outer shroud having a plurality of vane members attached to an inner surface and having a cantilevered end. The assembly further includes a inner shroud being formed by a plurality of segments. Each of the segments having a first end and a second end and having a recess positioned in each of the ends. The cantilevered end of the vane member being positioned in the recess. The airfoil and nozzle assembly being made from a material having a lower rate of thermal expansion than that of the components to which the airfoil and nozzle assembly is attached.

  9. An Iterative Decambering Approach for Post-Stall Prediction of Wing Characteristics using known Section Data

    NASA Technical Reports Server (NTRS)

    Mukherjee, Rinku; Gopalarathnam, Ashok; Kim, Sung Wan

    2003-01-01

    An iterative decambering approach for the post stall prediction of wings using known section data as inputs is presented. The method can currently be used for incompressible .ow and can be extended to compressible subsonic .ow using Mach number correction schemes. A detailed discussion of past work on this topic is presented first. Next, an overview of the decambering approach is presented and is illustrated by applying the approach to the prediction of the two-dimensional C(sub l) and C(sub m) curves for an airfoil. The implementation of the approach for iterative decambering of wing sections is then discussed. A novel feature of the current e.ort is the use of a multidimensional Newton iteration for taking into consideration the coupling between the di.erent sections of the wing. The approach lends itself to implementation in a variety of finite-wing analysis methods such as lifting-line theory, discrete-vortex Weissinger's method, and vortex lattice codes. Results are presented for a rectangular wing for a from 0 to 25 deg. The results are compared for both increasing and decreasing directions of a, and they show that a hysteresis loop can be predicted for post-stall angles of attack.

  10. Thick airfoil designs for the root of the 10MW INNWIND.EU wind turbine

    NASA Astrophysics Data System (ADS)

    Mu≁oz, A.; Méndez, B.; Munduate, X.

    2016-09-01

    The main objective of the “INNWIND.EU” project is to investigate and demonstrate innovative designs for 10-20MW offshore wind turbines and their key components, such as lightweight rotors. In this context, the present paper describes the development of two new airfoils for the blade root region. From the structural point of view, the root is the region in charge of transmitting all the loads of the blade to the hub. Thus, it is very important to include airfoils with adequate structural properties in this region. The present article makes use of high-thickness and blunt trailing edge airfoils to improve the structural characteristics of the airfoils used to build this blade region. CENER's (National Renewable Energy Center of Spain) airfoil design tool uses the airfoil software XFOIL to compute the aerodynamic characteristics of the designed airfoils. That software is based on panel methods which show some problems with the calculation of airfoils with thickness bigger than 35% and with blunt trailing edge. This drawback has been overcome with the development of an empirical correction for XFOIL lift and drag prediction based on airfoil experiments. From the aerodynamic point of view, thick airfoils are known to be very sensitive to surface contamination or turbulent inflow conditions. Consequently, the design optimization takes into account the aerodynamic torque in both clean and contaminated conditions. Two airfoils have been designed aiming to improve the structural and the aerodynamic behaviour of the blade in clean and contaminated conditions. This improvement has been corroborated with Blade Element Momentum (BEM) computations.

  11. Initial Circulation and Peak Vorticity Behavior of Vortices Shed from Airfoil Vortex Generators

    NASA Technical Reports Server (NTRS)

    Wendt, Bruce J.; Biesiadny, Tom (Technical Monitor)

    2001-01-01

    An extensive parametric study of vortices shed from airfoil vortex generators has been conducted to determine the dependence of initial vortex circulation and peak vorticity on elements of the airfoil geometry and impinging flow conditions. These elements include the airfoil angle of attack, chord length, span, aspect ratio, local boundary layer thickness, and free stream Mach number. In addition, the influence of airfoil-to-airfoil spacing on the circulation and peak vorticity has been examined for pairs of co-rotating and counter-rotating vortices. The vortex generators were symmetric airfoils having a NACA-0012 cross-sectional profile. These airfoils were mounted either in isolation, or in pairs, on the surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio was about 17 percent. The circulation and peak vorticity data were derived from cross-plane velocity measurements acquired with a seven-hole probe at one chord-length downstream of the airfoil trailing edge location. The circulation is observed to be proportional to the free-stream Mach number, the angle-of-attack, and the span-to-boundary layer thickness ratio. With these parameters held constant, the circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio. The peak vorticity is also observed to be proportional to the free-stream Mach number, the airfoil angle-of-attack, and the span-to-boundary layer thickness ratio. Unlike circulation, however, the peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at an aspect ratio of about 2.0 before falling off again at higher values of aspect ratio. Co-rotating vortices shed from closely spaced pairs of airfoils have values of circulation and peak vorticity under those values found for vortices shed from isolated airfoils of the same geometry. Conversely, counter-rotating vortices show enhanced values of circulation and peak vorticity when compared to values

  12. Effect of aileron displacement on wing characteristics

    NASA Technical Reports Server (NTRS)

    Heald, R H

    1933-01-01

    The effect of aileron displacement on wing characteristics has been investigated for the Clark Y and the U.S.A. 27 wing sections equipped with rectangular ailerons. The airfoils, rectangular in plan, and having a 10 inch chord and 60 inch span, were mounted on a model fuselage.

  13. Wind-tunnel tests on combinations of a wing with fixed auxiliary airfoils having various chords and profiles

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Sanders, Robert

    1934-01-01

    This report presents the results of wind tunnel tests on various auxiliary airfoils having three different airfoil sections and several different chord lengths in combination with a Clark y model wing in a sufficient number of relative positions to determine the optimum with regard to certain criterions of aerodynamic performance. The airfoil sections included a symmetrical profile, one of medium camber, and a highly cambered one. The chord sizes of the auxiliary airfoils ranged from 7.5 to 25 percent of the chord of the main wing, and the span was equal to that of the main wing.

  14. Aerodynamic performance of transonic and subsonic airfoils: Effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape

    NASA Astrophysics Data System (ADS)

    Zhang, Qiang

    The effects of surface roughness, turbulence intensity, Mach number, and streamline curvature-airfoil shape on the aerodynamic performance of turbine airfoils are investigated in compressible, high speed flows. The University of Utah Transonic Wind Tunnel is employed for the experimental part of the study. Two different test sections are designed to produce Mach numbers, Reynolds numbers, passage mass flow rates, and physical dimensions, which match values along turbine blades in operating engines: (i) a nonturning test section with a symmetric airfoil, and (ii) a cascade test section with a cambered turbine vane. The nonuniform, irregular, three-dimensional surface roughness is characterized using the equivalent sand grain roughness size. Changing the airfoil surface roughness condition has a substantial effect on wake profiles of total pressure loss coefficients, normalized Mach number, normalized kinetic energy, and on the normalized and dimensional magnitudes of Integrated Aerodynamic Losses produced by the airfoils. Comparisons with results for a symmetric airfoil and a cambered vane show that roughness has more substantial effects on losses produced by the symmetric airfoil than the cambered vane. Data are also provided that illustrate the larger loss magnitudes are generally present with flow turning and cambered airfoils, than with symmetric airfoils. Wake turbulence structure of symmetric airfoils and cambered vanes are also studied experimentally. The effects of surface roughness and freestream turbulence levels on wake distributions of mean velocity, turbulence intensity, and power spectral density profiles and vortex shedding frequencies are quantified one axial chord length downstream of the test airfoils. As the level of surface roughness increases, all wake profile quantities broaden significantly and nondimensional vortex shedding frequencies decrease. Wake profiles produced by the symmetric airfoil are more sensitive to variations of surface

  15. Nozzle airfoil having movable nozzle ribs

    DOEpatents

    Yu, Yufeng Phillip; Itzel, Gary Michael

    2002-01-01

    A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

  16. Numerical design of shockless airfoils

    NASA Technical Reports Server (NTRS)

    Garabedian, P. R.

    1979-01-01

    An attempt is made to indicate and briefly discuss only the most significant achievements of the research. The most successful contribution from the contract was the code for two dimensional analysis of airfoils in transonic flow.

  17. Numerical and Analytical Study of Nonlinear Effects of Transonic Flow Past a Wing Airfoil in Oscillation of its Surface Element

    NASA Astrophysics Data System (ADS)

    Aul'chenko, S. M.; Zamuraev, V. P.; Kalinina, A. P.

    2014-05-01

    The present work is devoted to a criterial analysis and mathematical modeling of the influence of forced oscillations of surface elements of a wing airfoil on the shock-wave structure of transonic flow past it. Parameters that govern the regimes of interaction of the oscillatory motion of airfoil sections with the breakdown compression shock have been established. The qualitative and quantitative influence of these parameters on the wave resistance of the airfoil has been investigated.

  18. Dynamic Stall on Advanced Airfoil Sections,

    DTIC Science & Technology

    1980-05-01

    boundary-layer transition, flow Fromme and Golberg " have indicated that 80-1-2 unsteady wall corrections can be greater A limited amount of static and...A., and Golberg , M. A. New Rotor Profile on the Basis "Unsteady Two-Dimensional Air- of Flow Phenomena; Aerofoil loads Acting on Oscillating

  19. The NASA Langley Laminar-Flow-Control (LFC) experiment on a swept, supercritical airfoil: Design overview

    NASA Technical Reports Server (NTRS)

    Harris, Charles D.; Harvey, William D.; Brooks, Cuyler W., Jr.

    1988-01-01

    A large-chord, swept, supercritical, laminar-flow-control (LFC) airfoil was designed and constructed and is currently undergoing tests in the Langley 8 ft Transonic Pressure Tunnel. The experiment was directed toward evaluating the compatibility of LFC and supercritical airfoils, validating prediction techniques, and generating a data base for future transport airfoil design as part of NASA's ongoing research program to significantly reduce drag and increase aircraft efficiency. Unique features of the airfoil included a high design Mach number with shock free flow and boundary layer control by suction. Special requirements for the experiment included modifications to the wind tunnel to achieve the necessary flow quality and contouring of the test section walls to simulate free air flow about a swept model at transonic speeds. Design of the airfoil with a slotted suction surface, the suction system, and modifications to the tunnel to meet test requirements are discussed.

  20. Summary of Section Data on Trailing-Edge High-Lift Devices

    NASA Technical Reports Server (NTRS)

    1948-01-01

    A summary has been made of available data on the characteristics of airfoil sections with trailing-edge high-lift devices. Data for plain, split, and slotted flaps are collected and analyzed. The effects of each of the variables involved in the design of the various types of flap are examined and, in cases where sufficient data are given, optimum configurations are deduced. Wherever possible, the effects of airfoil section, Reynolds number, and leading-edge roughness are shown. For single and double slotted flaps, where a great mass of unrelated date are available, maximum lift coefficients of a large number of configurations are presented in tables.

  1. Interference on an airfoil of finite span in an open wind tunnel

    NASA Technical Reports Server (NTRS)

    Theodorsen, Theodore

    1934-01-01

    The wall interference on an airfoil of finite span in an open-throat rectangular section has been treated theoretically and the result is presented in a convenient formula. Numerical results are given in tables and diagrams.

  2. Cyberdrugs: a cross-sectional study of online pharmacies characteristics.

    PubMed

    Orizio, Grazia; Schulz, Peter; Domenighini, Serena; Caimi, Luigi; Rosati, Cristina; Rubinelli, Sara; Gelatti, Umberto

    2009-08-01

    As e-commerce and online pharmacies (OPs) arose, the potential impact of the Internet on the world of health shifted from merely the spread of information to a real opportunity to acquire health services directly. Aim of the study was to investigate the offer of prescription drugs in OPs, analysing their characteristics, using the content analysis method. The research performed using the Google search engine led to an analysis of 118 online pharmacies. Only 51 (43.2%) of them stated their precise location. Ninety-six (81.4%) online pharmacies did not require a medical prescription from the customer's physician. Online pharmacies rise complex issues in terms of patient-doctor relationship, consumer empowerment, drug quality, regulation and public health implications.

  3. Analysis of high Reynolds numbers effects on a wind turbine airfoil using 2D wind tunnel test data

    NASA Astrophysics Data System (ADS)

    Pires, O.; Munduate, X.; Ceyhan, O.; Jacobs, M.; Snel, H.

    2016-09-01

    The aerodynamic behaviour of a wind turbine airfoil has been measured in a dedicated 2D wind tunnel test at the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. The tests have been performed on the DU00W212 airfoil at different Reynolds numbers: 3, 6, 9, 12 and 15 million, and at low Mach numbers (below 0.1). Both clean and tripped conditions of the airfoil have been measured. An analysis of the impact of a wide Reynolds number variation over the aerodynamic characteristics of this airfoil has been performed.

  4. Airfoil Vibration Dampers program

    NASA Technical Reports Server (NTRS)

    Cook, Robert M.

    1991-01-01

    The Airfoil Vibration Damper program has consisted of an analysis phase and a testing phase. During the analysis phase, a state-of-the-art computer code was developed, which can be used to guide designers in the placement and sizing of friction dampers. The use of this computer code was demonstrated by performing representative analyses on turbine blades from the High Pressure Oxidizer Turbopump (HPOTP) and High Pressure Fuel Turbopump (HPFTP) of the Space Shuttle Main Engine (SSME). The testing phase of the program consisted of performing friction damping tests on two different cantilever beams. Data from these tests provided an empirical check on the accuracy of the computer code developed in the analysis phase. Results of the analysis and testing showed that the computer code can accurately predict the performance of friction dampers. In addition, a valuable set of friction damping data was generated, which can be used to aid in the design of friction dampers, as well as provide benchmark test cases for future code developers.

  5. A supercritical airfoil experiment

    NASA Technical Reports Server (NTRS)

    Mateer, G. G.; Seegmiller, H. L.; Hand, L. A.; Szodruck, J.

    1994-01-01

    The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference connections can be made to the data sets.

  6. Airfoil deposition model

    NASA Technical Reports Server (NTRS)

    Kohl, F. J.

    1982-01-01

    The methodology to predict deposit evolution (deposition rate and subsequent flow of liquid deposits) as a function of fuel and air impurity content and relevant aerodynamic parameters for turbine airfoils is developed in this research. The spectrum of deposition conditions encountered in gas turbine operations includes the mechanisms of vapor deposition, small particle deposition with thermophoresis, and larger particle deposition with inertial effects. The focus is on using a simplified version of the comprehensive multicomponent vapor diffusion formalism to make deposition predictions for: (1) simple geometry collectors; and (2) gas turbine blade shapes, including both developing laminar and turbulent boundary layers. For the gas turbine blade the insights developed in previous programs are being combined with heat and mass transfer coefficient calculations using the STAN 5 boundary layer code to predict vapor deposition rates and corresponding liquid layer thicknesses on turbine blades. A computer program is being written which utilizes the local values of the calculated deposition rate and skin friction to calculate the increment in liquid condensate layer growth along a collector surface.

  7. Low speed aerodynamic characteristics of a transport model having 42.33 deg swept low wing with supercritical airfoil, double-slotted flaps, and T-tail or low tail

    NASA Technical Reports Server (NTRS)

    Fournier, P. G.

    1975-01-01

    A low-speed investigation was conducted in the Langley V/STOL tunnel over an angle-of-attack range of approximately 4 deg to 24 deg to determine the static longitudinal stability characteristics and high lift performance of a general research model which represented an advanced subsonic transport configuration. The model had a 42.33 deg swept, aspect ratio 7.05 wing with a supercritical airfoil and high lift system consisting of a leading edge device (slat or Kruger flap) and a double-slotted flap. The flaps were deflected for take off and landing configurations and were not deflected for tests of the clean configuration. The model was tested with the horizontal tail in either a T tail or low tail position. The effects of various arrangements of flowthrough nacelles which represent a three engine configuration (two large wing-mounted nacelles and a vertical tail mounted nacelle) and a four engine configuration (four smaller wing-mounted nacelles) were determined.

  8. A Numerical Evaluation of Icing Effects on a Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

    Chung, James J.; Addy, Harold E., Jr.

    2000-01-01

    As a part of CFD code validation efforts within the Icing Branch of NASA Glenn Research Center, computations were performed for natural laminar flow (NLF) airfoil, NLF-0414. with 6 and 22.5 minute ice accretions. Both 3-D ice castings and 2-D machine-generated ice shapes were used in wind tunnel tests to study the effects of natural ice is well as simulated ice. They were mounted in the test section of the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley that the 2-dimensionality of the flow can be maintained. Aerodynamic properties predicted by computations were compared to data obtained through the experiment by the authors at the LTPT. Computations were performed only in 2-D and in the case of 3-D ice, the digitized ice shape obtained at one spanwise location was used. The comparisons were mainly concentrated on the lift characteristics over Reynolds numbers ranging from 3 to 10 million and Mach numbers ranging from 0.12 to 0.29. WIND code computations indicated that the predicted stall angles were in agreement with experiment within one or two degrees. The maximum lift values obtained by computations were in good agreement with those of the experiment for the 6 minute ice shapes and the minute 3-D ice, but were somewhat lower in the case of the 22.5 minute 2-D ice. In general, the Reynolds number variation did not cause much change in the lift values while the variation of Mach number showed more change in the lift. The Spalart-Allmaras (S-A) turbulence model was the best performing model for the airfoil with the 22.5 minute ice and the Shear Stress Turbulence (SST) turbulence model was the best for the airfoil with the 6 minute ice and also for the clean airfoil. The pressure distribution on the surface of the iced airfoil showed good agreement for the 6 minute ice. However, relatively poor agreement of the pressure distribution on the upper surface aft of the leading edge horn for the 22.5 minute ice suggests that improvements are needed in the grid or

  9. Root region airfoil for wind turbine

    DOEpatents

    Tangler, James L.; Somers, Dan M.

    1995-01-01

    A thick airfoil for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%-26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4-1.6 that has minimum sensitivity to roughness effects.

  10. Advanced technology airfoil research, volume 2. [conferences

    NASA Technical Reports Server (NTRS)

    1979-01-01

    A comprehensive review of airfoil research is presented. The major thrust of the research is in three areas: development of computational aerodynamic codes for airfoil analysis and design, development of experimental facilities and test techniques, and all types of airfoil applications.

  11. Airfoil shape for a turbine nozzle

    DOEpatents

    Burdgick, Steven Sebastian; Patik, Joseph Francis; Itzel, Gary Michael

    2002-01-01

    A first-stage nozzle vane includes an airfoil having a profile according to Table I. The annulus profile of the hot gas path is defined in conjunction with the airfoil profile and the profile of the inner and outer walls by the Cartesian coordinate values given in Tables I and II, respectively. The airfoil is a three-dimensional bowed design, both in the airfoil body and in the trailing edge. The airfoil is steam and air-cooled by flowing cooling mediums through cavities extending in the vane between inner and outer walls.

  12. Hook nozzle arrangement for supporting airfoil vanes

    DOEpatents

    Shaffer, James E.; Norton, Paul F.

    1996-01-01

    A gas turbine engine's nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic.

  13. Effect of oscillation frequency on wind turbine airfoil dynamic stall

    NASA Astrophysics Data System (ADS)

    Zhou, Z.; Li, C.; Nie, J. B.; Chen, Y.

    2013-12-01

    At the same oscillation amplitude, Reynolds Number, mean angle of attack, the dynamic stall characteristics of the NREL S809 airfoil undergoing sinusoidal pitch oscillations of different oscillation frequencies were investigated with modified k-ω SST turbulence model of CFD solution for two-dimensional numerical simulation. The predicted lift, drag coefficients and moment coefficients were compared with the Ohio State University wind tunnel test results, which showed a good agreement. The birth, development and breaking off of eddies were analyzed through streamline distribution around airfoil and the influence of oscillation frequencies on dynamic stall characteristics was also described and analyzed in detail, which enrich the database of dynamic stall characteristics needed by the quantization of oscillation frequencies on dynamic characteristics and prove that sliding mesh method is reliable when dealing with dynamic stall problems.

  14. Flutter and Time Response Analyses of Three Degree of Freedom Airfoils in Transonic Flow

    DTIC Science & Technology

    1981-08-01

    Reference 10), and a NACA 64A010 (Reference 4) airfoil. They also used STRANS2 and UTRANS2 to analyze a TF-8A wing section (Reference 4). In the...used for the reduced frequency kc with values up to 1.0 and the entire Mach number range. He used the code to study the NACA 64A010 airfoil for two...structural equation of motion. Rizzetta (Reference 1G) performed a time-response analysis of a NACA 64A010 airfoil with a single pitch d.o.f, and

  15. Pneumatic Spoiler Controls Airfoil Lift

    NASA Technical Reports Server (NTRS)

    Hunter, D.; Krauss, T.

    1991-01-01

    Air ejection from leading edge of airfoil used for controlled decrease of lift. Pneumatic-spoiler principle developed for equalizing lift on helicopter rotor blades. Also used to enhance aerodynamic control of short-fuselage or rudderless aircraft such as "flying-wing" airplanes. Leading-edge injection increases maneuverability of such high-performance fixed-wing aircraft as fighters.

  16. Tests in the variable-density wind tunnel of the NACA 23012 airfoil with plain and split flaps

    NASA Technical Reports Server (NTRS)

    Abbott, Ira H; Greenberg, Harry

    1939-01-01

    Section characteristics for use in wing design are presented for the NACA 23012 airfoil with plain and split flaps of 20 percent wing chord at a value of the effective Reynolds number of about 8,000,000. The flap deflections covered a range from 60 degrees upward to 75 degrees downward for the plain flap and from neutral to 90 degrees downward for the split flap. The split flap was aerodynamically superior to the plain flap in producing high maximum lift coefficients and in having lower profile-drag coefficients at high lift coefficients.

  17. The modelling of symmetric airfoil vortex generators

    NASA Technical Reports Server (NTRS)

    Reichert, B. A.; Wendt, B. J.

    1996-01-01

    An experimental study is conducted to determine the dependence of vortex generator geometry and impinging flow conditions on shed vortex circulation and crossplane peak vorticity for one type of vortex generator. The vortex generator is a symmetric airfoil having a NACA 0012 cross-sectional profile. The geometry and flow parameters varied include angle-of-attack alfa, chordlength c, span h, and Mach number M. The vortex generators are mounted either in isolation or in a symmetric counter-rotating array configuration on the inside surface of a straight pipe. The turbulent boundary layer thickness to pipe radius ratio is delta/R = 0. 17. Circulation and peak vorticity data are derived from crossplane velocity measurements conducted at or about 1 chord downstream of the vortex generator trailing edge. Shed vortex circulation is observed to be proportional to M, alfa, and h/delta. With these parameters held constant, circulation is observed to fall off in monotonic fashion with increasing airfoil aspect ratio AR. Shed vortex peak vorticity is also observed to be proportional to M, alfa, and h/delta. Unlike circulation, however, peak vorticity is observed to increase with increasing aspect ratio, reaching a peak value at AR approx. 2.0 before falling off.

  18. Transition and separation control on a low-Reynolds number airfoil

    NASA Technical Reports Server (NTRS)

    Mangalam, S. M.; Bar-Sever, A.; Zaman, K. B. M. Q.; Harvey, W. D.

    1986-01-01

    The major problem associated with the aerodynamic performance of airfoils at low Reynolds numbers is the presence of extensive laminar boundary-layer separation resulting in a large increase in presssure drag and a decrease in lift. The rapid deterioration in airfoil characteristics can be largely eliminated by artificially controlling the flow through the introduction of suitable disturbances in the boundary layer such that transition occurs ahead of the anticipated laminar separation. This paper presents the results of wind-tunnel tests conducted on a 10-cm model of LRN (1)-1007 airfoil with passive (roughness trips) and active (acoustic excitation) controls to trigger transition and suppress separation. Significant improvements in the aerodynamic characteristics of the airfoil were observed. Results of this study for a chord Reynolds number range of 40,000 to 250,000 are presented in this paper.

  19. SiC/SiC Leading Edge Turbine Airfoil Tested Under Simulated Gas Turbine Conditions

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Hatton, Kenneth S.

    1999-01-01

    Silicon-based ceramics have been proposed as component materials for use in gas turbine engine hot-sections. A high pressure burner rig was used to expose both a baseline metal airfoil and ceramic matrix composite leading edge airfoil to typical gas turbine conditions to comparatively evaluate the material response at high temperatures. To eliminate many of the concerns related to an entirely ceramic, rotating airfoil, this study has focused on equipping a stationary metal airfoil with a ceramic leading edge insert to demonstrate the feasibility and benefits of such a configuration. Here, the idea was to allow the SiC/SiC composite to be integrated as the airfoil's leading edge, operating in a "free-floating" or unrestrained manner. and provide temperature relief to the metal blade underneath. The test included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were air-cooled, uniquely instrumented, and exposed to the same internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). Results show the leading edge insert remained structurally intact after 200 simulated flight cycles with only a slightly oxidized surface. The instrumentation clearly suggested a significant reduction (approximately 600 F) in internal metal temperatures as a result of the ceramic leading edge. The object of this testing was to validate the design and analysis done by Materials Research and Design of Rosemont, PA and to determine the feasibility of this design for the intended application.

  20. Some observations of surface pressures and the near wake of a blunt trailing edge airfoil

    NASA Technical Reports Server (NTRS)

    Digumarthi, R. V.; Koutsoyannis, S. P.; Karamcheti, K.

    1981-01-01

    Experiments with a truncated and untruncated airfoils of profiles NACA 640A10, were carried out in subsonic wind tunnels in a velocity range of 19m/s to 54m/s corresponding to Reynolds numbers of 200,000 to 468,000 based on the chord. Airfoil spanned the test section to achieve two dimensionality of the model. Velocity measurements, pressure measurements, and vortex shedding in the wake were measured using a hotwire and pressure transducers. The measured chordwise static pressure distribution on the smooth trailing edge airfoil along the midspan plane, agreed with the theoretical results calculated on the basis of the potential flow for that airfoil. Boundary layer profiles measured in the midspan plane, behind the maximum thickness of the airfoil show no separation of the flow. Spanwise distribution of the measured static pressure on the upper surface of the airfoil shows uniformity for both configurations with and without the boundary layer trip. This uniformity of pressure distribution and separation indicates that the flow on the airfoil was uniform and two dimensional in character.

  1. A Two Element Laminar Flow Airfoil Optimized for Cruise. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Steen, Gregory Glen

    1994-01-01

    Numerical and experimental results are presented for a new two-element, fixed-geometry natural laminar flow airfoil optimized for cruise Reynolds numbers on the order of three million. The airfoil design consists of a primary element and an independent secondary element with a primary to secondary chord ratio of three to one. The airfoil was designed to improve the cruise lift-to-drag ratio while maintaining an appropriate landing capability when compared to conventional airfoils. The airfoil was numerically developed utilizing the NASA Langley Multi-Component Airfoil Analysis computer code running on a personal computer. Numerical results show a nearly 11.75 percent decrease in overall wing drag with no increase in stall speed at sailplane cruise conditions when compared to a wing based on an efficient single element airfoil. Section surface pressure, wake survey, transition location, and flow visualization results were obtained in the Texas A&M University Low Speed Wind Tunnel. Comparisons between the numerical and experimental data, the effects of the relative position and angle of the two elements, and Reynolds number variations from 8 x 10(exp 5) to 3 x 10(exp 6) for the optimum geometry case are presented.

  2. Family of airfoil shapes for rotating blades. [for increased power efficiency and blade stability

    NASA Technical Reports Server (NTRS)

    Noonan, K. W. (Inventor)

    1983-01-01

    An airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is described. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.

  3. Experimental Investigation on Airfoil Shock Control by Plasma Aerodynamic Actuation

    NASA Astrophysics Data System (ADS)

    Sun, Quan; Cheng, Bangqin; Li, Yinghong; Cui, Wei; Jin, Di; Li, Jun

    2013-11-01

    An experimental investigation on airfoil (NACA64—215) shock control is performed by plasma aerodynamic actuation in a supersonic tunnel (Ma = 2). The results of schlieren and pressure measurement show that when plasma aerodynamic actuation is applied, the position moves forward and the intensity of shock at the head of the airfoil weakens. With the increase in actuating voltage, the total pressure measured at the head of the airfoil increases, which means that the shock intensity decreases and the control effect increases. The best actuation effect is caused by upwind-direction actuation with a magnetic field, and then downwind-direction actuation with a magnetic field, while the control effect of aerodynamic actuation without a magnetic field is the most inconspicuous. The mean intensity of the normal shock at the head of the airfoil is relatively decreased by 16.33%, and the normal shock intensity is relatively reduced by 27.5% when 1000 V actuating voltage and upwind-direction actuation are applied with a magnetic field. This paper theoretically analyzes the Joule heating effect generated by DC discharge and the Lorentz force effect caused by the magnetic field. The discharge characteristics are compared for all kinds of actuation conditions to reveal the mechanism of shock control by plasma aerodynamic actuation.

  4. Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System

    NASA Technical Reports Server (NTRS)

    Gowan, William H., Jr.; Mulholland, Donald R.

    1951-01-01

    Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.

  5. Force and pressure tests of the GA(W)-1 airfoil with a 20% aileron and pressure tests with a 30% Fowler flap

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.; Seetharam, H. C.; Fiscko, K. A.

    1977-01-01

    Wind tunnel force and pressure tests were conducted for the GA(W)-1 airfoil equipped with a 20% aileron, and pressure tests were conducted with a 30% Fowler flap. All tests were conducted at a Reynolds number of 2.2 and a Mach number of 0.13. The aileron provides control effectiveness similar to ailerons applied to more conventional airfoils. Effects of aileron gaps from 0% to 2% chord were evaluated, as well as hinge moment characteristics. The aft camber of the GA(W)-1 section results in a substantial up-aileron moment, but the hinge moments associated with aileron deflection are similar to other configurations. Fowler flap pressure distributions indicate that unseparated flow is achieved for flap settings up to 40 deg., over a limited angle of attack range. Theoretical pressure distributions compare favorably with experiments for low flap deflections, but show substantial errors at large deflections.

  6. OUT Success Stories: Advanced Airfoils for Wind Turbines

    DOE R&D Accomplishments Database

    Jones, J.; Green, B.

    2000-08-01

    New airfoils have substantially increased the aerodynamic efficiency of wind turbines. It is clear that these new airfoils substantially increased energy output from wind turbines. Virtually all new blades built in this country today use these advanced airfoil designs.

  7. Root region airfoil for wind turbine

    DOEpatents

    Tangler, J.L.; Somers, D.M.

    1995-05-23

    A thick airfoil is described for the root region of the blade of a wind turbine. The airfoil has a thickness in a range from 24%--26% and a Reynolds number in a range from 1,000,000 to 1,800,000. The airfoil has a maximum lift coefficient of 1.4--1.6 that has minimum sensitivity to roughness effects. 3 Figs.

  8. Boundary-layer stability and airfoil design

    NASA Technical Reports Server (NTRS)

    Viken, Jeffrey K.

    1986-01-01

    Several different natural laminar flow (NLF) airfoils have been analyzed for stability of the laminar boundary layer using linear stability codes. The NLF airfoils analyzed come from three different design conditions: incompressible; compressible with no sweep; and compressible with sweep. Some of the design problems are discussed, concentrating on those problems associated with keeping the boundary layer laminar. Also, there is a discussion on how a linear stability analysis was effectively used to improve the design for some of the airfoils.

  9. Airfoil seal system for gas turbine engine

    SciTech Connect

    None, None

    2013-06-25

    A turbine airfoil seal system of a turbine engine having a seal base with a plurality of seal strips extending therefrom for sealing gaps between rotational airfoils and adjacent stationary components. The seal strips may overlap each other and may be generally aligned with each other. The seal strips may flex during operation to further reduce the gap between the rotational airfoils and adjacent stationary components.

  10. Adjoint-based airfoil shape optimization in transonic flow

    NASA Astrophysics Data System (ADS)

    Gramanzini, Joe-Ray

    The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.

  11. Blade design trade-offs using low-lift airfoils for stall-regulated HAWTs

    SciTech Connect

    Giguere, P.; Selig, M.S.; Tangler, J.L.

    1999-11-01

    A systematic blade design study was conducted to explore the trade-offs in using low-lift airfoils for a 750-kilowatt stall-regulated wind turbine. Tip-region airfoils having a maximum-lift coefficient ranging from 0.7--1.2 were considered in this study, with the main objective of identifying the practical lower limit for the maximum-life coefficient. Blades were optimized for both maximum annual energy production and minimum cost of energy using a method that takes into account aerodynamic and structural considerations. The results indicate that the effect of the maximum-lift coefficient on the cost of energy is small with a slight advantage to the highest maximum lift coefficient airfoils for the tip-region of the blade become more desirable as machine size increases, provided the airfoils yield acceptable stall characteristics. The conclusions are applicable to large wind turbines that use passive or active stall to regulate peak power.

  12. Bionic Design of Wind Turbine Blade Based on Long-Eared Owl's Airfoil.

    PubMed

    Tian, Weijun; Yang, Zhen; Zhang, Qi; Wang, Jiyue; Li, Ming; Ma, Yi; Cong, Qian

    2017-01-01

    The main purpose of this paper is to demonstrate a bionic design for the airfoil of wind turbines inspired by the morphology of Long-eared Owl's wings. Glauert Model was adopted to design the standard blade and the bionic blade, respectively. Numerical analysis method was utilized to study the aerodynamic characteristics of the airfoils as well as the blades. Results show that the bionic airfoil inspired by the airfoil at the 50% aspect ratio of the Long-eared Owl's wing gives rise to a superior lift coefficient and stalling performance and thus can be beneficial to improving the performance of the wind turbine blade. Also, the efficiency of the bionic blade in wind turbine blades tests increases by 12% or above (up to 44%) compared to that of the standard blade. The reason lies in the bigger pressure difference between the upper and lower surface which can provide stronger lift.

  13. Bionic Design of Wind Turbine Blade Based on Long-Eared Owl's Airfoil

    PubMed Central

    Li, Ming

    2017-01-01

    The main purpose of this paper is to demonstrate a bionic design for the airfoil of wind turbines inspired by the morphology of Long-eared Owl's wings. Glauert Model was adopted to design the standard blade and the bionic blade, respectively. Numerical analysis method was utilized to study the aerodynamic characteristics of the airfoils as well as the blades. Results show that the bionic airfoil inspired by the airfoil at the 50% aspect ratio of the Long-eared Owl's wing gives rise to a superior lift coefficient and stalling performance and thus can be beneficial to improving the performance of the wind turbine blade. Also, the efficiency of the bionic blade in wind turbine blades tests increases by 12% or above (up to 44%) compared to that of the standard blade. The reason lies in the bigger pressure difference between the upper and lower surface which can provide stronger lift. PMID:28243053

  14. Application of shock tubes to transonic airfoil testing at high Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.

    1978-01-01

    Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.

  15. Three-dimensional effects on airfoils

    NASA Technical Reports Server (NTRS)

    Chevallier, J. P.

    1983-01-01

    The effects of boundary layer flows along the walls of wind tunnels were studied to validate the transfer of two dimensional calculations to three dimensional transonic flowfield calculations. Results from trials in various wind tunnels were examind to determine the effects of the wall boundary flow on the control surfaces of an airfoil. Models sliding along a groove in the wall of a channel at sub- and transonic speeds were examined, with the finding that with either nonuniformities in the groove, or even if the channel walls are uniform, the lateral boundary layer can cause variations in the central flow region or alter the onset of shock at the transition point. Models for the effects in both turbulence and in the absence of turbulence are formulated, and it is noted that the characteristics of individual wind tunnels must be studied to quantify any existing three dimensional effects.

  16. Experimental airfoil characterization under tailored turbulent conditions

    NASA Astrophysics Data System (ADS)

    Heißelmann, Hendrik; Peinke, Joachim; Hölling, Michael

    2016-09-01

    Studies of the impact of turbulent inflow conditions on the airfoil characteristics were performed within the EU FP7 project AVATAR. The aim of this study is to provide data for the validation of simulations and the improvement of engineering tools. Chord-wise pressure distributions and highly-resolved force data of the wind turbine dedicated DU 00-W-212 profile were measured in the wind tunnel in two tailored turbulent inflow conditions generated with an active grid. A sinusoidal and an intermittent pattern with customized inflow angle fluctuations were generated providing two significantly different distributions of reduced frequencies. The obtained pressure distributions and polars from the unsteady patterns are compared to the laminar baseline case.

  17. An experimental investigation into cavitation behaviour and pressure characteristics of alternative blade sections for propellers

    NASA Astrophysics Data System (ADS)

    Korkut, Emin; Atlar, Mehmet; Wang, Dazheng

    2013-03-01

    During the final quarter of the last century considerable efforts have been spent to reduce the hull pressure fluctuations caused by unsteady propeller cavitation. This has resulted in further changes in propeller design characteristics including increased skew, tip unloading and introduction of "New Blade Sections" (NBS) designed on the basis of the so-called Eppler code. An experimental study was carried out to investigate flow characteristics of alternative two-dimensional (2-D) blade sections of rectangular planform, one of which was the New Blade Section (NBS) developed in Newcastle University and other was based on the well-known National Advisory Committee for Aeronautics (NACA) section. The experiments comprised the cavitation observations and the measurements of the local velocity distribution around the blade sections by using a 2-D Laser Doppler Anemometry (LDA) system. Analysis of the cavitation tests demonstrated that the two blade sections presented very similar bucket shapes with virtually no width at the bottom but relatively favourable buckets arms at the suction and pressure sides for the NACA section. Similarly, pressure analysis of the sections displayed a slightly larger value for the NBS pressure peak. The comparative overall pressure distributions around the sections suggested that the NBS might be more susceptible to cavitation than the NACA section. This can be closely related to the fundamental shape of the NBS with very fine leading edge. Therefore a further investigation into the modification of the leading edge should be considered to improve the cavitation behaviour of the NBS.

  18. Inverse transonic airfoil design including viscous interaction

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

    A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.

  19. Transonic airfoil flowfield analysis using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1975-01-01

    A numerical technique for analyzing transonic airfoils is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that Cartesian coordinates are used rather than a grid which fits the airfoil, such as the conformal circle-plane or 'sheared parabolic' coordinates which were used previously. Comparison with previous results shows that it is not necessary to match the computational grid to the airfoil surface, and that accurate results can be obtained with a Cartesian grid for lifting supercritical airfoils.

  20. Wavy flow cooling concept for turbine airfoils

    DOEpatents

    Liang, George

    2010-08-31

    An airfoil including an outer wall and a cooling cavity formed therein. The cooling cavity includes a leading edge flow channel located adjacent a leading edge of the airfoil and a trailing edge flow channel located adjacent a trailing edge of the airfoil. Each of the leading edge and trailing edge flow channels define respective first and second flow axes located between pressure and suction sides of the airfoil. A plurality of rib members are located within each of the flow channels, spaced along the flow axes, and alternately extending from opposing sides of the flow channels to define undulating flow paths through the flow channels.

  1. Schooling behavior of heaving flexible airfoils

    NASA Astrophysics Data System (ADS)

    Im, Sunghyuk; Sung, Hyung Jin

    2016-11-01

    The schooling behavior of rigid and flexible NACA0017 airfoils in the heaving motion is experimentally explored in a merry-go-round equipment. The airfoil was attached to the end of a horizontal support bar whose other end was connected to the freely rotating vertical axis. The axis was forced to undergo a sinusoidal motion in the vertical direction to make a pure heaving motion of the airfoils in the frequency range of 0.5 to 5 Hz. The propulsion due to the heaving airfoils is expressed by a horizontally rotating speed of the support bar. This experimental setup is simulating infinite schooling situations of airfoils in an in-phase heaving motion with the streamwise distance d. The ratio of the distance to the chord length d/ c was determined by the number of airfoils (1 <= n <= 8) . The rotational frequency F according to the heaving frequency f was measured with different experimental parameters. The schooling number S = f /(nF), representing the number of heaving oscillations between each airfoil, was introduced to explain the schooling behavior of the airfoils. The effects of the flexibility, d/ c and f on the propulsive performance were examined with the schooling behavior of the airfoils. This work was supported by the Creative Research Initiatives (No. 2016-004749) program of the National Research Foundation of Korea (MSIP).

  2. Design of a family of new advanced airfoils for low wind class turbines

    NASA Astrophysics Data System (ADS)

    Grasso, Francesco

    2014-12-01

    In order to maximize the ratio of energy capture and reduce the cost of energy, the selection of the airfoils to be used along the blade plays a crucial role. Despite the general usage of existing airfoils, more and more, families of airfoils specially tailored for specific applications are developed. The present research is focused on the design of a new family of airfoils to be used for the blade of one megawatt wind turbine working in low wind conditions. A hybrid optimization scheme has been implemented, combining together genetic and gradient based algorithms. Large part of the work is dedicated to present and discuss the requirements that needed to be satisfied in order to have a consistent family of geometries with high efficiency, high lift and good structural characteristics. For each airfoil, these characteristics are presented and compared to the ones of existing airfoils. Finally, the aerodynamic design of a new blade for low wind class turbine is illustrated and compared to a reference shape developed by using existing geometries. Due to higher lift performance, the results show a sensitive saving in chords, wetted area and so in loads in idling position.

  3. Design of a 3 kW wind turbine generator with thin airfoil blades

    SciTech Connect

    Ameku, Kazumasa; Nagai, Baku M.; Roy, Jitendro Nath

    2008-09-15

    Three blades of a 3 kW prototype wind turbine generator were designed with thin airfoil and a tip speed ratio of 3. The wind turbine has been controlled via two control methods: the variable pitch angle and by regulation of the field current of the generator and examined under real wind conditions. The characteristics of the thin airfoil, called ''Seven arcs thin airfoil'' named so because the airfoil is composed of seven circular arcs, are analyzed with the airfoil design and analysis program XFOIL. The thin airfoil blade is designed and calculated by blade element and momentum theory. The performance characteristics of the machine such as rotational speed, generator output as well as stability for wind speed changes are described. In the case of average wind speeds of 10 m/s and a maximum of 19 m/s, the automatically controlled wind turbine ran safely through rough wind conditions and showed an average generator output of 1105 W and a power coefficient 0.14. (author)

  4. The construction of airfoil pressure models by the plate method: Achievements, current research, technology development and potential applications

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.

    1985-01-01

    A method of constructing airfoils by inscribing pressure channels on the face of opposing plates, bonding them together to form one plate with integral channels, and contour machining this plate to form an airfoil model is described. The research and development program to develop the bonding technology is described as well as the construction and testing of an airfoil model. Sample aerodynamic data sets are presented and discussed. Also, work currently under way to produce thin airfoils with camber is presented. Samples of the aft section of a 6 percent airfoil with complete pressure instrumentation including the trailing edge are pictured and described. This technique is particularly useful in fabricating models for transonic cryogenic testing, but it should find application in a wide ange of model construction projects, as well as the fabrication of fuel injectors, space hardware, and other applications requiring advanced bonding technology and intricate fluid passages.

  5. Optimization of Wind Turbine Airfoils/Blades and Wind Farm Layouts

    NASA Astrophysics Data System (ADS)

    Chen, Xiaomin

    Shape optimization is widely used in the design of wind turbine blades. In this dissertation, a numerical optimization method called Genetic Algorithm (GA) is applied to address the shape optimization of wind turbine airfoils and blades. In recent years, the airfoil sections with blunt trailing edge (called flatback airfoils) have been proposed for the inboard regions of large wind-turbine blades because they provide several structural and aerodynamic performance advantages. The FX, DU and NACA 64 series airfoils are thick airfoils widely used for wind turbine blade application. They have several advantages in meeting the intrinsic requirements for wind turbines in terms of design point, off-design capabilities and structural properties. This research employ both single- and multi-objective genetic algorithms (SOGA and MOGA) for shape optimization of Flatback, FX, DU and NACA 64 series airfoils to achieve maximum lift and/or maximum lift to drag ratio. The commercially available software FLUENT is employed for calculation of the flow field using the Reynolds-Averaged Navier-Stokes (RANS) equations in conjunction with a two-equation Shear Stress Transport (SST) turbulence model and a three equation k-kl-o turbulence model. The optimization methodology is validated by an optimization study of subsonic and transonic airfoils (NACA0012 and RAE 2822 airfoils). In this dissertation, we employ DU 91-W2-250, FX 66-S196-V1, NACA 64421, and Flat-back series of airfoils (FB-3500-0050, FB-3500-0875, and FB-3500-1750) and compare their performance with S809 airfoil used in NREL Phase II and III wind turbines; the lift and drag coefficient data for these airfoils sections are available. The output power of the turbine is calculated using these airfoil section blades for a given B and lambda and is compared with the original NREL Phase II and Phase III turbines using S809 airfoil section. It is shown that by a suitable choice of airfoil section of HAWT blade, the power generated

  6. Development of two supercritical airfoils with a thickness-to-chord ratio of 0.20 and design lift coefficients of 0.3 and 0.4

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.

    1976-01-01

    Two supercritical airfoils were developed specifically for application to span distributed loading cargo aircraft. These airfoils have a thickness-to-chord ratio of 0.20 and design lift coefficients of 0.3 and 0.4, and were derived by modifying a recently developed supercritical airfoil having a thickness-to-chord ratio of 0.18 and a design lift coefficient of 0.5. The aerodynamic characteristics were calculated using a theoretical method which computes the flow field about an airfoil having supercritical surface velocities.

  7. Propulsion of a flapping and oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Garrick, I E

    1937-01-01

    Formulas are given for the propelling or drag force experience in a uniform air stream by an airfoil or an airfoil-aileron combination, oscillating in any of three degrees of freedom; vertical flapping, torsional oscillations about a fixed axis parallel to the span, and angular oscillations of the aileron about a hinge.

  8. Measuring Lift with the Wright Airfoils

    ERIC Educational Resources Information Center

    Heavers, Richard M.; Soleymanloo, Arianne

    2011-01-01

    In this laboratory or demonstration exercise, we mount a small airfoil with its long axis vertical at one end of a nearly frictionless rotating platform. Air from a leaf blower produces a sidewise lift force L on the airfoil and a drag force D in the direction of the air flow (Fig. 1). The rotating platform is kept in equilibrium by adding weights…

  9. Subsonic flow over thin oblique airfoils at zero lift

    NASA Technical Reports Server (NTRS)

    Jones, Robert T

    1948-01-01

    A previous report gave calculations for the pressure distribution over thin oblique airfoils at supersonic speed. The present report extends the calculations to subsonic speeds. It is found that the flows again can be obtained by the superposition of elementary conical flow fields. In the case of the swept-back wing the pressure distributions remain qualitatively similar at subsonic and supersonic speeds. Thus a distribution similar to the Ackeret type of distribution appears on the root sections of the swept-back wing at Mach=0. The resulting positive pressure drag on the root section is balanced by negative drags on outboard sections.

  10. Numerical Simulations of Subscale Wind Turbine Rotor Inboard Airfoils at Low Reynolds Number

    SciTech Connect

    Blaylock, Myra L.; Maniaci, David Charles; Resor, Brian R.

    2015-04-01

    New blade designs are planned to support future research campaigns at the SWiFT facility in Lubbock, Texas. The sub-scale blades will reproduce specific aerodynamic characteristics of utility-scale rotors. Reynolds numbers for megawatt-, utility-scale rotors are generally above 2-8 million. The thickness of inboard airfoils for these large rotors are typically as high as 35-40%. The thickness and the proximity to three-dimensional flow of these airfoils present design and analysis challenges, even at the full scale. However, more than a decade of experience with the airfoils in numerical simulation, in the wind tunnel, and in the field has generated confidence in their performance. Reynolds number regimes for the sub-scale rotor are significantly lower for the inboard blade, ranging from 0.7 to 1 million. Performance of the thick airfoils in this regime is uncertain because of the lack of wind tunnel data and the inherent challenge associated with numerical simulations. This report documents efforts to determine the most capable analysis tools to support these simulations in an effort to improve understanding of the aerodynamic properties of thick airfoils in this Reynolds number regime. Numerical results from various codes of four airfoils are verified against previously published wind tunnel results where data at those Reynolds numbers are available. Results are then computed for other Reynolds numbers of interest.

  11. Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Sewall, William G.

    1995-01-01

    Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

  12. Using potential flow theory and conformal mapping technique to measure pressure differential on airfoil

    NASA Astrophysics Data System (ADS)

    Mughal, Umair Najeeb

    2017-01-01

    Flow around an airfoil to calculate pressure co-efficient variations at different relative velocities have always been an important/basic part of Aerodynamic Study. Potential flow theory is used to study flow behavior on rankine half body, non-rotating cylinder and rotating cylinder as it is more trackable. Falkan-Skan Similarity Solution is taken to simulate the flow behavior on wedge. However, to use potential flow theory on usable airfoils the author have used conformal mapping to show a relation between realistic airfoil shapes and the knowledge gained from flow about cylinders. This method can further be used in the designing of an airfoil section. The author has used Joukowski Tranform to generate the flow around airfoils of various geometries and then utilized Kutta condition to force the stagnation point at the trailing edge. Co-efficient of pressure over the entire airfoil surface were calculated and corrected using Karman-Tsien compressibility correction equations. On the basis of this, the location of the ports to install the flush measurement system is suggested.

  13. A critical assessment of UH-60 main rotor blade airfoil data

    NASA Technical Reports Server (NTRS)

    Totah, Joseph

    1993-01-01

    Many current comprehensive rotorcraft analyses employ lifting-line methods that require main rotor blade airfoil data, typically obtained from wind tunnel tests. In order to effectively evaluate these lifting-line methods, it is of the utmost importance to ensure that the airfoil section data are free of inaccuracies. A critical assessment of the SC1095 and SC1094R8 airfoil data used on the UH-60 main rotor blade was performed for that reason. Nine sources of wind tunnel data were examined, all of which contain SC1095 data and four of which also contain SC1094R8 data. Findings indicate that the most accurate data were generated in 1982 at the 11-Foot Wind Tunnel Facility at NASA Ames Research Center and in 1985 at the 6-inch by 22-inch transonic wind tunnel facility at Ohio State University. It has not been determined if data from these two sources are sufficiently accurate for their use in comprehensive rotorcraft analytical models of the UH-60. It is recommended that new airfoil tables be created for both airfoils using the existing data. Additional wind tunnel experimentation is also recommended to provide high quality data for correlation with these new airfoil tables.

  14. A critical assessment of UH-60 main rotor blade airfoil data

    NASA Technical Reports Server (NTRS)

    Totah, Joseph

    1993-01-01

    Many current comprehensive rotorcraft analyses employ lifting-line methods that require main rotor blade airfoil data, typically obtained from wind tunnel tests. In order to effectively evaluate these lifting-line methods, it is of the utmost importance to ensure that the airfoil section data are free of inaccuracies. A critical assessment of the SC1095 and SC1094R8 airfoil data used on the UH-60 main rotor blade was performed for that reason. Nine sources of wind tunnel data were examined, all of which contain SC1095 data and four of which also contain SC1094R8 data. Findings indicate that the most accurate data were generated in 1982 at the 11-Foot Wind Tunnel Facility at NASA Ames Research Center and in 1985 at the 6-inch-by-22-inch transonic wind tunnel facility at Ohio State University. It has not been determined if data from these two sources are sufficiently accurate for their use in comprehensive rotorcraft analytical models of the UH-60. It is recommended that new airfoil tables be created for both airfoils using the existing data. Additional wind tunnel experimentation is also recommended to provide high quality data for correlation with these new airfoil tables.

  15. On the acoustic signature of tandem airfoils: The sound of an elastic airfoil in the wake of a vortex generator

    NASA Astrophysics Data System (ADS)

    Manela, A.

    2016-07-01

    The acoustic signature of an acoustically compact tandem airfoil setup in uniform high-Reynolds number flow is investigated. The upstream airfoil is considered rigid and is actuated at its leading edge with small-amplitude harmonic pitching motion. The downstream airfoil is taken passive and elastic, with its motion forced by the vortex-street excitation of the upstream airfoil. The non-linear near-field description is obtained via potential thin-airfoil theory. It is then applied as a source term into the Powell-Howe acoustic analogy to yield the far-field dipole radiation of the system. To assess the effect of downstream-airfoil elasticity, results are compared with counterpart calculations for a non-elastic setup, where the downstream airfoil is rigid and stationary. Depending on the separation distance between airfoils, airfoil-motion and airfoil-wake dynamics shift between in-phase (synchronized) and counter-phase behaviors. Consequently, downstream airfoil elasticity may act to amplify or suppress sound through the direct contribution of elastic-airfoil motion to the total signal. Resonance-type motion of the elastic airfoil is found when the upstream airfoil is actuated at the least stable eigenfrequency of the downstream structure. This, again, results in system sound amplification or suppression, depending on the separation distance between airfoils. With increasing actuation frequency, the acoustic signal becomes dominated by the direct contribution of the upstream airfoil motion, whereas the relative contribution of the elastic airfoil to the total signature turns negligible.

  16. An airfoil pitch apparatus-modeling and control design

    NASA Astrophysics Data System (ADS)

    Andrews, Daniel R.

    1989-03-01

    The study of dynamic stall of rapidly pitching airfoils is being conducted at NASA Ames Research Center. Understanding this physical phenomenon will aid in improving the maneuverability of fighter aircraft as well as civilian aircraft. A wind tunnel device which can linearly pitch and control an airfoil with rapid dynamic response is needed for such tests. To develop a mechanism capable of high accelerations, an accurate model and control system is created. The model contains mathematical representations of the mechanical system, including mass, spring, and damping characteristics for each structural element, as well as coulomb friction and servovalve saturation. Electrical components, both digital and analog, linear and nonlinear, are simulated. The implementation of such a high-performance system requires detailed control design as well as state-of-the-art components. This paper describes the system model, states the system requirements, and presents results of its theoretical performance which maximizes the structural and hydraulic aspects of this system.

  17. Prediction of Film Cooling on Gas Turbine Airfoils

    NASA Technical Reports Server (NTRS)

    Garg, Vijay K.; Gaugler, Raymond E.

    1994-01-01

    A three-dimensional Navier-Stokes analysis tool has been developed in order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine airfoils. An existing code (Arnone et al., 1991) has been modified for the purpose. The code is an explicit, multigrid, cell-centered, finite volume code with an algebraic turbulence model. Eigenvalue scaled artificial dissipation and variable-coefficient implicit residual smoothing are used with a full-multigrid technique. Moreover, Mayle's transition criterion (Mayle, 1991) is used. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the airfoil surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison is fair with near mid-span experimental data for four and nine rows of cooling holes, five on the shower head, and two rows each on the pressure and suction surfaces. The computations, however, show a strong spanwise variation of the heat transfer coefficient on the airfoil surface, specially with shower-head cooling.

  18. Kasprzyk airfoil. The first wind-tunnel tests

    NASA Technical Reports Server (NTRS)

    Wusatowski, T.

    1984-01-01

    The Kasprzyk slotted flap glider airfoil (the Kasper wing) enabling glider flight at 32 km/h and 0.5 m/sec descent speed was wind tunnel tested in the U.S. The test layout is described and reasons offered for discrepancies between wind tunnel results and Polish in flight data: high induced drag caused by relative size of model wing span and tunnel, by vortex attenuators on the model and their proximity to the tunnel wall, nonsimilarity between flow over a smooth wing and flow over the Kasprzyk wing with bound vortices, obstruction of the tunnel test chamber cross section by the model wing, discrepant Reynolds numbers, and model airfoil aspect ratio much smaller than the prototype. The overall results offer partial confirmation of the Kasprzyk theory, but further in tunnel and in flight studies are recommended.

  19. CAST-10-2/DOA 2 Airfoil Studies Workshop Results

    NASA Technical Reports Server (NTRS)

    Ray, Edward J. (Compiler); Hill, Acquilla S. (Compiler)

    1989-01-01

    During the period of September 23 through 27, 1988, the Transonic Aerodynamics Division at the Langely Research Center hosted an International Workshop on CAST-10-2/DOA 2 Airfoil Studies. The CAST-10 studies were the outgrowth of several cooperative study agreements among the NASA, the NAE of Canada, the DLR of West Germany, and the ONERA of France. Both theoretical and experimental CAST-10 airfoil results that were obtained form an extensive series of tests and studies, were reviewed. These results provided an opportunity to make direct comparisons of adaptive wall test section (AWTS) results from the NASA 0.3-meter Transonic Cryogenic Tunnel and ONERA T-2 AWTS facilities with conventional ventilated wall wind tunnel results from the Canadian high Reynolds number two-dimensional test facility. Individual papers presented during the workshop are included.

  20. A Computational Modeling Mystery Involving Airfoil Trailing Edge Treatments

    NASA Astrophysics Data System (ADS)

    Choo, Yeunun; Epps, Brenden

    2015-11-01

    In a curious result, Fairman (2002) observed that steady RANS calculations predicted larger lift than the experimentally-measured data for six different airfoils with non-traditional trailing edge treatments, whereas the time average of unsteady RANS calculations matched the experiments almost exactly. Are these results reproducible? If so, is the difference between steady and unsteady RANS calculations a numerical artifact, or is there a physical explanation? The goals of this project are to solve this thirteen year old mystery and further to model viscous/load coupling for airfoils with non-traditional trailing edges. These include cupped, beveled, and blunt trailing edges, which are common anti-singing treatments for marine propeller sections. In this talk, we present steady and unsteady RANS calculations (ANSYS Fluent) with careful attention paid to the possible effects of asymmetric unsteady vortex shedding and the modeling of turbulence anisotropy. The effects of non-traditional trailing edge treatments are visualized and explained.

  1. An experimental study of transonic flow about a supercritical airfoil

    NASA Technical Reports Server (NTRS)

    Spaid, F. W.; Dahlin, J. A.; Bachalo, W. D.; Stivers, L. S., Jr.

    1983-01-01

    A series of experiments was conducted on flow fields about two airfoil models whose sections are slight modifications of the original Whitcomb supercritical airfoil section. Data obtained include surface static-pressure distributions, far-wake surveys, oil-flow photographs, pitot-pressure surveys in the viscous regions, and holographic interferograms. These data were obtained for different combinations of lift coefficient and free-stream Mach number, which included both subcritical cases and flows with upper-surface shock waves. The availability of both pitot-pressure data and density data from interferograms allowed determination of flow-field properties in the vicinity of the trailing edge and in the wake without recourse to any assumptions about the local static pressure. The data show that significant static-pressure gradients normal to viscous layers exist in this region, and that they persist to approximately 10% chord downstream of the trailing edge. Comparisons are made between measured boundary-layer properties and results from boundary-layer computations that employed measured static-pressure distributions, as well as comparisons between data and results of airfoil flow-field computations.

  2. Nonlinear effects of flow unsteadiness on the acoustic radiation of a heaving airfoil

    NASA Astrophysics Data System (ADS)

    Manela, Avshalom

    2013-12-01

    The study considers the combined effects of boundary animation (small-amplitude heaving) and incoming flow unsteadiness (incident vorticity) on the vibroacoustic signature of a thin rigid airfoil in low-Mach number flow. The potential-flow problem is analysed using the Brown and Michael equation, yielding the incident vortex trajectory and time evolution of trailing edge wake. The dynamical description serves as an effective source term to evaluate the far-field sound using Powell-Howe analogy. The results identify the fluid-airfoil system as a dipole-type source, and demonstrate the significance of nonlinear eddy-airfoil interactions on the acoustic radiation. Based on the value of scaled heaving frequency ωa/U (with ω the dimensional heaving frequency, a the airfoil half-chord, and U the mean flow speed), the system behaviour can be divided into two characteristic regimes: (i) for ωa/U≪1, the effect of heaving is minor, and the acoustic response is well approximated by considering the interaction of a line vortex with a stationary airfoil; (ii) for ωa/U≫1, the impact of heaving is dominant, radiating sound through an “airfoil motion” dipole oriented along the direction of heaving. In between (for ωa/U~O(1)), an intermediate regime takes place. The results indicate that trailing edge vorticity has a two-fold impact on the acoustic far field: while reducing pressure fluctuations generated by incident vortex interaction with the airfoil, trailing edge vortices transmit sound along the mean-flow direction, characterized by airfoil heaving frequency. The “silencing” effect of trailing edge vorticity is particularly efficient when the incident vortex passes close to the airfoil trailing edge: at that time, application of the Kutta condition implies the release of a trailing edge vortex in the opposite direction to the incident vortex; the released vortex then detaches from the airfoil and follows the incident vortex, forming a “silent” vortex pair

  3. Aerodynamic Characteristics of Airfoils. Volume 4.

    DTIC Science & Technology

    1927-01-01

    8-6-4 -2 02a46 a 10 12 14 16is 20 2Angie of Attack in Degrees. Angle of Attack in Deprese . I~r xi N . A rxi (𔃺\\I ) M ITTll IK’ FORi AERONAUT~ICS RFM...8217 ~ 86--20a2 46 a10 12 14 163826 D Aogle of Attack in Deprese . Angle of it tak in Deprese . A II~ il A CHC ARACTERIISTICS OF Al RFOILS-i V 205 nzmuRRCS

  4. Theory and Simulation for Traffic Characteristics on the Highway with a Slowdown Section.

    PubMed

    Xu, Dejie; Mao, Baohua; Rong, Yaping; Wei, Wei

    2015-01-01

    We study the traffic characteristics on a single-lane highway with a slowdown section using the deterministic cellular automaton (CA) model. Based on the theoretical analysis, the relationships among local mean densities, velocities, traffic fluxes, and global densities are derived. The results show that two critical densities exist in the evolutionary process of traffic state, and they are significant demarcation points for traffic phase transition. Furthermore, the changing laws of the two critical densities with different length of limit section are also investigated. It is shown that only one critical density appears if a highway is not slowdown section; nevertheless, with the growing length of slowdown section, one critical density separates into two critical densities; if the entire highway is slowdown section, they finally merge into one. The contrastive analysis proves that the analytical results are consistent with the numerical ones.

  5. Theory and Simulation for Traffic Characteristics on the Highway with a Slowdown Section

    PubMed Central

    Xu, Dejie; Mao, Baohua; Rong, Yaping; Wei, Wei

    2015-01-01

    We study the traffic characteristics on a single-lane highway with a slowdown section using the deterministic cellular automaton (CA) model. Based on the theoretical analysis, the relationships among local mean densities, velocities, traffic fluxes, and global densities are derived. The results show that two critical densities exist in the evolutionary process of traffic state, and they are significant demarcation points for traffic phase transition. Furthermore, the changing laws of the two critical densities with different length of limit section are also investigated. It is shown that only one critical density appears if a highway is not slowdown section; nevertheless, with the growing length of slowdown section, one critical density separates into two critical densities; if the entire highway is slowdown section, they finally merge into one. The contrastive analysis proves that the analytical results are consistent with the numerical ones. PMID:26089864

  6. Flexural fatigue life prediction of closed hat-section using materially nonlinear axial fatigue characteristics

    NASA Technical Reports Server (NTRS)

    Razzaq, Zia

    1989-01-01

    Straight or curved hat-section members are often used as structural stiffeners in aircraft. For instance, they are employed as stiffeners for the dorsal skin as well as in the aerial refueling adjacent area structure in F-106 aircraft. The flanges of the hat-section are connected to the aircraft skin. Thus, the portion of the skin closing the hat-section interacts with the section itself when resisting the stresses due to service loads. The flexural fatigue life of such a closed section is estimated using materially nonlinear axial fatigue characteristics. It should be recognized that when a structural shape is subjected to bending, the fatigue life at the neutral axis is infinity since the normal stresses are zero at that location. Conversely, the fatigue life at the extreme fibers where the normal bending stresses are maximum can be expected to be finite. Thus, different fatigue life estimates can be visualized at various distances from the neural axis. The problem becomes compounded further when significant portions away from the neutral axis are stressed into plastic range. A theoretical analysis of the closed hat-section subjected to flexural cyclic loading is first conducted. The axial fatigue characteristics together with the related axial fatigue life formula and its inverted form given by Manson and Muralidharan are adopted for an aluminum alloy used in aircraft construction. A closed-form expression for predicting the flexural fatigue life is then derived for the closed hat-section including materially nonlinear action. A computer program is written to conduct a study of the variables such as the thicknesses of the hat-section and the skin, and the type of alloy used. The study has provided a fundamental understanding of the flexural fatigue life characteristics of a practical structural component used in aircraft when materially nonlinear action is present.

  7. Aerodynamic sound of flow past an airfoil

    NASA Technical Reports Server (NTRS)

    Wang, Meng

    1995-01-01

    The long term objective of this project is to develop a computational method for predicting the noise of turbulence-airfoil interactions, particularly at the trailing edge. We seek to obtain the energy-containing features of the turbulent boundary layers and the near-wake using Navier-Stokes Simulation (LES or DNS), and then to calculate the far-field acoustic characteristics by means of acoustic analogy theories, using the simulation data as acoustic source functions. Two distinct types of noise can be emitted from airfoil trailing edges. The first, a tonal or narrowband sound caused by vortex shedding, is normally associated with blunt trailing edges, high angles of attack, or laminar flow airfoils. The second source is of broadband nature arising from the aeroacoustic scattering of turbulent eddies by the trailing edge. Due to its importance to airframe noise, rotor and propeller noise, etc., trailing edge noise has been the subject of extensive theoretical (e.g. Crighton & Leppington 1971; Howe 1978) as well as experimental investigations (e.g. Brooks & Hodgson 1981; Blake & Gershfeld 1988). A number of challenges exist concerning acoustic analogy based noise computations. These include the elimination of spurious sound caused by vortices crossing permeable computational boundaries in the wake, the treatment of noncompact source regions, and the accurate description of wave reflection by the solid surface and scattering near the edge. In addition, accurate turbulence statistics in the flow field are required for the evaluation of acoustic source functions. Major efforts to date have been focused on the first two challenges. To this end, a paradigm problem of laminar vortex shedding, generated by a two dimensional, uniform stream past a NACA0012 airfoil, is used to address the relevant numerical issues. Under the low Mach number approximation, the near-field flow quantities are obtained by solving the incompressible Navier-Stokes equations numerically at chord

  8. An evaluation of three helicopter rotor sections

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Collins, L. J.

    1985-01-01

    Three helicopter rotor sections were tested in the NASA Ames Research Center 2- by 2-Foot Transonic Wind Tunnel over a Mach range from 0.2 to 0.88. The sections tested had maximum thickness/chord ratios of 0.078, 0.09, and 0.10. The thickest section was of early technology and had been tested previously in other wind tunnels. This section was included in the investigation to establish a basis for comparing the two thinner sections, which were of recent design. The results of the investigation showed that the pitching-moment characteristics for the three airfoil sections were acceptable. The drag divergence Mach numbers for the three sections were 0.80, 0.825, and 0.845 in order of decreasing thickness.

  9. Airfoil flutter model suspension system

    NASA Technical Reports Server (NTRS)

    Reed, Wilmer H. (Inventor)

    1987-01-01

    A wind tunnel suspension system for testing flutter models under various loads and at various angles of attack is described. The invention comprises a mounting bracket assembly affixing the suspension system to the wind tunnel, a drag-link assembly and a compound spring arrangement comprises a plunge spring working in opposition to a compressive spring so as to provide a high stiffness to trim out steady state loads and simultaneously a low stiffness to dynamic loads. By this arrangement an airfoil may be tested for oscillatory response in both plunge and pitch modes while being held under high lifting loads in a wind tunnel.

  10. Design optimization of transonic airfoils

    NASA Technical Reports Server (NTRS)

    Joh, C.-Y.; Grossman, B.; Haftka, R. T.

    1991-01-01

    Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.

  11. Adaptation of the Theodorsen theory to the representation of an airfoil as a combination of a lifting line and a thickness distribution

    NASA Technical Reports Server (NTRS)

    Barger, R. L.

    1975-01-01

    The theory provides a direct method for resolving an airfoil into a lifting line and a thickness distribution as well as a means of synthesizing thickness and lift components into a resultant airfoil and computing its aerodynamic characteristics. Specific applications of the technique are discussed.

  12. Spline-Based Smoothing of Airfoil Curvatures

    NASA Technical Reports Server (NTRS)

    Li, W.; Krist, S.

    2008-01-01

    Constrained fitting for airfoil curvature smoothing (CFACS) is a splinebased method of interpolating airfoil surface coordinates (and, concomitantly, airfoil thicknesses) between specified discrete design points so as to obtain smoothing of surface-curvature profiles in addition to basic smoothing of surfaces. CFACS was developed in recognition of the fact that the performance of a transonic airfoil is directly related to both the curvature profile and the smoothness of the airfoil surface. Older methods of interpolation of airfoil surfaces involve various compromises between smoothing of surfaces and exact fitting of surfaces to specified discrete design points. While some of the older methods take curvature profiles into account, they nevertheless sometimes yield unfavorable results, including curvature oscillations near end points and substantial deviations from desired leading-edge shapes. In CFACS as in most of the older methods, one seeks a compromise between smoothing and exact fitting. Unlike in the older methods, the airfoil surface is modified as little as possible from its original specified form and, instead, is smoothed in such a way that the curvature profile becomes a smooth fit of the curvature profile of the original airfoil specification. CFACS involves a combination of rigorous mathematical modeling and knowledge-based heuristics. Rigorous mathematical formulation provides assurance of removal of undesirable curvature oscillations with minimum modification of the airfoil geometry. Knowledge-based heuristics bridge the gap between theory and designers best practices. In CFACS, one of the measures of the deviation of an airfoil surface from smoothness is the sum of squares of the jumps in the third derivatives of a cubicspline interpolation of the airfoil data. This measure is incorporated into a formulation for minimizing an overall deviation- from-smoothness measure of the airfoil data within a specified fitting error tolerance. CFACS has been

  13. Forcing function effects on unsteady aerodynamic gust response: Part 2--Low solidity airfoil row response

    SciTech Connect

    Henderson, G.H.; Fleeter, S. . School of Mechanical Engineering)

    1993-10-01

    The fundamental gust modeling assumption is investigated by means of series of experiments performed in the Purdue Annular Cascade Research Facility. The unsteady periodic flow field is generated by rotating rows of perforated plates and airfoil cascades, with the resulting unsteady periodic chord wise pressure response of a downstream low-solidity stator row determined by miniature pressure transducers embedded within selected airfoils. When the forcing function exhibited the characteristic of a linear-theory vortical gust, as was the case for the perforated-plate wake generators, the resulting response on the downstream stator airfoils was in excellent agreement with the linear-theory models. In contrast, when the forcing function did not exhibit linear-theory vortical gust characteristics, i.e., for the airfoil wake generators, the resulting unsteady aerodynamic responses of the downstream stators were much more complex and correlated poorly with the linear-theory gust predictions. Thus, this investigation has quantitatively shown that the forcing function generator significantly affects the resulting gust response, with the complexity of the response characteristics increasing from the perforated-plate to the airfoil-cascade forcing functions.

  14. Numerical analysis of combustion characteristics of hybrid rocket motor with multi-section swirl injection

    NASA Astrophysics Data System (ADS)

    Li, Chengen; Cai, Guobiao; Tian, Hui

    2016-06-01

    This paper is aimed to analyse the combustion characteristics of hybrid rocket motor with multi-section swirl injection by simulating the combustion flow field. Numerical combustion flow field and combustion performance parameters are obtained through three-dimensional numerical simulations based on a steady numerical model proposed in this paper. The hybrid rocket motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Multiple injection sections are set along the axis of the solid fuel grain, and the oxidizer enters the combustion chamber by means of tangential injection via the injector ports in the injection sections. Simulation results indicate that the combustion flow field structure of the hybrid rocket motor could be improved by multi-section swirl injection method. The transformation of the combustion flow field can greatly increase the fuel regression rate and the combustion efficiency. The average fuel regression rate of the motor with multi-section swirl injection is improved by 8.37 times compared with that of the motor with conventional head-end irrotational injection. The combustion efficiency is increased to 95.73%. Besides, the simulation results also indicate that (1) the additional injection sections can increase the fuel regression rate and the combustion efficiency; (2) the upstream offset of the injection sections reduces the combustion efficiency; and (3) the fuel regression rate and the combustion efficiency decrease with the reduction of the number of injector ports in each injection section.

  15. Uncertainty Quantification for Airfoil Icing

    NASA Astrophysics Data System (ADS)

    DeGennaro, Anthony Matteo

    Ensuring the safety of airplane flight in icing conditions is an important and active arena of research in the aerospace community. Notwithstanding the research, development, and legislation aimed at certifying airplanes for safe operation, an analysis of the effects of icing uncertainties on certification quantities of interest is generally lacking. The central objective of this thesis is to examine and analyze problems in airfoil ice accretion from the standpoint of uncertainty quantification. We focus on three distinct areas: user-informed, data-driven, and computational uncertainty quantification. In the user-informed approach to uncertainty quantification, we discuss important canonical icing classifications and show how these categories can be modeled using a few shape parameters. We then investigate the statistical effects of these parameters. In the data-driven approach, we build statistical models of airfoil ice shapes from databases of actual ice shapes, and quantify the effects of these parameters. Finally, in the computational approach, we investigate the effects of uncertainty in the physics of the ice accretion process, by perturbing the input to an in-house numerical ice accretion code that we develop in this thesis.

  16. Flow control at low Reynolds numbers using periodic airfoil morphing

    NASA Astrophysics Data System (ADS)

    Jones, Gareth; Santer, Matthew; Papadakis, George; Bouremel, Yann; Debiasi, Marco; Imperial-NUS Joint PhD Collaboration

    2014-11-01

    The performance of airfoils operating at low Reynolds numbers is known to suffer from flow separation even at low angles of attack as a result of their boundary layers remaining laminar. The lack of mixing---a characteristic of turbulent boundary layers---leaves laminar boundary layers with insufficient energy to overcome the adverse pressure gradient that occurs in the pressure recovery region. This study looks at periodic surface morphing as an active flow control technique for airfoils in such a flight regime. It was discovered that at sufficiently high frequencies an oscillating surface is capable of not only reducing the size of the separated region---and consequently significantly reducing drag whilst simultaneously increasing lift---but it is also capable of delaying stall and as a result increasing CLmax. Furthermore, by bonding Macro Fiber Composite actuators (MFCs) to the underside of an airfoil skin and driving them with a sinusoidal frequency, it is shown that this control technique can be practically implemented in a lightweight, energy efficient way. Imperial-NUS Joint Ph.D. Programme.

  17. Aerodynamic Control of a Pitching Airfoil by Distributed Bleed Actuation

    NASA Astrophysics Data System (ADS)

    Kearney, John; Glezer, Ari

    2013-11-01

    The aerodynamic forces and moments on a dynamically pitching 2-D airfoil model are controlled in wind tunnel experiments using distributed active bleed. Bleed flow on the suction surface downstream of the leading edge is driven by pressure differences across the airfoil and is regulated by low-power louver actuators. The bleed interacts with cross flows to effect time-dependent variations of the vorticity flux and thereby alters the local flow attachment, resulting in significant changes in pre- and post-stall lift and pitching moment (over 50% increase in baseline post-stall lift). The flow field over the airfoil is measured using high-speed (2000 fps) PIV, resolving the dynamics and characteristic time-scales of production and advection of vorticity concentrations that are associated with transient variations in the aerodynamic forces and moments. In particular, it is shown that the actuation improves the lift hysteresis and pitch stability during the oscillatory pitching by altering the evolution of the dynamic stall vortex and the ensuing flow attachment during the downstroke. Supported by the Rotorcraft Center (VLRCOE) at Georgia Tech.

  18. Wind turbine airfoil investigations in customized turbulent inflow

    NASA Astrophysics Data System (ADS)

    Heisselmann, Hendrik; Peinke, Joachim; Hoelling, Michael

    2016-11-01

    Experimental airfoil characterizations are usually performed in laminar or unsteady periodical flows. Neither of these matches the flow conditions of natural atmospheric flows as experienced by wind turbine blades. In the presented experimental study, an active grid is used to generate turbulent inflow with customized properties, like reduced frequencies or inflow angles. This is used not only to tune flow properties, but also to mimic time series of measured atmospheric wind speeds and inflow angles in the wind tunnel. Experiments were performed on a wind turbine dedicated DU 00-W-212 airfoil to obtain highly resolved force data and chord-wise pressure distributions at Re=500,000 and Re=900,000. Additional to a laminar baseline case, unsteady sinusoidal inflow fluctuations were applied as well as three different turbulent inflows with comparable turbulence intensity, but different inflow angle fluctuations to grasp the impact of inflow characteristics on the airfoil performance. In comparison with the laminar inflow case, the lift peak of the polar is shifted to higher angles of attack in the turbulent flows. While the laminar lift polars show a rather sudden transition to stall, a softer transition with an extended stall region is found for all turbulent cases. The presented work was performed within the project AVATAR and is funded from the European Unions Seventh Program for research, technological development and demonstration under Grand Agreement No FP7-ENERGY-2013-1/n 608396.

  19. Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter

    NASA Technical Reports Server (NTRS)

    Mahajan, A. J.; Kaza, K. R. V.; Dowell, E. H.

    1993-01-01

    A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.

  20. Semi-empirical model for prediction of unsteady forces on an airfoil with application to flutter

    NASA Technical Reports Server (NTRS)

    Mahajan, Aparajit J.; Kaza, Krishna Rao V.

    1992-01-01

    A semi-empirical model is described for predicting unsteady aerodynamic forces on arbitrary airfoils under mildly stalled and unstalled conditions. Aerodynamic forces are modeled using second order ordinary differential equations for lift and moment with airfoil motion as the input. This model is simultaneously integrated with structural dynamics equations to determine flutter characteristics for a two degrees-of-freedom system. Results for a number of cases are presented to demonstrate the suitability of this model to predict flutter. Comparison is made to the flutter characteristics determined by a Navier-Stokes solver and also the classical incompressible potential flow theory.

  1. Navier-Stokes simulations of WECS airfoil flowfields

    SciTech Connect

    Homicz, G.F.

    1994-06-01

    Sandia National Laboratories has initiated an effort to apply Computational Fluid Dynamics (CFD) to the study of WECS aerodynamics. Preliminary calculations are presented for the flow past a SAND 0018/50 airfoil. The flow solver used is F3D, an implicitly, finite-difference code which solves the Thin-Layer Navier-airfoil. The flow solver used is F3D, an implicit, finite-difference code which solves the Thin-Layer Navier-Stokes equations. 2D steady-state calculations are presented at various angles of attack, {alpha}. Sectional lift and drag coefficient, as well as surface pressure distributions, are compared with wind tunnel data, and exhibit reasonable agreement at low to moderate angles of attack. At high {alpha}, where the airfoil is stalled, a converged solution to the steady-state equations could not be obtained. The flowfield continued to change with successive iterations, which is consistent with the fact that the actual flow is inherently transient, and requires the solution of the full unsteady form of the equations.

  2. Explanation of the effects of leading-edge tubercles on the aerodynamics of airfoils and finite wings

    NASA Astrophysics Data System (ADS)

    Saadat, Mehdi; Haj-Hariri, Hossein; Fish, Frank

    2010-11-01

    A computational study was conducted to explain the aerodynamic effect of leading edge tubercles on maximum lift coefficient, stall angle of attack (AoA), drag, and post stall characteristics for airfoils as well as finite wings. Past experiments demonstrated airfoils with leading edge tubercles do not improve Clmax, drag, or stall AoA but smoothen post stall characteristics to a great degree. In contrast to airfoils, finite wings with L.E. tubercles improved all aerodynamic characteristics. We explain the stall mechanism of the tubercled wing by considering each L.E. tubercle as a combination of a swept forward and a swept backward wing.There are 3 mechanisms (streamline curvature, accelerated stall, and upwash) that cause Clmax of airfoils with L.E. tubercles always be lower than that of smooth airfoils. We also identify two additional mechanisms which are responsible for improved post-stall characteristics of airfoils with L.E. tubercles. Finally, we discuss why finite wings with L.E. tubercles have higher Clmax and lower drag than their smooth L.E. counterparts by studying effects of wing tip, sweep, and taper ratio.

  3. A comparative study on two characteristic parametrizations for high energy pp and overlinepp total cross sections

    NASA Astrophysics Data System (ADS)

    Bueno, A.; Velasco, J.

    1996-02-01

    Available high energy data for both pp and overlinepp total cross sections ( f GeV < s < 1.8 TeV ) are described by means of two well-known distinct parametrizations, characteristic of theoretical (“Regge-like” expression) and experimental (“Froissart-Martin-like” expression) practices, respectively. Both are compared from the statistical point of view. For the whole set of present data statistical analysis ( χ2/d.o.f.) seems to favour a “Froissart-like” ((ln s) γ≈2 ) rise of the total cross section rather than a “Regge-like” ( sɛ) one.

  4. Characteristic plateau in the L1 -subshell ionization cross section of Ag induced by proton collisions

    NASA Astrophysics Data System (ADS)

    Kavčič, M.; Šmit, Ž.

    2009-12-01

    A characteristic dependence of the proton-induced 2s subshell ionization cross section generated by the nodal structure of the 2s wave function was probed via high-resolution measurements of the Lβ3 (L1-M3) and Lβ6 (L3-N1) x-ray lines of Ag. The intensity ratio of the measured lines depends strongly on the proton energy exhibiting a pronounced maximum around 0.4 MeV. The experimental intensities are critically compared to the theoretically predicted values. The latter were obtained using the ionization cross sections calculated within the semiclassical approximation employing hydrogenlike and Dirac-Fock electron wave functions.

  5. Turbine airfoil with outer wall thickness indicators

    DOEpatents

    Marra, John J; James, Allister W; Merrill, Gary B

    2013-08-06

    A turbine airfoil usable in a turbine engine and including a depth indicator for determining outer wall blade thickness. The airfoil may include an outer wall having a plurality of grooves in the outer surface of the outer wall. The grooves may have a depth that represents a desired outer surface and wall thickness of the outer wall. The material forming an outer surface of the outer wall may be removed to be flush with an innermost point in each groove, thereby reducing the wall thickness and increasing efficiency. The plurality of grooves may be positioned in a radially outer region of the airfoil proximate to the tip.

  6. High Reynolds number tests of a Boeing BAC I airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, W. G., Jr.; Hill, A. S.; Ray, E. J.; Rozendaal, R. A.; Butler, T. W.

    1982-01-01

    A wind tunnel investigation of an advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents the first in a series of NASA/U.X. industry two dimensional airfoil studies to be completed in the Advanced Technology Airfoil Test program. Test temperature was varied from ambient to about 100 K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from about 0.40 to 0.80. These variables provided a Reynolds number (based on airfoil chord) range from about .0000044 to .00005. This investigation was specifically designed to: (1) test a Boeing advanced airfoil from low to flight-equivalent Reynolds numbers; (2) provide the industry participant (Boeing) with experience in cryogenic wind-tunnel model design and testing techniques; and (3) demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. All the objectives of the cooperative test were met. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics of the airfoil. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

  7. Modeling and computation of flow in a passage with 360 deg turning and multiple airfoils

    NASA Astrophysics Data System (ADS)

    Shyy, W.; Vu, T. C.

    1991-06-01

    Numerical modeling of the three-dimensional flows in a spiral casing of a hydraulic turbine, containing a passage of 360-deg turning and multiple elements of airfoils (the so-called distributor), is made. The physical model is based on a novel two-level approach, comprising of (1) a global model that adequately accounts for the geometry of the spiral casing but smears out the details of the distributor and represents the multiple airfoils by a porous medium treatment; and (2) a local model that performs detailed analysis of flow in the distributor region. The global analysis supplies the inlet flow condition for the individual cascade of distributor airfoils, while the distributor analysis yields the information needed for modeling the characteristics of the porous medium. Comparisons of pressure and velocity profiles between measurement and prediction have been made to assess the validity of the present approach. Flow characteristics in the spiral casing are also discussed.

  8. Transitory Control of the Aerodynamic Loads on an Airfoil in Dynamic Pitch and Plunge

    NASA Astrophysics Data System (ADS)

    Tan, Yuehan; Crittenden, Thomas; Glezer, Ari

    2016-11-01

    Transitory control and regulation of trapped vorticity concentrations are exploited in wind tunnel experiments for control of the aerodynamic loads on an airfoil moving in time-periodic 2-DOF (pitch and plunge) beyond the dynamic stall margin. Actuation is effected using a spanwise array of integrated miniature chemical (combustion based) high-impulse actuators that are triggered intermittently relative to the airfoil's motion. Each actuation pulse has sufficient control authority to alter the global aerodynamic performance throughout the motion cycle on a characteristic time scale that is an order of magnitude shorter than the airfoil's convective time scale. The effects of the actuation on the aerodynamic characteristics of the airfoil are assessed using time-dependent measurements of the lift force and pitching moment coupled with time-resolved particle image velocimetry that is acquired phased-locked to the motion of the airfoil. It is shown that the aerodynamic loads can be significantly altered using actuation programs based on multiple actuation pulses during the time-periodic pitch/plunge cycle. Superposition of such actuation programs leads to enhancement of cycle lift and pitch stability, and reduced cycle hysteresis and peak pitching moment. Supported by GT-VLRCOE.

  9. Unsteady Aerodynamic Response of a Linear Cascade of Airfoils in Separated Flow

    NASA Technical Reports Server (NTRS)

    Capece, Vincent R.; Ford, Christopher; Bone, Christopher; Li, Rui

    2004-01-01

    The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.

  10. Two-dimensional cascade test of a highly loaded, low-solidity, tandem airfoil turbine rotor blade

    NASA Technical Reports Server (NTRS)

    Kline, J. F.; Stabe, R. G.

    1973-01-01

    A tip region section of a low-solidity tandem airfoil blade for a turbine rotor was tested in a two-dimensional cascade tunnel at solidities of 0.736 and 0.912. Blade surface static pressures and blade exit total and static pressure and flow angle were surveyed. Blade surface velocities, wake shapes, and kinetic energy losses were analyzed and compared with values for 1.852 solidity tandem airfoil blading.

  11. Theory and Low-Order Modeling of Unsteady Airfoil Flows

    NASA Astrophysics Data System (ADS)

    Ramesh, Kiran

    Unsteady flow phenomena are prevalent in a wide range of problems in nature and engineering. These include, but are not limited to, aerodynamics of insect flight, dynamic stall in rotorcraft and wind turbines, leading-edge vortices in delta wings, micro-air vehicle (MAV) design, gust handling and flow control. The most significant characteristics of unsteady flows are rapid changes in the circulation of the airfoil, apparent-mass effects, flow separation and the leading-edge vortex (LEV) phenomenon. Although experimental techniques and computational fluid dynamics (CFD) methods have enabled the detailed study of unsteady flows and their underlying features, a reliable and inexpensive loworder method for fast prediction and for use in control and design is still required. In this research, a low-order methodology based on physical principles rather than empirical fitting is proposed. The objective of such an approach is to enable insights into unsteady phenomena while developing approaches to model them. The basis of the low-order model developed here is unsteady thin-airfoil theory. A time-stepping approach is used to solve for the vorticity on an airfoil camberline, allowing for large amplitudes and nonplanar wakes. On comparing lift coefficients from this method against data from CFD and experiments for some unsteady test cases, it is seen that the method predicts well so long as LEV formation does not occur and flow over the airfoil is attached. The formation of leading-edge vortices (LEVs) in unsteady flows is initiated by flow separation and the formation of a shear layer at the airfoil's leading edge. This phenomenon has been observed to have both detrimental (dynamic stall in helicopters) and beneficial (high-lift flight in insects) effects. To predict the formation of LEVs in unsteady flows, a Leading Edge Suction Parameter (LESP) is proposed. This parameter is calculated from inviscid theory and is a measure of the suction at the airfoil's leading edge. It

  12. Tissue culture characteristics of maize (Zea mays L.) haploid coleoptile sections.

    PubMed

    Jiang, L; Jing, G X; Li, X Y; Wang, X Q; Xing, Z; Deng, P K; Zhao, R G

    2015-12-08

    Doubled haploid (DH) technology, which is used for rapidly purifying genetic resources, is a key technology in modern maize breeding. The present study evaluated the tissue culture characteristics of maize haploid coleoptile sections, in order to provide a new way of haploid doubling. With 20 combinations of haploid coleoptile sections, obtained by hybridization within Reid, Tangsipingtou, and Term-tropical groups, as explants, we analyzed the induction and differentiation rate of callus, observed the number of root tip chromosomes in regenerated plants, and analyzed the pollen fertility. In addition, we used 47 SSR markers to analyze the genotypes of regenerated plants. The Reid and Tangsipingtou groups had significantly higher induction rates of haploid coleoptile callus compared to the Term-tropical group. Fifteen haploid plants were obtained which had 10 chromosomes in the root tips as assessed by I-KI staining. It was also noticed that the pollen of pollinated anthers were partially fertile. The haploid plants had genetic stability and showed no variation. The Reid and Tangsipingtou groups had good culture characteristics of haploid coleoptile sections, while the Term-tropical group had poor culture characteristics. Genotypes of haploid plants generated by tissue culture were evidenced to come from recombinant types of parents. Thus, this study established a tissue culture system of maize haploid coleoptile.

  13. Preliminary Design and Evaluation of an Airfoil with Continuous Trailing-Edge Flap

    NASA Technical Reports Server (NTRS)

    Shen, Jinwei; Thornburgh, Robert P.; Kreshock, Andrew R.; Wilbur, Matthew L.; Liu, Yi

    2012-01-01

    This paper presents the preliminary design and evaluation of an airfoil with active continuous trailing-edge flap (CTEF) as a potential rotorcraft active control device. The development of structural cross-section models of a continuous trailing-edge flap airfoil is described. The CTEF deformations with MFC actuation are predicted by NASTRAN and UM/VABS analyses. Good agreement is shown between the predictions from the two analyses. Approximately two degrees of CTEF deflection, defined as the rotation angle of the trailing edge, is achieved with the baseline MFC-PZT bender. The 2D aerodynamic characteristics of the continuous trailing-edge flap are evaluated using a CFD analysis. The aerodynamic efficiency of a continuous trailing-edge flap is compared to that of a conventional discrete trailing-edge flap (DTEF). It is found that the aerodynamic characteristics of a CTEF are equivalent to those of a conventional DTEF with the same deflection angle but with a smaller flap chord. A fluid structure interaction procedure is implemented to predict the deflection of the continuous trailingedge flap under aerodynamic pressure. The reductions in CTEF deflection are overall small when aerodynamic pressure is applied: 2.7% reduction is shown with a CTEF deflection angle of two degrees and at angle of attack of six degrees. In addition, newly developed MFC-PMN actuator is found to be a good supplement to MFC-PZT when applied as the bender outside layers. A mixed MFC-PZT and MFC-PMN bender generates 3% more CTEF deformation than an MFC-PZT only bender and 5% more than an MFC-PMN only bender under aerodynamic loads.

  14. Modal characteristics in a single-nanowire cavity with a triangular cross section.

    PubMed

    Seo, Min-Kyo; Yang, Jin-Kyu; Jeong, Kwang-Yong; Park, Hong-Gyu; Qian, Fang; Ee, Ho-Seok; No, You-Shin; Leet, Yong-Hee

    2008-12-01

    In this study, the modal characteristics of a single-GaN nanowire cavity with a triangular cross section surrounded by air or located on a silicon dioxide substrate have been analyzed. Two transverse resonant modes, transverse electric-like and transverse magnetic-like modes, are dominantly excited for nanowire cavities that have a small cross-sectional size of <300 nm and length of 10 microm. Using the three-dimensional finite-difference time-domain simulation method, quality factors, confinement factors, single-mode conditions, and far-field emission patterns are investigated for a nanowire cavity as a function of one length of the triangular cross section. The results of these simulations provide information that will be vital for the design and development of efficient nanowire lasers and light sources in ultracompact nanophotonic integrated circuits.

  15. Sedimentological and mineralogical characteristics of the Pleistocene loess/paleosol sections in the Eastern Croatia

    NASA Astrophysics Data System (ADS)

    Galović, Lidija

    2016-03-01

    In the Eastern Croatia impressive loess-paleosol successions up to 30 m thick are exposed. In the Zmajevac I section three paleosols are intercalated in loess while in the Zmajevac, Erdut and Šarengrad sections there are four paleosols are intercalated in loess. IRSL age estimates of 17.8 ± 1.9 and 217 ± 22 ka. In all investigated sections, alluvial sediments are intercalated in the loess deposits, indicating periods of fluvial activity. Strongly abraded typical aeolian spherical grains characterized by pitted well-rounded surface that was developed during transportation have original crystal surface almost destroyed. Surface of quartz grains preserves micro textures characteristic for all transport medias that it has been exposed to. However, muscovite grain surface enable successful distinguishing if the last transport was by wind or by aquatic media. Characteristic of all horizons with muscovite as a dominant mineral is recent settling of organisms. Beside the Danube, Drava and Sava River flood plains, part of the analyzed sediments also originates from regional Tertiary sediments which are rich in granite (as a muscovite-bearing rock), indicating the local influence. Enrichment of pyroxenes in the Šarengrad section points to the Dinaride Ophiolite Zone as its source of origin eroded by the Sava River southern tributaries. Šarengrad section is the southernmost among the analyzed sections and the southern edge of the Carpathian Basin. Thus, beside the Alpine region, the mineral composition is influenced by minerals from the Dinaride Ophiolite Zone in Bosnia. Warming periods are not represented just by paleosols, but also with laminated alluvial sediments.

  16. Full-scale semispan tests of a business-jet wing with a natural laminar flow airfoil

    NASA Technical Reports Server (NTRS)

    Hahne, David E.; Jordan, Frank L., Jr.

    1991-01-01

    A full-scale semispan model was investigated to evaluate and document the low-speed, high-lift characteristics of a business-jet class wing that utilized the HSNLF(1)-0213 airfoil section and a single-slotted flap system. Also, boundary-layer transition effects were examined, a segmented leading-edge droop for improved stall/spin resistance was studied, and two roll-controlled devices were evaluated. The wind-tunnel investigation showed that deployment of single-slotted, trailing-edge flap was effective in providing substantial increments in lift required for takeoff and landing performance. Fixed-transition studies to investigate premature tripping of the boundary layer indicated no adverse effects in lift and pitching-moment characteristics for either the cruise or landing configuration. The full-scale results also suggested the need to further optimize the leading-edge droop design that was developed in the subscale tests.

  17. Ice Accretions on Modern Airfoils Investigated

    NASA Technical Reports Server (NTRS)

    Addy, Harold E., Jr.

    2000-01-01

    The Icing Branch at the NASA Glenn Research Center at Lewis Field initiated and conducted the Modern Airfoils Ice Accretions project to identify ice shapes and determine their effects on the aerodynamic performance of aircraft, particularly on lift and drag. Previous aircraft ice shape and performance documentation focused on a few, older airfoils. This permitted more basic studies of the ice accretion process to be undertaken. However, having established both a working data base of ice shapes and the capability to predict these shapes for basic airfoils, questions arose about how ice might accrete differently on airfoils more representative of those being designed and flown on various aircraft today. Similarly, information about how these ice shapes would affect aerodynamic performance was needed.

  18. Low speed airfoil design and analysis

    NASA Technical Reports Server (NTRS)

    Eppler, R.; Somers, D. M.

    1979-01-01

    A low speed airfoil design and analysis program was developed which contains several unique features. In the design mode, the velocity distribution is not specified for one but many different angles of attack. Several iteration options are included which allow the trailing edge angle to be specified while other parameters are iterated. For airfoil analysis, a panel method is available which uses third-order panels having parabolic vorticity distributions. The flow condition is satisfied at the end points of the panels. Both sharp and blunt trailing edges can be analyzed. The integral boundary layer method with its laminar separation bubble analog, empirical transition criterion, and precise turbulent boundary layer equations compares very favorably with other methods, both integral and finite difference. Comparisons with experiment for several airfoils over a very wide Reynolds number range are discussed. Applications to high lift airfoil design are also demonstrated.

  19. Turbine airfoil to shroud attachment method

    SciTech Connect

    Campbell, Christian X; Kulkarni, Anand A; James, Allister W; Wessell, Brian J; Gear, Paul J

    2014-12-23

    Bi-casting a platform (50) onto an end portion (42) of a turbine airfoil (31) after forming a coating of a fugitive material (56) on the end portion. After bi-casting the platform, the coating is dissolved and removed to relieve differential thermal shrinkage stress between the airfoil and platform. The thickness of the coating is varied around the end portion in proportion to varying amounts of local differential process shrinkage. The coating may be sprayed (76A, 76B) onto the end portion in opposite directions parallel to a chord line (41) of the airfoil or parallel to a mid-platform length (80) of the platform to form respective layers tapering in thickness from the leading (32) and trailing (34) edges along the suction side (36) of the airfoil.

  20. Airfoil for a gas turbine

    DOEpatents

    Liang, George [Palm City, FL

    2011-01-18

    An airfoil is provided for a gas turbine comprising an outer structure comprising a first wall, an inner structure comprising a second wall spaced relative to the first wall such that a cooling gap is defined between at least portions of the first and second walls, and seal structure provided within the cooling gap between the first and second walls for separating the cooling gap into first and second cooling fluid impingement gaps. An inner surface of the second wall may define an inner cavity. The inner structure may further comprise a separating member for separating the inner cavity of the inner structure into a cooling fluid supply cavity and a cooling fluid collector cavity. The second wall may comprise at least one first impingement passage, at least one second impingement passage, and at least one bleed passage.

  1. Transonic airfoil design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

    A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

  2. Modeling and Grid Generation of Iced Airfoils

    NASA Technical Reports Server (NTRS)

    Vickerman, Mary B.; Baez, Marivell; Braun, Donald C.; Hackenberg, Anthony W.; Pennline, James A.; Schilling, Herbert W.

    2007-01-01

    SmaggIce Version 2.0 is a software toolkit for geometric modeling and grid generation for two-dimensional, singleand multi-element, clean and iced airfoils. A previous version of SmaggIce was described in Preparing and Analyzing Iced Airfoils, NASA Tech Briefs, Vol. 28, No. 8 (August 2004), page 32. To recapitulate: Ice shapes make it difficult to generate quality grids around airfoils, yet these grids are essential for predicting ice-induced complex flow. This software efficiently creates high-quality structured grids with tools that are uniquely tailored for various ice shapes. SmaggIce Version 2.0 significantly enhances the previous version primarily by adding the capability to generate grids for multi-element airfoils. This version of the software is an important step in streamlining the aeronautical analysis of ice airfoils using computational fluid dynamics (CFD) tools. The user may prepare the ice shape, define the flow domain, decompose it into blocks, generate grids, modify/divide/merge blocks, and control grid density and smoothness. All these steps may be performed efficiently even for the difficult glaze and rime ice shapes. Providing the means to generate highly controlled grids near rough ice, the software includes the creation of a wrap-around block (called the "viscous sublayer block"), which is a thin, C-type block around the wake line and iced airfoil. For multi-element airfoils, the software makes use of grids that wrap around and fill in the areas between the viscous sub-layer blocks for all elements that make up the airfoil. A scripting feature records the history of interactive steps, which can be edited and replayed later to produce other grids. Using this version of SmaggIce, ice shape handling and grid generation can become a practical engineering process, rather than a laborious research effort.

  3. Transonic Airfoils with a Given Pressure Distribution,

    DTIC Science & Technology

    1981-06-01

    erovse sidst necesosar mod Ideatify b lock mmb)L An inverse design procedure for airfoils, based on hodograph techniques, has been developed. For...w L-:- " " -- - r- L i -- _ 9 ABSTRACT An inverse design procedure for airfoils, based on hodograph tech...generated in the hodograph plane by Nieuwand,5 Bauer, Garabedian and Korn,6 Boerstoel and Huizing,7 and Sobieczky.8 More recently, the development of

  4. Unsteady Pressure Distributions on Airfoils in Cascade.

    DTIC Science & Technology

    1980-04-01

    of thin airfoil theory has been used by Henderson (-ftj’ and Bruce (1-7-)’to derive expressions for the unsteady response which includes the cascade...model in conjunction with the assumptions of thin airfoil theory has been used by Henderson (16) and Bruce (17) to derive expressions for the unsteady...effect, that is, a sharp change in the unsteady lift when the disturbance wavelength equals the blade spacing. Bruce (19) further extends this theory to

  5. Determination of tire cross-sectional geometric characteristics from a digitally scanned image

    NASA Technical Reports Server (NTRS)

    Danielson, Kent T.

    1995-01-01

    A semi-automated procedure is described for the accurate determination of geometrical characteristics using a scanned image of the tire cross-section. The procedure can be useful for cases when CAD drawings are not available or when a description of the actual cured tire is desired. Curves representing the perimeter of the tire cross-section are determined by an edge tracing scheme, and the plyline and cord-end positions are determined by locations of color intensities. The procedure provides an accurate description of the perimeter of the tire cross-section and the locations of plylines and cord-ends. The position, normals, and curvatures of the cross-sectional surface are included in this description. The locations of the plylines provide the necessary information for determining the ply thicknesses and relative position to a reference surface. Finally, the locations of the cord-ends provide a means to calculate the cord-ends per inch (epi). Menu driven software has been developed to facilitate the procedure using the commercial code, PV-Wave by Visual Numerics, Inc., to display the images. From a single user interface, separate modules are executed for image enhancement, curve fitting the edge trace of the cross-sectional perimeter, and determining the plyline and cord-end locations. The code can run on SUN or SGI workstations and requires the use of a mouse to specify options or identify items on the scanned image.

  6. Propulsion by active and passive airfoil oscillation

    NASA Astrophysics Data System (ADS)

    Mackowski, A. W.; Williamson, C. H. K.

    2013-11-01

    Oscillating airfoils have been the subject of much research both as a mechanism of propulsion in engineering devices as well as a model of understanding how fish, birds, and insects produce thrust and maneuvering forces. Additionally, the jet or wake generated by an oscillating airfoil exhibits a multitude of vortex patterns, which are an interesting study in their own right. We present PIV measurements of the vortex flow behind an airfoil undergoing controlled pitching oscillations at moderate Reynolds number. As a method of propulsion, oscillating foils have been found to be capable performers when undergoing both pitching and heaving motions [Anderson et al. 1998]. While an airfoil undergoing only pitching motion is a relatively inefficient propulsor, we examine the effect of adding passive dynamics to the system: for example, actuated pitching with a passive spring in the heave direction. Practically speaking, a mechanical system with such an arrangement has the potential to reduce the cost and complexity of an oscillating airfoil propulsor. To study an airfoil undergoing both active and passive motion, we employ our ``cyber-physical fluid dynamics'' technique [Mackowski & Williamson, 2011] to simulate the effects of passive dynamics in a physical experiment.

  7. Numerical investigation of multi-element airfoils

    NASA Technical Reports Server (NTRS)

    Cummings, Russell M.

    1993-01-01

    The flow over multi-element airfoils with flat-plate lift-enhancing tabs was numerically investigated. Tabs ranging in height from 0.25 percent to 1.25 percent of the reference airfoil chord were studied near the trailing edge of the main-element. This two-dimensional numerical simulation employed an incompressible Navier-Stokes solver on a structured, embedded grid topology. New grid refinements were used to improve the accuracy of the solution near the overlapping grid boundaries. The effects of various tabs were studied at a constant Reynolds number on a two-element airfoil with a slotted flap. Both computed and measured results indicated that a tab in the main-element cove improved the maximum lift and lift-to-drag ratio relative to the baseline airfoil without a tab. Computed streamlines revealed that the additional turning caused by the tab may reduce the amount of separated flow on the flap. A three-element airfoil was also studied over a range of Reynolds numbers. For the optimized flap rigging, the computed and measured Reynolds number effects were similar. When the flap was moved from the optimum position, numerical results indicated that a tab may help to reoptimize the airfoil to within 1 percent of the optimum flap case.

  8. Simulation of a Controlled Airfoil with Jets

    NASA Technical Reports Server (NTRS)

    Allan, Brian G.; Holt, Maurice; Packard, Andrew

    1997-01-01

    Numerical simulations of a two-dimensional airfoil, controlled by an applied moment in pitch and an airfoil controlled by jets, were investigated. These simulations couple the Reynolds-averaged Navier-Stokes equations and Euler's equations of rigid body motion, with an active control system. Controllers for both systems were designed to track altitude commands and were evaluated by simulating a closed-loop altitude step response using the coupled system. The airfoil controlled by a pitching moment used an optimal state feedback controller. A closed-loop simulation, of the airfoil with an applied moment, showed that the trajectories compared very well with quasi-steady aerodynamic theory, providing a measure of validation. The airfoil with jets used a controller designed by robust control methods. A linear plant model for this system was identified using open-loop data generated by the nonlinear coupled system. A closed-loop simulation of the airfoil with jets, showed good tracking of an altitude command. This simulation also showed oscillations in the control input as a result of dynamics not accounted for in the control design. This research work demonstrates how computational fluid dynamics, coupled with rigid body dynamics, and a control law can be used to prototype control systems in problematic nonlinear flight regimes.

  9. Computational studies on small wind turbine performance characteristics

    NASA Astrophysics Data System (ADS)

    Karthikeyan, N.; Suthakar, T.

    2016-10-01

    To optimize the selection of suitable airfoils for small wind turbine applications, computational investigation on aerodynamic characteristics of low Re airfoils MID321a, MID321d, SG6040, SG6041, SG6042 and SG6043 are carried out for the Reynolds number range of (0.5- 2)×105. The BEM method is used to determine the power coefficient of the rotor from the airfoil characteristics; in addition, the blade parameters like chord and twist are also determined. The newly designed MID321a airfoil shows better aerodynamic performance and maximum power coefficient as compared with other investigated airfoils for wider operating ranges.

  10. Isolated and cascade airfoils with prescribed velocity distribution

    NASA Technical Reports Server (NTRS)

    Goldstein, Arthur W; Jerison, Meyer

    1947-01-01

    An exact solution of the problem of designing an airfoil with a prescribed velocity distribution on the suction surface in a given uniform flow of an incompressible perfect fluid is obtained by replacing the boundary of the airfoil by vortices. By this device, a method of solution is developed that is applicable both to isolated airfoils and to airfoils in cascade. The conformal transformation of the designed airfoil into a circle can then be obtained and the velocity distribution at any angle of attack computed. Numerical illustrations of the method are given for the airfoil in cascade.

  11. Comparative Drag Measurements at Transonic Speeds of Rectangular Sweptback NACA 65-009 Airfoils Mounted on a Freely Falling Body

    NASA Technical Reports Server (NTRS)

    Mathews, Charles W; Thompson, Jim Rogers

    1950-01-01

    Directly comparable drag measurements have been made of an airfoil with a conventional rectangular plan form and an airfoil with a sweptback plan form mounted on freely falling bodies. Both airfoils had NACA 65-009 sections and were identical in span, frontal area, and chord perpendicular to the leading edge. The sweptback plan form incorporated a sweepback angle of 45 degrees. The data obtained have been used to establish the relation between the airfoil drag coefficients and the free-stream Mach number over a range of Mach numbers from 0.90 to 1.27. The results of the measurements indicate that the drag of the sweptback plan form is less than 0.3 that of the rectangular plan form at a Mach number of 1.00 and is less than 0.4 that at a Mach number of 1.20.

  12. 2D CFD Analysis of an Airfoil with Active Continuous Trailing Edge Flap

    NASA Astrophysics Data System (ADS)

    Jaksich, Dylan; Shen, Jinwei

    2014-11-01

    Efficient and quieter helicopter rotors can be achieved through on-blade control devices, such as active Continuous Trailing-Edge Flaps driven by embedded piezoelectric material. This project aims to develop a CFD simulation tool to predict the aerodynamic characteristics of an airfoil with CTEF using open source code: OpenFOAM. Airfoil meshes used by OpenFOAM are obtained with MATLAB scripts. Once created it is possible to rotate the airfoil to various angles of attack. When the airfoil is properly set up various OpenFOAM properties, such as kinematic viscosity and flow velocity, are altered to achieve the desired testing conditions. Upon completion of a simulation, the program gives the lift, drag, and moment coefficients as well as the pressure and velocity around the airfoil. The simulation is then repeated across multiple angles of attack to give full lift and drag curves. The results are then compared to previous test data and other CFD predictions. This research will lead to further work involving quasi-steady 2D simulations incorporating NASTRAN to model aeroelastic deformation and eventually to 3D aeroelastic simulations. NSF ECE Grant #1358991 supported the first author as an REU student.

  13. Investigation to optimize the passive shock wave/boundary layer control for supercritical airfoil drag reduction

    NASA Technical Reports Server (NTRS)

    Nagamatsu, H. T.; Dyer, R.

    1984-01-01

    The passive shock wave/boundary layer control for reducing the drag of 14%-thick supercritical airfoil was investigated in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel with and without the top wall insert at transonic Mach numbers. Top wall insert was installed to increase the flow Mach number to 0.90 with the model mounted on the test section bottom wall. Various porous surfaces with a cavity underneath were positioned on the area of the airfoil where the shock wave occurs. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and enthropy increase hrough the shock wave. 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 2.8% uniform porosity the normal shock wave for the solid surface was changed to a lambda shock wave, and the wake impact pressure data indicate a drag coefficient reduction as much as 45% lower than for the solid surface airfoil at high transonic Mach numbers.

  14. Flow past a self-oscillating airfoil with two degrees of freedom: measurements and simulations

    NASA Astrophysics Data System (ADS)

    Šidlof, Petr; Štěpán, Martin; Vlček, Václav; Řidký, Václav; Šimurda, David; Horáček, Jaromír

    2014-03-01

    The paper focuses on investigation of the unsteady subsonic airflow past an elastically supported airfoil for subcritical flow velocities and during the onset of the flutter instability. A physical model of the NACA0015 airfoil has been designed and manufactured, allowing motion with two degrees of freedom: pitching (rotation about the elastic axis) and plunging (vertical motion). The structural mass and stiffness matrix can be tuned to certain extent, so that the natural frequencies of the two modes approach as needed. The model was placed in the measuring section of the wind tunnel in the aerodynamic laboratory of the Institute of Thermomechanics in Nový Knín, and subjected to low Mach number airflow up to the flow velocities when self-oscillation reach amplitudes dangerous for the structural integrity of the model. The motion of the airfoil was registered by a high-speed camera, with synchronous measurement of the mechanic vibration and discrete pressure sensors on the surface of the airfoil. The results of the measurements are presented together with numerical simulation results, based on a finite volume CFD model of airflow past a vibrating airfoil.

  15. Development of heat flux sensors for turbine airfoils

    NASA Astrophysics Data System (ADS)

    Atkinson, William H.; Cyr, Marcia A.; Strange, Richard R.

    1985-10-01

    The objectives of this program are to develop heat flux sensors suitable for installation in hot section airfoils of advanced aircraft turbine engines and to experimentally verify the operation of these heat flux sensors in a cylinder in a cross flow experiment. Embedded thermocouple and Gardon gauge sensors were developed and fabricated into both blades and vanes. These were then calibrated using a quartz lamp bank heat source and finally subjected to thermal cycle and thermal soak testing. These sensors were also fabricated into cylindrical test pieces and tested in a burner exhaust to verify heat flux measurements produced by these sensors. The results of the cylinder in cross flow tests are given.

  16. Evaluation of a research circulation control airfoil using Navier-Stokes methods

    NASA Technical Reports Server (NTRS)

    Shrewsbury, George D.

    1987-01-01

    The compressible Reynolds time averaged Navier-Stokes equations were used to obtain solutions for flows about a two dimensional circulation control airfoil. The governing equations were written in conservation form for a body-fitted coordinate system and solved using an Alternating Direction Implicit (ADI) procedure. A modified algebraic eddy viscosity model was used to define the turbulent characteristics of the flow, including the wall jet flow over the Coanda surface at the trailing edge. Numerical results are compared to experimental data obtained for a research circulation control airfoil geometry. Excellent agreement with the experimental results was obtained.

  17. Computer investigations of the turbulent flow around a NACA2415 airfoil wind turbine

    NASA Astrophysics Data System (ADS)

    Driss, Zied; Chelbi, Tarek; Abid, Mohamed Salah

    2015-12-01

    In this work, computer investigations are carried out to study the flow field developing around a NACA2415 airfoil wind turbine. The Navier-Stokes equations in conjunction with the standard k-ɛ turbulence model are considered. These equations are solved numerically to determine the local characteristics of the flow. The models tested are implemented in the software "SolidWorks Flow Simulation" which uses a finite volume scheme. The numerical results are compared with experiments conducted on an open wind tunnel to validate the numerical results. This will help improving the aerodynamic efficiency in the design of packaged installations of the NACA2415 airfoil type wind turbine.

  18. Effects of laminar separation bubbles and turbulent separation on airfoil stall

    SciTech Connect

    Dini, P.; Coiro, D.P.

    1997-12-31

    An existing two-dimensional, interactive, stall prediction program is extended by improving its laminar separation bubble model. The program now accounts correctly for the effects of the bubble on airfoil performance characteristics when it forms at the mid-chord and on the leading edge. Furthermore, the model can now predict bubble bursting on very sharp leading edges at high angles of attack. The details of the model are discussed in depth. Comparisons of the predicted stall and post-stall pressure distributions show excellent agreement with experimental measurements for several different airfoils at different Reynolds numbers.

  19. Icing Test Results on an Advanced Two-Dimensional High-Lift Multi-Element Airfoil

    NASA Technical Reports Server (NTRS)

    Shin, Jaiwon; Wilcox, Peter; Chin, Vincent; Sheldon, David

    1994-01-01

    An experimental study has been conducted to investigate ice accretions on a high-lift, multi-element airfoil in the Icing Research Tunnel at the NASA Lewis Research Center. The airfoil is representative of an advanced transport wing design. The experimental work was conducted as part of a cooperative program between McDonnell Douglas Aerospace and the NASA Lewis Research Center to improve current understanding of ice accretion characteristics on the multi-element airfoil. The experimental effort also provided ice shapes for future aerodynamic tests at flight Reynolds numbers to ascertain high-lift performance effects. Ice shapes documented for a landing configuration over a variety of icing conditions are presented along with analyses.

  20. Atmospheric performance of the special-purpose Solar Energy Research Institute (SERI) thin-airfoil family

    SciTech Connect

    Tangler, J; Smith, B; Jager, D; Olsen, T

    1990-09-01

    The Solar Energy Research Institute (SERI), in cooperation with SeaWest Energy Group, has completed extensive atmospheric testing of the special-purpose SERI thin-airfoil family during the 1990 wind season. The purpose of this test program was to experimentally verify the predicted performance characteristics of the thin-airfoil family on a geometrically optimized blade, and to compare it to original-equipment blades under atmospheric wind conditions. The tests were run on two identical Micon 65/13 horizontal-axis wind turbines installed side-by-side in a wind farm. The thin-airfoil family 7.96 m blades were installed on one turbine, and AeroStar 7.41 m blades were installed on the other. This paper presents final performance results of the side-by-side comparative field test for both clean and dirty blade conditions. 7 refs., 11 figs., 1 tab.

  1. An airfoil flutter model suspension system to accommodate large static transonic airloads

    NASA Technical Reports Server (NTRS)

    Reed, W. H., III

    1985-01-01

    A pitch/plunge flutter model suspension system and associated two-dimensional MBB-A3 airfoil models is described. The system is designed for installation in the Langley 6-by-19-inch and 6-by-18-inch transonic blowdown wind tunnels to enable systematic study of the transonic flutter characteristics and static pressure distributions of supercritical airfoils at transonic Mach numbers. A compound spring suspension concept is introduced which simultaneously meets requirements for low plunge-mode stiffness, lightweight suspended model, and large steady lift due to angle of attack without the need for excessive static deflections of the plunge spring. The system features variable pitch and plunge frequencies, changeable airfoil rotation axes, and a self aligning control system to maintain a constant mean position of the model with changing airload.

  2. Influence of horizontally curved roadway section characteristics on motorcycle-to-barrier crash frequency.

    PubMed

    Gabauer, Douglas J; Li, Xiaolong

    2015-04-01

    The purpose of this study was to investigate motorcycle-to-barrier crash frequency on horizontally curved roadway sections in Washington State using police-reported crash data linked with roadway data and augmented with barrier presence information. Data included 4915 horizontal curved roadway sections with 252 of these sections experiencing 329 motorcycle-to-barrier crashes between 2002 and 2011. Negative binomial regression was used to predict motorcycle-to-barrier crash frequency using horizontal curvature and other roadway characteristics. Based on the model results, the strongest predictor of crash frequency was found to be curve radius. This supports a motorcycle-to-barrier crash countermeasure placement criterion based, at the very least, on horizontal curve radius. With respect to the existing horizontal curve criterion of 820 feet or less, curves meeting this criterion were found to increase motorcycle-to-barrier crash frequency rate by a factor of 10 compared to curves not meeting this criterion. Other statistically significant predictors were curve length, traffic volume and the location of adjacent curves. Assuming curves of identical radius, the model results suggest that longer curves, those with higher traffic volume, and those that have no adjacent curved sections within 300 feet of either curve end would likely be better candidates for a motorcycle-to-barrier crash countermeasure.

  3. Application of numerical optimization to the design of advanced supercritical airfoils

    NASA Technical Reports Server (NTRS)

    Johnson, R. R.; Hicks, R. M.

    1979-01-01

    An application of numerical optimization to the design of advanced airfoils for transonic aircraft showed that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while minimizing the drag coefficient at the design Mach number. This multiple design-point numerical optimization has been implemented with the use of airfoil shape functions which permit a wide range of attainable profiles during the optimization process. Analytical data for the starting airfoil shape, a single design-point optimized shape, and a double design-point optimized shape are presented. Experimental data obtained in the NASA Ames two-by two-foot wind tunnel are also presented and discussed.

  4. Interference method for obtaining the potential flow past an arbitrary cascade of airfoils

    NASA Technical Reports Server (NTRS)

    Katzoff, S; Finn, Robert S; Laurence, James C

    1947-01-01

    A procedure is presented for obtaining the pressure distribution on an arbitrary airfoil section in cascade in a two-dimensional, incompressible, and nonviscous flow. The method considers directly the influence on a given airfoil of the rest of the cascade and evaluates this interference by an iterative process, which appeared to converge rapidly in the cases tried (about unit solidity, stagger angles of 0 degree and 45 degrees). Two variations of the basic interference calculations are described. One, which is accurate enough for most purposes, involves the substitution of sources, sinks, and vortices for the interfering airfoils; the other, which may be desirable for the final approximation, involves a contour integration. The computations are simplified by the use of a chart presented by Betz in a related paper. Illustrated examples are included.

  5. Discussion of test results in the design of laminar airfoils for competition gliders

    NASA Technical Reports Server (NTRS)

    Ostrowski, J.; Skrzynski, S.; Litwinczyk, M.

    1980-01-01

    The deformation of flow in the boundary layer and the local separation of a laminar layer (laminar bubbles) from various airfoils were investigated. These phenomena were classified and their influence is discussed. Various aerodynamic characteristics are discussed and the principles for prescribing pressure distribution to attain a high value of c sub z max with a possibly low drag coefficient are described.

  6. Experimental Results with Airfoils Tested in the High-speed Tunnel at Guidonia

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1940-01-01

    The results are presented of a triple series of tests using force measurements, pressure-distribution measurements, and air flow photographs on airfoil sections suitably selected so that comparison could be made between the experimental and theoretical results. The comparison with existing theory is followed by a discussion of the divergences found, and an attempt is made to find their explanation.

  7. Experimental Investigation of a 2D Supercritical Circulation-Control Airfoil Using Particle Image Velocimetry

    NASA Technical Reports Server (NTRS)

    Jones, Gregory S.; Yao, Chung-Sheng; Allan, Brian G.

    2006-01-01

    Recent efforts in extreme short takeoff and landing aircraft configurations have renewed the interest in circulation control wing design and optimization. The key to accurately designing and optimizing these configurations rests in the modeling of the complex physics of these flows. This paper will highlight the physics of the stagnation and separation regions on two typical circulation control airfoil sections.

  8. Improved design of subcritical and supercritical cascades using complex characteristics and boundary-layer correction

    NASA Technical Reports Server (NTRS)

    Sanz, J. M.

    1984-01-01

    The method of complex characteristics and hodograph transformation for the design of shockless airfoils was extended to design supercritical cascades with high solidities and large inlet angles. This capability was achieved by introducing a conformal mapping of the hodograph domain onto an ellipse and expanding the solution in terms of Tchebycheff polynomials. A computer code was developed based on this idea. A number of airfoils designed with the code are presented. Various supercritical and subcritical compressor, turbine and propeller sections are shown. The lag-entrainment method for the calculation of a turbulent boundary layer was incorporated to the inviscid design code. The results of this calculation are shown for the airfoils described. The elliptic conformal transformation developed to map the hodograph domain onto an ellipse can be used to generate a conformal grid in the physical domain of a cascade of airfoils with open trailing edges with a single transformation. A grid generated with this transformation is shown for the Korn airfoil. Previously announced in STAR as N83-24474

  9. Ocean Thermohydrodynamic Characteristics Obtained By Synthesizing A Woce Hydrographic Section With Climatic Data.

    NASA Astrophysics Data System (ADS)

    Sarkisyan, A.; Levitus, S.

    As a main chain for synthesizing a WOCE "synoptic" hydrographic section with the neighbouring climatic data serves a modern high resolution 3DPEM. The calculation procedure is as follows: 1.Replacing climatic T,S data by observed unfiltered hydro- logic section; 2.Making a short model time diagnostic-prognostic investigation by a special freezing-defreezing method. As a result we obtain mutually adjusted synoptic fields of all physical characteristics at narrow strip embracing the WOCE section.This method, after some preliminary test, was validated by using the observed data of the R/V "AKADEMIK Sergei Vavilov" in her 11th cruise to the Barents Sea on Sept. 1997 (Sarkisyan and Sidorova, 1998). More comprehensive calculations were made by using the WOCE A05 section data, 24.5N in the North Atlantic (S.Levitus and Sarkisyan, 2001) The main results of this work are as follows. (1) A clearly defined asymmetry occurs in the spatial distribution of the effect of inlaying syn-optic data. Namely, (a) the effect of synoptic information decays in the meridional direction at a distance of a few degrees from the section, and it is obvious that this effect also decays with depth because of baroclinicity; (b) in the zonal direction, the effect is propagated westward under the action of Rossby waves and other factors, and, in the continental slope (east of 80W), this effect enhances through die JEBAR and is propagated north- ward by the Antilles Current and further by the Gulf Stream. This western intensifi- cation is clearly reflected in the mass and heat transport values. (2) Due to eddies and other-type spatial variability, as well as to the action of the JEBAR, meridional trans- port processes in the region of the section are an order of magnitude stronger than those in the case of computations based on climatic data, which is confirmed by direct observations in the Bahamas region. From computations, it is inferred that the western intensification and a high degree of

  10. Program manual for the Eppler airfoil inversion program

    NASA Technical Reports Server (NTRS)

    Thomson, W. G.

    1975-01-01

    A computer program is described for calculating the profile of an airfoil as well as the boundary layer momentum thickness and energy form parameter. The theory underlying the airfoil inversion technique developed by Eppler is discussed.

  11. AirfoilPrep.py Documentation: Release 0.1.0

    SciTech Connect

    Ning, S. A.

    2013-09-01

    AirfoilPrep.py provides functionality to preprocess aerodynamic airfoil data. Essentially, the module is an object oriented version of the AirfoilPrep spreadsheet with additional functionality and is written in the Python language. It allows the user to read in two-dimensional aerodynamic airfoil data, apply three-dimensional rotation corrections for wind turbine applications, and extend the datato very large angles of attack. This document discusses installation, usage, and documentation of the module.

  12. Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes

    NASA Technical Reports Server (NTRS)

    Ivey, Margaret F

    1945-01-01

    Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corresponding cutouts made in wing were tested for various flap deflections, chord-wise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80 percent wing chord behind leading edge. Airflow was smooth and buffeting negligible.

  13. Forcing function effects on unsteady aerodynamic gust response. II - Low solidity airfoil row response

    NASA Technical Reports Server (NTRS)

    Henderson, Gregory H.; Fleeter, Sanford

    1992-01-01

    The paper investigates the fundamental gust modeling assumption on the basis of a series of experiments performed in the Purdue Annular Cascade Research Facility. The unsteady period flow field is generated by rotating flows of perforated plates and airfoil cascades, with the resulting unsteady periodic chordwise pressure response of a downstream low solidity stator row determined by miniature pressure transducers embedded within selected airfoils. When the forcing function exhibited the characteristics of a linear-theory gust, the resulting response on the downstream stator airfoils was in excellent agreement with the linear-theory models. When the forcing function did not exhibit linear-theory gust characteristics, the resulting unsteady aerodynamic response of the downstream stators was much more complex and correlated poorly with the linear-theory gust predictions. It is shown that the forcing function generator significantly affects the resulting gust response, with the complexity of the response characteristics increasing from the perforated-plate to the airfoil-cascade forcing functions.

  14. Coating-Substrate Systems for Thermomechanically Durable Turbine Airfoils

    DTIC Science & Technology

    2015-06-30

    Technical Report 4. TITLE AND SUBTITLE Coating - Substrate Systems for Thermomechanically Durable Turbine Airfoils 6. AUTHOR(S) Dr. Tresa Pollock 3...Thermomechanically Durable Turbine Airfoils Final Report ONRGrant#N00014-l 1-1-0616 Technical Contact (Principal Investigator) Tresa M. Pollock Materials...Substrate Systems for Thermomechanically Durable Turbine Airfoils 1. Summary In the severe operating environments encountered in Naval ship

  15. Deposition characteristics of methamphetamine and amphetamine in fingernail clippings and hair sections.

    PubMed

    Lin, Dong-Liang; Yin, Rea-Ming; Liu, Hsiu-Chuan; Wang, Chung-Yi; Liu, Ray H

    2004-09-01

    Fingernail clippings collected from 97 consenting females, who admitted amphetamines and/or opiates use and are currently under treatment, were quantitatively analyzed for the presence of methamphetamine and amphetamine. Sixty-two subjects were found positive for methamphetamine/amphetamine. Paired nail-hair specimens were collected from 6 of these subjects for a 12-week period and analyzed to determine the duration of detectability and deposition characteristics of amphetamines in fingernails; whether data derived from the analysis of nail clippings and hair sections are reflective of drug use patterns; and whether there is a relationship between the analytical data derived from the paired nail-hair specimens. Typical sample pre-treatment procedures and GC-MS protocols were evaluated to establish the validity of various analytical parameters and to ensure that the resulting data can be properly interpreted. Major findings include 1. Methamphetamine was found in the nails of 62 subjects collected in Week 0. The distribution of methamphetamine concentrations (ng/mg) in these nail samples are range, 0.46-61.50; mean, 9.96; and standard deviation: 13.33. The corresponding data for amphetamine are < 0.20-5.42, 0.93, and 1.01, respectively. 2. Sectional analyses of hair samples collected from 6 subjects in Week 0 show methamphetamine concentrations peak at different distances from the root. 3. The concentrations of methamphetamine and amphetamine in nail clippings are generally lower than the first 1.5-cm section of hair samples collected at the same time from the same individual. 4. Amphetamine/ methamphetamine concentration ratios in nail clippings and hair samples are comparable. 5. Methamphetamine concentration in the nail clippings collected at Weeks 0, 4, 8, and 12 decreases in a pattern similar to that exhibited by the first 1.5-cm sections of the hair samples collected at the same time.

  16. Epidemiology and Reporting Characteristics of Systematic Reviews of Biomedical Research: A Cross-Sectional Study

    PubMed Central

    Page, Matthew J.; Shamseer, Larissa; Altman, Douglas G.; Tetzlaff, Jennifer; Tricco, Andrea C.; Catalá-López, Ferrán; Li, Lun; Reid, Emma K.; Sarkis-Onofre, Rafael; Moher, David

    2016-01-01

    Background Systematic reviews (SRs) can help decision makers interpret the deluge of published biomedical literature. However, a SR may be of limited use if the methods used to conduct the SR are flawed, and reporting of the SR is incomplete. To our knowledge, since 2004 there has been no cross-sectional study of the prevalence, focus, and completeness of reporting of SRs across different specialties. Therefore, the aim of our study was to investigate the epidemiological and reporting characteristics of a more recent cross-section of SRs. Methods and Findings We searched MEDLINE to identify potentially eligible SRs indexed during the month of February 2014. Citations were screened using prespecified eligibility criteria. Epidemiological and reporting characteristics of a random sample of 300 SRs were extracted by one reviewer, with a 10% sample extracted in duplicate. We compared characteristics of Cochrane versus non-Cochrane reviews, and the 2014 sample of SRs versus a 2004 sample of SRs. We identified 682 SRs, suggesting that more than 8,000 SRs are being indexed in MEDLINE annually, corresponding to a 3-fold increase over the last decade. The majority of SRs addressed a therapeutic question and were conducted by authors based in China, the UK, or the US; they included a median of 15 studies involving 2,072 participants. Meta-analysis was performed in 63% of SRs, mostly using standard pairwise methods. Study risk of bias/quality assessment was performed in 70% of SRs but was rarely incorporated into the analysis (16%). Few SRs (7%) searched sources of unpublished data, and the risk of publication bias was considered in less than half of SRs. Reporting quality was highly variable; at least a third of SRs did not report use of a SR protocol, eligibility criteria relating to publication status, years of coverage of the search, a full Boolean search logic for at least one database, methods for data extraction, methods for study risk of bias assessment, a primary

  17. Compressor airfoil tip clearance optimization system

    DOEpatents

    Little, David A.; Pu, Zhengxiang

    2015-08-18

    A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.

  18. Options for Robust Airfoil Optimization under Uncertainty

    NASA Technical Reports Server (NTRS)

    Padula, Sharon L.; Li, Wu

    2002-01-01

    A robust optimization method is developed to overcome point-optimization at the sampled design points. This method combines the best features from several preliminary methods proposed by the authors and their colleagues. The robust airfoil shape optimization is a direct method for drag reduction over a given range of operating conditions and has three advantages: (1) it prevents severe degradation in the off-design performance by using a smart descent direction in each optimization iteration, (2) it uses a large number of spline control points as design variables yet the resulting airfoil shape does not need to be smoothed, and (3) it allows the user to make a tradeoff between the level of optimization and the amount of computing time consumed. For illustration purposes, the robust optimization method is used to solve a lift-constrained drag minimization problem for a two-dimensional (2-D) airfoil in Euler flow with 20 geometric design variables.

  19. Turbine airfoil fabricated from tapered extrusions

    DOEpatents

    Marra, John J

    2013-07-16

    An airfoil (30) and fabrication process for turbine blades with cooling channels (26). Tapered tubes (32A-32D) are bonded together in a parallel sequence, forming a leading edge (21), a trailing edge (22), and pressure and suction side walls (23, 24) connected by internal ribs (25). The tapered tubes may be extruded without camber to simplify the extrusion process, then bonded along matching surfaces (34), forming a non-cambered airfoil (28), which may be cambered in a hot forming process and cut (48) to length. The tubes may have tapered walls that are thinner at the blade tip (T1) than at the base (T2), reducing mass. A cap (50) may be attached to the blade tip. A mounting lug (58) may be forged (60) on the airfoil base and then machined, completing the blade for mounting in a turbine rotor disk.

  20. Comparative Study of Airfoil Flow Separation Criteria

    NASA Astrophysics Data System (ADS)

    Laws, Nick; Kahouli, Waad; Epps, Brenden

    2015-11-01

    Airfoil flow separation impacts a multitude of applications including turbomachinery, wind turbines, and bio-inspired micro-aerial vehicles. In order to achieve maximum performance, some devices operate near the edge of flow separation, and others use dynamic flow separation advantageously. Numerous criteria exist for predicting the onset of airfoil flow separation. This talk presents a comparative study of a number of such criteria, with emphasis paid to speed and accuracy of the calculations. We evaluate the criteria using a two-dimensional unsteady vortex lattice method, which allows for rapid analysis (on the order of seconds instead of days for a full Navier-Stokes solution) and design of optimal airfoil geometry and kinematics. Furthermore, dynamic analyses permit evaluation of dynamic stall conditions for enhanced lift via leading edge vortex shedding, commonly present in small flapping-wing flyers such as the bumblebee and hummingbird.

  1. Vortex noise from nonrotating cylinders and airfoils

    NASA Technical Reports Server (NTRS)

    Schlinker, R. H.; Amiet, R. K.; Fink, M. R.

    1976-01-01

    An experimental study of vortex-shedding noise was conducted in an acoustic research tunnel over a Reynolds-number range applicable to full-scale helicopter tail-rotor blades. Two-dimensional tapered-chord nonrotating models were tested to simulate the effect of spanwise frequency variation on the vortex-shedding mechanism. Both a tapered circular cylinder and tapered airfoils were investigated. The results were compared with data for constant-diameter cylinder and constant-chord airfoil models also tested during this study. Far-field noise, surface pressure fluctuations, and spanwise correlation lengths were measured for each configuration. Vortex-shedding noise for tapered cylinders and airfoils was found to contain many narrowband-random peaks which occurred within a range of frequencies corresponding to a predictable Strouhal number referenced to the maximum and minimum chord. The noise was observed to depend on surface roughness and Reynolds number.

  2. Near-wall serpentine cooled turbine airfoil

    SciTech Connect

    Lee, Ching-Pang

    2013-09-17

    A serpentine coolant flow path (54A-54G) formed by inner walls (50, 52) in a cavity (49) between pressure and suction side walls (22, 24) of a turbine airfoil (20A). A coolant flow (58) enters (56) an end of the airfoil, flows into a span-wise channel (54A), then flows forward (54B) over the inner surface of the pressure side wall, then turns behind the leading edge (26), and flows back along a forward part of the suction side wall, then follows a loop (54E) forward and back around an inner wall (52), then flows along an intermediate part of the suction side wall, then flows into an aft channel (54G) between the pressure and suction side walls, then exits the trailing edge (28). This provides cooling matched to the heating topography of the airfoil, minimizes differential thermal expansion, revives the coolant, and minimizes the flow volume needed.

  3. Effect of cross-sectional shape on the propagation characteristics of fundamental plasmon mode guided along Au and Ag nanowires

    NASA Astrophysics Data System (ADS)

    Jitender; Kumar, Arun

    2017-01-01

    We numerically examine the effect of the cross-sectional shape on the propagation characteristics of the fundamental surface plasmon polariton (SPP) mode guided along Au and Ag nanowires. It is observed that as the cross-sectional symmetry decreases from circular to triangular via pentagonal and square, the modal confinement, loss and dispersion increases. Further the difference in propagation characteristics between circular and other cross-sections increases rapidly as the wavelength of operation decreases. The study should be useful for selecting the cross-sectional shape of the nanowire for a particular application.

  4. Blowing Circulation Control on a Seaplane Airfoil

    NASA Astrophysics Data System (ADS)

    Guo, B. D.; Liu, P. Q.; Qu, Q. L.

    2011-09-01

    RANS simulations are presented for blowing circulation control on a seaplane airfoil. Realizable k-epsilon turbulent model and pressure-based coupled algorithm with second-order discretization were adopted to simulate the compressible flow. Both clear and simple flap configuration were simulated with blowing momentum coefficient Cμ = 0, 0.15 and 0.30. The results show that blowing near the airfoil trailing edge could enhance the Coanda effect, delay the flow separation, and increase the lift coefficient dramatically. The blowing circulation control is promising to apply to taking off and landing of an amphibious aircraft or seaplane.

  5. TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE

    NASA Technical Reports Server (NTRS)

    Dougherty, F. C.

    1994-01-01

    The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters

  6. Multi-pass cooling for turbine airfoils

    DOEpatents

    Liang, George

    2011-06-28

    An airfoil for a turbine vane of a gas turbine engine. The airfoil includes an outer wall having pressure and suction sides, and a radially extending cooling cavity located between the pressure and suction sides. A plurality of partitions extend radially through the cooling cavity to define a plurality of interconnected cooling channels located at successive chordal locations through the cooling cavity. The cooling channels define a serpentine flow path extending in the chordal direction. Further, the cooling channels include a plurality of interconnected chambers and the chambers define a serpentine path extending in the radial direction within the serpentine path extending in the chordal direction.

  7. Nonlinear transonic Wall-Interference Assessment/Correction (WIAC) procedures and application to cast-10 airfoil results from the NASA 0.3-m TCT 8- by 24-inch Slotted Wall Test Section (SWTS)

    NASA Technical Reports Server (NTRS)

    Gumbert, Clyde R.; Green, Lawrence L.; Newman, Perry A.

    1989-01-01

    From the time that wind tunnel wall interference was recognized to be significant, researchers have been developing methods to alleviate or account for it. Despite the best effort so far, it appears that no method is available which completely eliminates the effects due to the wind tunnel walls. This report discusses procedures developed for slotted wall and adaptive wall test sections of the Langley 0.3-m Transonic Cryogenic Tunnel (TCT) to assess and correct for the residual interference by methods consistent with the transonic nature of the tests.

  8. Advanced technology airfoil research, volume 1, part 2

    NASA Technical Reports Server (NTRS)

    1978-01-01

    This compilation contains papers presented at the NASA Conference on Advanced Technology Airfoil Research held at Langley Research Center on March 7-9, 1978, which have unlimited distribution. This conference provided a comprehensive review of all NASA airfoil research, conducted in-house and under grant and contract. A broad spectrum of airfoil research outside of NASA was also reviewed. The major thrust of the technical sessions were in three areas: development of computational aerodynamic codes for airfoil analysis and design, development of experimental facilities and test techniques, and all types of airfoil applications.

  9. An experimental study of airfoil instability tonal noise with trailing edge serrations

    NASA Astrophysics Data System (ADS)

    Chong, Tze Pei; Joseph, Phillip F.

    2013-11-01

    This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien-Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack. Larger Δf, which is defined as (fn+1-fn). In other words, a larger margin of velocity increase is required in order to "shift" the fn and fn+1 across fs

  10. Experimental Study of the Effects of Finite Surface Disturbances and Angle of Attack on the Laminar Boundary Layer of an NACA 64A010 Airfoil with Area Suction

    NASA Technical Reports Server (NTRS)

    Schwartzberg, Milton A; Braslow, Albert L

    1952-01-01

    A Langley low-turbulence wind-tunnel investigation of a porous NACA 64A010 airfoil section has been made to determine the effectiveness of area suction in maintaining full-chord laminar flow behind finite disturbances and at angles of attacks other than 0 degrees. Aero suction resulted in only a small increase in the size of a finite disturbance required to cause premature boundary-layer transition as compared with that for the airfoil without suction. Combined wake and suction drags lower than the drag of the plain airfoil were obtained through a range of low lift coefficient by the use of area suction.

  11. On the acoustic radiation of a pitching airfoil

    NASA Astrophysics Data System (ADS)

    Manela, A.

    2013-07-01

    We examine the acoustic far field of a thin elastic airfoil, immersed in low-Mach non-uniform stream flow, and actuated by small-amplitude sinusoidal pitching motion. The near-field fluid-structure interaction problem is analyzed using potential thin-airfoil theory, combined with a discrete vortex model to describe the evolution of airfoil trailing edge wake. The leading order dipole-sound signature of the system is investigated using Powell-Howe acoustic analogy. Compared with a pitching rigid airfoil, the results demonstrate a two-fold effect of structure elasticity on airfoil acoustic field: at actuation frequencies close to the system least stable eigenfrequency, elasticity amplifies airfoil motion amplitude and associated sound levels; however, at frequencies distant from this eigenfrequency, structure elasticity acts to absorb system kinetic energy and reduce acoustic radiation. In the latter case, and with increasing pitching frequency ωp, a rigid-airfoil setup becomes significantly noisier than an elastic airfoil, owing to an ω _p^{5/2} increase of its direct motion noise component. Unlike rigid airfoil signature, it is shown that wake sound contribution to elastic airfoil radiation is significant for all ωp. Remarkably, this contribution contains, in addition to the fundamental pitching frequency, its odd multiple harmonics, which result from nonlinear interactions between the airfoil and the wake. The results suggest that structure elasticity may serve as a viable means for design of flapping flight noise control methodologies.

  12. Lifetime prediction modeling of airfoils for advanced power generation

    NASA Astrophysics Data System (ADS)

    Karaivanov, Ventzislav Gueorguiev

    The use of gases produced from coal as a turbine fuel offers an attractive means for efficiently generating electric power from our Nation's most abundant fossil fuel resource. The oxy-fuel and hydrogen-fired turbine concepts promise increased efficiency and low emissions on the expense of increased turbine inlet temperature (TIT) and different working fluid. Developing the turbine technology and materials is critical to the creation of these near-zero emission power generation technologies. A computational methodology, based on three-dimensional finite element analysis (FEA) and damage mechanics is presented for predicting the evolution of creep and fatigue in airfoils. We took a first look at airfoil thermal distributions in these advanced turbine systems based on CFD analysis. The damage mechanics-based creep and fatigue models were implemented as user modified routine in commercial package ANSYS. This routine was used to visualize the creep and fatigue damage evolution over airfoils for hydrogen-fired and oxy-fuel turbines concepts, and regions most susceptible to failure were indentified. Model allows for interaction between creep and fatigue damage thus damage due to fatigue and creep processes acting separately in one cycle will affect both the fatigue and creep damage rates in the next cycle. Simulation results were presented for various thermal conductivity of the top coat. Surface maps were created on the airfoil showing the development of the TGO scale and the Al depletion of the bond coat. In conjunction with model development, laboratory-scale experimental validation was executed to evaluate the influence of operational compressive stress levels on the performance of the TBC system. TBC coated single crystal coupons were exposed isothermally in air at 900, 1000, 1100oC with and without compressive load. Exposed samples were cross-sectioned and evaluated with scanning electron microscope (SEM). Performance data was collected based on image analysis

  13. Trailing edge flow conditions as a factor in airfoil design

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.; Maughmer, M. D.

    1984-01-01

    Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the role played by the non-dimensional trailing edge velocity in design is considered and the relationship between the specified value and the resulting airfoil geometry is explored. In addition, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution of the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler is presented and the modified scheme employed to generate a number of airfoils for consideration. The detailed viscous analysis of airfoils having finite trailing edge pressure gradients demonstrates a reduction in the strong inviscid-viscid interactions generally present near the trailing edge of an airfoil.

  14. Aerodynamic Simulation of Ice Accretion on Airfoils

    NASA Technical Reports Server (NTRS)

    Broeren, Andy P.; Addy, Harold E., Jr.; Bragg, Michael B.; Busch, Greg T.; Montreuil, Emmanuel

    2011-01-01

    This report describes recent improvements in aerodynamic scaling and simulation of ice accretion on airfoils. Ice accretions were classified into four types on the basis of aerodynamic effects: roughness, horn, streamwise, and spanwise ridge. The NASA Icing Research Tunnel (IRT) was used to generate ice accretions within these four types using both subscale and full-scale models. Large-scale, pressurized windtunnel testing was performed using a 72-in.- (1.83-m-) chord, NACA 23012 airfoil model with high-fidelity, three-dimensional castings of the IRT ice accretions. Performance data were recorded over Reynolds numbers from 4.5 x 10(exp 6) to 15.9 x 10(exp 6) and Mach numbers from 0.10 to 0.28. Lower fidelity ice-accretion simulation methods were developed and tested on an 18-in.- (0.46-m-) chord NACA 23012 airfoil model in a small-scale wind tunnel at a lower Reynolds number. The aerodynamic accuracy of the lower fidelity, subscale ice simulations was validated against the full-scale results for a factor of 4 reduction in model scale and a factor of 8 reduction in Reynolds number. This research has defined the level of geometric fidelity required for artificial ice shapes to yield aerodynamic performance results to within a known level of uncertainty and has culminated in a proposed methodology for subscale iced-airfoil aerodynamic simulation.

  15. Exact solutions in oscillating airfoil theory

    NASA Technical Reports Server (NTRS)

    Williams, M. H.

    1977-01-01

    A result obtained by Williams (1977) for two-dimensional airfoils oscillating in an arbitrary subsonic parallel flowfield is reformulated to show that the pressure distribution induced by any deformation can be construed from the particular solutions for heaving and pitching motions. Specific formulas are presented for an oscillating control surface with a sealed gap.

  16. Near-wall serpentine cooled turbine airfoil

    DOEpatents

    Lee, Ching-Pang

    2014-10-28

    A serpentine coolant flow path is formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil, the cavity partitioned by one or more transverse partitions into a plurality of continuous serpentine cooling flow streams each having a respective coolant inlet.

  17. Downwash and Wake Behind Plain and Flapped Airfoils

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe; Katzoff, S; Bullivant, W Kenneth

    1939-01-01

    Extensive experimental measurements have been made of the downwash angles and the wake characteristics behind airfoils with and without flaps and the data have been analyzed and correlated with the theory. A detailed study was made of the errors involved in applying lifting-line theory, such as the effects of a finite wing chord, the rolling-up of the trailing vortex sheet, and the wake. The downwash angles, as computed from the theoretical span load distribution by means of the Biot-Savart equation, were found to be in satisfactory agreement with the experimental results. The rolling-up of the trailing vortex sheet may be neglected, but the vertical displacement of the vortex sheet requires consideration. By the use of a theoretical treatment indicated by Prandtl, it has been possible to generalize the available experimental results so the predictions can be made of the important wake parameters in terms of the distance behind the airfoil trailing edge and the profile-drag coefficient. The method of application of the theory to design and the satisfactory agreement between predicted and experimental results when applied to an airplane are demonstrated.

  18. CFD simulation of turbulent airflow around wind turbine airfoils

    NASA Astrophysics Data System (ADS)

    Halbrooks, David N.

    The airflow around wind turbines has proved to be a difficult problem to approach by means of today's Computational Fluid Dynamics (CFD) codes. One reason for this difficulty lies within the stall characteristics of turbine airfoils. For the purposes of this research, the popular commercial CFD code, FLUENT was employed to facilitate the understanding of airflow around wind turbines through the study of various turbulence models. Parallel processing was employed to enhance computational performance as well as lower simulation times. The system used for simulation is the National Renewable Energy Laboratory (NREL) Phase VI Wind Turbine. The coefficients of pressure for the airfoil were extracted from the simulated data and compared against data obtained during the NREL Phase VI Wind Turbine data campaign. Since power is a driving factor of the design of wind turbine blades, the aspect of power was also examined and compared. After the completion of the baseline study, a parametric study was carried out to examine the effects of rotor speed downstream of the turbine blades.

  19. Epidemiological and Clinical Characteristics of Fall Injuries in East Azerbaijan, Iran; A Cross-Sectional Study

    PubMed Central

    Ghaffari-Fam, Saber; Sarbazi, Ehsan; Daemi, Amin; Sarbazi, Mohamadreza; Riyazi, Lachin; Sadeghi-Bazargani, Homayoun; Allahyari, Ali

    2015-01-01

    Objective: To describe the epidemiological and clinical characteristics of fall injuries in East Azerbaijan, Iran. Methods: This cross-sectional study was based on Hospital Information System (HIS) data for patients referred to the Imam Reza Hospital between 2008 and 2013. We recorded the demographic characteristics and epidemiological patterns of patients who were admitted to our center due to fall injuries. To standardize the reports the International Classification of Diseases (ICD), the International Classification of Diseases 9 Clinical Modification (ICD-9-CM) was used. Equally, the hospitalization period and number, admission ward, and the final status of victims after discharge from the hospital were extracted from the HIS. Results: Overall we included a total number of 3397 patients with mean age of 39.2±22.7 years. There were 2501 (73.6%) men among the patients. Long bone fracture (48.1%) and intracranial injury (24.2%) were the most frequent injuries among fall injury victims. Operations on spinal cord and spinal canal structures (12.0%), Operations on nose (11.6%) were the most common operations being performed in these patients. The survival was significantly lower in patients with age more than 60 years when compared to other age groups (p=0.001). The survival rate was significantly lower in age group of >60 years, compared to other age groups (p=0.001). Conclusion: Given the high rate of fall injuries and death among the elderly that increases with age, appropriate measures must be taken to control and prevent injuries while prioritizing the elderly. PMID:27162912

  20. Increased cesarean section rate in Central Saudi Arabia: a change in practice or different maternal characteristics

    PubMed Central

    Al-Kadri, Hanan M; Al-Anazi, Sultana A; Tamim, Hani M

    2015-01-01

    Background Cesarean section (CS) rate has shown an alarming increase. We aimed in this work to identify factors contributing to the increasing rate of CS in central Saudi Arabia. Methods A retrospective cohort study was conducted at King Abdulaziz Medical City. Two groups of women were included (G1 and G2). G1 had delivered by CS during the year 2002 (CS rate 12%), and G2 had delivered by CS during the year 2009 (CS rate 20%). We compared the included women’s characteristics, neonates, CS indications, and complications. Data were analyzed using SPSS version 15 program. Odds ratios and confidence intervals were calculated to report precision of categorical data results. A P-value of ≤0.05 was considered significant. Results A total of 198 women were included in G1 and 200 in G2. Both groups had comparable maternal and fetal characteristics; however, absence of antenatal care has resulted in 70% increase in CS deliveries for G2, P=0.008, OR =0.30, CI 0.12–0.76. Previous vaginal surgeries have contributed to tenfold increase in CS deliveries for G2, P=0.006, OR =10.37, CI 1.32–81.78. G2 had eight times increased CS deliveries than G1 due to intrauterine growth restriction, P=0.02, OR =8.21, CI 1.02–66.25, and 80% increased risk of CS was based on maternal demand, P=0.02, OR =0.20, CI 0.02–1.71. Decision taken by less-experienced staff was associated with 2.5-fold increase in CS deliveries for G2, P=0.002, OR =2.62, CI 1.39–4.93. There was a significant increase in CS deliveries under regional analgesia and shorter duration of hospital stay for G2, P=0.0001 and P=0.001, respectively. G2 women had 2.75-fold increase in neonatal intensive care unit admission, P=0.03, OR =2.75, CI 1.06–7.15. Conclusion CS delivery rate significantly increased within the studied population. The increased rate of CS may be related to a change in physician’s practice rather than a change in maternal characteristics, and it appears to be reducible. PMID:26203285

  1. [Characteristics of "pre-ALIP" in bone marrow sections of patients with acute myeloid leukemia].

    PubMed

    Yu, Ye-Hua; Yuan, Ying-Hua; Li, Ling; Cao, Gui-Tao; Li, Jing; Tao, Ying; Jiang, Hua; Hou, Jian; Shi, Jun

    2011-10-01

    To detect the characteristics of "pre-ALIP" and to investigate their relevance with the development of acute myeloid leukemia (AML) by computer image procession technology, bone marrow (BM) was collected by aspiration/trephine biopsy from AML patients during the complete remission (CR). BM sections were stained by HGF (haematoxylin-Giemsa-acid fuchsin) and photographed by optical microscope imaging system. 4 kinds of computer image segmentation technologies were compared to select the best one for detecting the localization and quantitation of the precursor cells. Planimetry was combined with morphology to segment bone trabeculae. The number of single and double-cluster precursor cells and their distance from bone trabeculae was detected with Euclidean distance change method in BM images of AML patients, and compared with the normal controls. Moreover, the morphological characteristics of "pre-ALIP" were investigated, and the correlation with the development of AML was analyzed. The results showed that the computer image segmentation method based on morphology could identify the precursor cells and bone trabeculae more exactly in BM image, as compared with the methods of 8-Sobel operater. Canny operator and watershed algorithm. Bone trabeculae could be segmented with combinative methods of morphology and planimetry. The number of single precursor cells (19.27 ± 11.60)/mm(2) and double-cluster precursor cells (1.77 ± 1.76)/mm(2) in CR group were higher than that in normal controls (p < 0.05). The distance of single precursor cells from bone trabeculae in CR group were closer to bone trabeculae than that in controls [(230.12 ± 97.68) µm vs (260.92 ± 99.88 µm)] (p < 0.05), but the distance of double-cluster precursor cells from bone trabeculae in AML patients was (274.56 ± 139.48) µm, which showed no statistically significant different from controls (p > 0.05), while the double-cluster precursor cells showed the tendency of migrating to the intermediate zone

  2. Airfoil Ice-Accretion Aerodynamics Simulation

    NASA Technical Reports Server (NTRS)

    Bragg, Michael B.; Broeren, Andy P.; Addy, Harold E.; Potapczuk, Mark G.; Guffond, Didier; Montreuil, E.

    2007-01-01

    NASA Glenn Research Center, ONERA, and the University of Illinois are conducting a major research program whose goal is to improve our understanding of the aerodynamic scaling of ice accretions on airfoils. The program when it is completed will result in validated scaled simulation methods that produce the essential aerodynamic features of the full-scale iced-airfoil. This research will provide some of the first, high-fidelity, full-scale, iced-airfoil aerodynamic data. An initial study classified ice accretions based on their aerodynamics into four types: roughness, streamwise ice, horn ice, and spanwise-ridge ice. Subscale testing using a NACA 23012 airfoil was performed in the NASA IRT and University of Illinois wind tunnel to better understand the aerodynamics of these ice types and to test various levels of ice simulation fidelity. These studies are briefly reviewed here and have been presented in more detail in other papers. Based on these results, full-scale testing at the ONERA F1 tunnel using cast ice shapes obtained from molds taken in the IRT will provide full-scale iced airfoil data from full-scale ice accretions. Using these data as a baseline, the final step is to validate the simulation methods in scale in the Illinois wind tunnel. Computational ice accretion methods including LEWICE and ONICE have been used to guide the experiments and are briefly described and results shown. When full-scale and simulation aerodynamic results are available, these data will be used to further develop computational tools. Thus the purpose of the paper is to present an overview of the program and key results to date.

  3. Site insolation and wind power characteristics: technical report western region (north section)

    SciTech Connect

    1980-08-01

    This phase of the Site Insolation and Wind Power Characteristics Study was performed to provide statistical information on the expected future availability of solar and wind power at various sites in the Western Region (North Section) of the US Historic data (SOLMET), at 21 National Weather Service stations with hourly solar insolation and collateral meteorological information, were interrogated to provide an estimate of future trends. Solar data are global radiation incident on a horizontal surface, and wind data represent wind power normal to the air flow. Selected insolation and wind power conditions were investigated for their occurrence and persistence, for defined periods of time, on a monthly basis. Global horizontal insolation is related to inclined surfaces at each site. Ratios are provided, monthly, for multiplying global insolation to obtain insolation estimates on south-facing surfaces inclined at different angles with respect to the horizontal. Also, joint probability distribution tables are constructed showing the number of occurrences, out of a finite sample size, of daily average solar and wind power within selected intervals, by month. Information of this nature is intended as an aid to preliminary planning activities for the design and operation of solar and wind energy utilization and conversion systems.

  4. High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, Charles B.; Dress, David A.; Hill, Acquilla S.; Wilcox, Peter A.; Bui, Minh H.

    1986-01-01

    A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

  5. Experimental and Theoretical Studies of Area Suction for the Control of the Laminar Boundary Layer on an NACA 64a010 Airfoil

    NASA Technical Reports Server (NTRS)

    Braslow, Albert L; Burrows, Dale L; Tetervin, Neal; Visconti, Fioravante

    1951-01-01

    A low-turbulence wind-tunnel investigation was made of an NACA 64a010 airfoil having a porous surface to determine the reduction in section total-drag coefficient that might be obtained at large Reynolds numbers by the use of suction to produce continuous inflow through the surface of the airfoil (area suction). In addition to the experimental investigation, a related theoretical analysis was made to provide a basis of comparison for the test results.

  6. Experimental Measurement and CFD Model Development of Thick Wind Turbine Airfoils with Leading Edge Erosion

    NASA Astrophysics Data System (ADS)

    Maniaci, David C.; White, Edward B.; Wilcox, Benjamin; Langel, Christopher M.; van Dam, C. P.; Paquette, Joshua A.

    2016-09-01

    Leading edge erosion and roughness accumulation is an issue observed with great variability by wind plant operators, but with little understanding of the effect on wind turbine performance. In wind tunnels, airfoil models are typically tested with standard grit roughness and trip tape to simulate the effects of roughness and erosion observed in field operation, but there is a lack of established relation between field measurements and wind tunnel test conditions. A research collaboration between lab, academic, and industry partners has sought to establish a method to estimate the effect of erosion in wind turbine blades that correlates to roughness and erosion measured in the field. Measurements of roughness and erosion were taken off of operational utility wind turbine blades using a profilometer. The field measurements were statistically reproduced in the wind tunnel on representative tip and midspan airfoils. Simultaneously, a computational model was developed and calibrated to capture the effect of roughness and erosion on airfoil transition and performance characteristics. The results indicate that the effects of field roughness fall between clean airfoil performance and the effects of transition tape. Severe leading edge erosion can cause detrimental performance effects beyond standard roughness. The results also indicate that a heavily eroded wind turbine blade can reduce annual energy production by over 5% for a utility scale wind turbine.

  7. Three-dimensional effects on airfoil measurements at high Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Kiefer, Janik; Miller, Mark; Hultmark, Marcus; Hansen, Martin

    2016-11-01

    Blade Element Momentum codes (BEM) are widely used in the wind turbine industry to determine a turbine's operational range and its limits. Empirical two-dimensional airfoil data serve as the primary and fundamental input to the BEM code. Consequently, the results of BEM simulations are strongly dependent on the accuracy of these data. In this presentation, an experimental study is described in which airfoils of different aspect ratios were tested at identical Reynolds numbers. A high-pressure wind tunnel facility is used to achieve large Reynolds numbers of Rec = 3 ×106 , even with small chord lengths. This methodology enables testing of very high aspect ratio airfoils to characterize 3-D effects on the lift and drag data. The tests were performed over a large range of angles of attack, which is especially important for wind turbines. The effect of varying aspect ratio on the aerodynamic characteristics of the airfoil is discussed with emphasis on the outcome of a BEM simulation. The project was partially funded by NSF CBET-1435254 (program manager Dr. Gregory Rorrer).

  8. Wall-Modeled Large-Eddy Simulation of Turbulent Flow Past an Airfoil

    NASA Astrophysics Data System (ADS)

    Gao, Wei; Zhang, Wei; Samtaney, Ravi

    2015-11-01

    We present wall-modeled large-eddy simulations (WMLES) for turbulent flows incompressible past an airfoil. The virtual wall model, originally developed by Chung & Pullin (J. of Fluid Mech., 2009), is extended to generalized curvilinear coordinates and implemented using a body-fitted structured C-grid for airfoils. This model dynamically couples the outer resolved region with the wall region, and imposes a slip velocity boundary condition for the filtered velocity field on the ``virtual'' wall. The virtual wall model is combined with the stretched spiral vortex sub-grid scale model in a self-consistent framework, which is tested in WMLES of flow past a NACA0012 airfoil at different Reynolds number (Re) and angle of attack. The numerical results show that the wall model is able to accurately predict mean flow characteristics, including the formation of the separation bubble. Some high-order turbulence quantities are also compared with the direct numerical simulation results (Re =104) of flow past the same airfoil. We will present verification test cases to quantify the effectiveness of the wall model in both attached and separated flow regimes. Supported by the KAUST Office of Competitive Research Funds under Award No. URF/1/1394-01. The IBM Blue Gene/P Shaheen at KAUST was utilized for the simulations.

  9. The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

    NASA Technical Reports Server (NTRS)

    Schwind, R. G.; Allen, H. J.

    1973-01-01

    High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

  10. A sampling scheme to assess persistence and transport characteristics of xenobiotics within an urban river section

    NASA Astrophysics Data System (ADS)

    Schwientek, Marc; Guillet, Gaelle; Kuch, Bertram; Rügner, Hermann; Grathwohl, Peter

    2014-05-01

    Xenobiotic contaminants such as pharmaceuticals or personal care products typically are continuously introduced into the receiving water bodies via wastewater treatment plant (WWTP) outfalls and, episodically, via combined sewer overflows in the case of precipitation events. Little is known about how these chemicals behave in the environment and how they affect ecosystems and human health. Examples of traditional persistent organic pollutants reveal, that they may still be present in the environment even decades after they have been released. In this study a sampling strategy was developed which gives valuable insights into the environmental behaviour of xenobiotic chemicals. The method is based on the Lagrangian sampling scheme by which a parcel of water is sampled repeatedly as it moves downstream while chemical, physical, and hydrologic processes altering the characteristics of the water mass can be investigated. The Steinlach is a tributary of the River Neckar in Southwest Germany with a catchment area of 140 km². It receives the effluents of a WWTP with 99,000 inhabitant equivalents 4 km upstream of its mouth. The varying flow rate of effluents induces temporal patterns of electrical conductivity in the river water which enable to track parcels of water along the subsequent urban river section. These parcels of water were sampled a) close to the outlet of the WWTP and b) 4 km downstream at the confluence with the Neckar. Sampling was repeated at a 15 min interval over a complete diurnal cycle and 2 h composite samples were prepared. A model-based analysis demonstrated, on the one hand, that substances behaved reactively to a varying extend along the studied river section. On the other hand, it revealed that the observed degradation rates are likely dependent on the time of day. Some chemicals were degraded mainly during daytime (e.g. the disinfectant Triclosan or the phosphorous flame retardant TDCP), others as well during nighttime (e.g. the musk fragrance

  11. Estimating HIV-1 fitness characteristics from cross-sectional genotype data.

    PubMed

    Gopalakrishnan, Sathej; Montazeri, Hesam; Menz, Stephan; Beerenwinkel, Niko; Huisinga, Wilhelm

    2014-11-01

    Despite the success of highly active antiretroviral therapy (HAART) in the management of human immunodeficiency virus (HIV)-1 infection, virological failure due to drug resistance development remains a major challenge. Resistant mutants display reduced drug susceptibilities, but in the absence of drug, they generally have a lower fitness than the wild type, owing to a mutation-incurred cost. The interaction between these fitness costs and drug resistance dictates the appearance of mutants and influences viral suppression and therapeutic success. Assessing in vivo viral fitness is a challenging task and yet one that has significant clinical relevance. Here, we present a new computational modelling approach for estimating viral fitness that relies on common sparse cross-sectional clinical data by combining statistical approaches to learn drug-specific mutational pathways and resistance factors with viral dynamics models to represent the host-virus interaction and actions of drug mechanistically. We estimate in vivo fitness characteristics of mutant genotypes for two antiretroviral drugs, the reverse transcriptase inhibitor zidovudine (ZDV) and the protease inhibitor indinavir (IDV). Well-known features of HIV-1 fitness landscapes are recovered, both in the absence and presence of drugs. We quantify the complex interplay between fitness costs and resistance by computing selective advantages for different mutants. Our approach extends naturally to multiple drugs and we illustrate this by simulating a dual therapy with ZDV and IDV to assess therapy failure. The combined statistical and dynamical modelling approach may help in dissecting the effects of fitness costs and resistance with the ultimate aim of assisting the choice of salvage therapies after treatment failure.

  12. Characteristics of street children in Cameroon: A cross-sectional study

    PubMed Central

    Cumber, Samuel N.

    2016-01-01

    Introduction The issue of street children is one of the global social problems rising in low- and middle-income countries. These children are vulnerable, but because of a lack of sufficient information, it is very difficult for stakeholders to address their plight in Cameroon. Aim To examine the situation and characteristics of street children in three Cameroonian cities. Objectives To describe the demographic, socio-economic and behavioural profiles of street children. To identify challenges of street children and to compare the results from the three cities on account of their different settings, cultural history and challenges. Materials and methods The study was an analytical cross-sectional survey conducted through researcher-administered questionnaires to 399 street children (homeless for at least a month), in three Cameroonian cities from 1 January 2015 to 30 March 2015. Results The majority of the participants were boys, more than 70% were homeless for less than 12 months and poverty was found to be the most common reason for being on the street. Most of the participants earned less than 500CFA francs (USD 0.85), with many of them resorting to begging, drug abuse, sex work and other risky behaviours. Only two of the respondents (0.5%) regarded the public attitude towards them as supportive. Conclusion As children roam the streets in search of shelter, food and other basic needs, their future hangs in the balance. Understanding the plight of street children highlights the need for immediate design and implementation of intervention strategies to prevent children from living in the streets and assist those who have become street children. PMID:28155316

  13. Turbine airfoil with laterally extending snubber having internal cooling system

    SciTech Connect

    Scribner, Carmen Andrew; Messmann, Stephen John; Marsh, Jan H.

    2016-09-06

    A turbine airfoil usable in a turbine engine and having at least one snubber with a snubber cooling system positioned therein and in communication with an airfoil cooling system is disclosed. The snubber may extend from the outer housing of the airfoil toward an adjacent turbine airfoil positioned within a row of airfoils. The snubber cooling system may include an inner cooling channel separated from an outer cooling channel by an inner wall. The inner wall may include a plurality of impingement cooling orifices that direct impingement fluid against an outer wall defining the outer cooling channel. In one embodiment, the cooling fluids may be exhausted from the snubber, and in another embodiment, the cooling fluids may be returned to the airfoil cooling system. Flow guides may be positioned in the outer cooling channel, which may reduce cross-flow by the impingement orifices, thereby increasing effectiveness.

  14. New airfoils for small horizontal axis wind turbines

    SciTech Connect

    Giguere, P.; Selig, M.S.

    1997-12-31

    In a continuing effort to enhance the performance of small energy systems, one root airfoil and three primary airfoils were specifically designed for small horizontal axis wind turbines. These airfoils are intended primarily for 1-10 kW variable-speed wind turbines for both conventional (tapered/twisted) or pultruded blades. The four airfoils were wind-tunnel tested at Reynolds numbers between 100,000 and 500,000. Tests with simulated leading-edge roughness were also conducted. The results indicate that small variable-speed wind turbines should benefit from the use of the new airfoils which provide enhanced lift-to-drag ratio performance as compared with previously existing airfoils.

  15. Quiet airfoils for small and large wind turbines

    DOEpatents

    Tangler, James L [Boulder, CO; Somers, Dan L [Port Matilda, PA

    2012-06-12

    Thick airfoil families with desirable aerodynamic performance with minimal airfoil induced noise. The airfoil families are suitable for a variety of wind turbine designs and are particularly well-suited for use with horizontal axis wind turbines (HAWTs) with constant or variable speed using pitch and/or stall control. In exemplary embodiments, a first family of three thick airfoils is provided for use with small wind turbines and second family of three thick airfoils is provided for use with very large machines, e.g., an airfoil defined for each of three blade radial stations or blade portions defined along the length of a blade. Each of the families is designed to provide a high maximum lift coefficient or high lift, to exhibit docile stalls, to be relatively insensitive to roughness, and to achieve a low profile drag.

  16. Investigation of low-speed turbulent separated flow around airfoils

    NASA Technical Reports Server (NTRS)

    Wadcock, Alan J.

    1987-01-01

    Described is a low-speed wind tunnel experiment to measure the flowfield around a two-dimensional airfoil operating close to maximum lift. Boundary layer separation occurs on the upper surface at x/c=0.85. A three-component laser velocimeter, coupled with a computer-controlled data acquisition system, was used to obtain three orthogonal mean velocity components and three components of the Reynolds stress tensor in both the boundary layer and wake of the airfoil. Pressure distributions on the airfoil, skin friction distribution on the upper surface of the airfoil, and integral properties of the airfoil boudary layer are also documented. In addition to these near-field flow properties, static pressure distributions, both upstream and downstream from the airfoil and on the walls of the wind tunnel, are also presented.

  17. Three-dimensional unsteady viscous flow analysis over airfoil sections

    NASA Technical Reports Server (NTRS)

    Weinberg, B. C.; Shamroth, S. J.

    1984-01-01

    A three-dimensional solution procedure for the approximate form of the Navier-Stokes equation was exercised in the two- and three-dimensional modes to compute the unsteady turbulent boundary layer on a flat plate corresponding to the data of Karlsson. The procedure is based on the use of a consistently split Linearized Block Implicit technique in conjunction with a QR operator scheme. New time-dependent upstream boundary conditions were developed that yielded realistic solutions for the interior in the vicinity of the upstream boundary. Comparisons of the computation employing these boundary conditions with the data indicate that both qualitative and quantitative agreement was obtained for the mean velocity and the in phase and out of phase components of the first harmonic of the velocity. In addition, the calculation gave results for the skin friction phase angle that had expected physical behavior for large distances downstream of the inflow boundary. For the three-dimensional case, the two-dimensional data of Karlsson was considered, but in a coordinate system skewed at 45 deg to the free stream direction. The results of the calculations were in excellent agreement with the data and the two-dimensional computations.

  18. Three-dimensional unsteady viscous flow analysis over airfoil sections

    NASA Astrophysics Data System (ADS)

    Weinberg, B. C.; Shamroth, S. J.

    1984-06-01

    A three-dimensional solution procedure for the approximate form of the Navier-Stokes equation was exercised in the two- and three-dimensional modes to compute the unsteady turbulent boundary layer on a flat plate corresponding to the data of Karlsson. The procedure is based on the use of a consistently split Linearized Block Implicit technique in conjunction with a QR operator scheme. New time-dependent upstream boundary conditions were developed that yielded realistic solutions for the interior in the vicinity of the upstream boundary. Comparisons of the computation employing these boundary conditions with the data indicate that both qualitative and quantitative agreement was obtained for the mean velocity and the in phase and out of phase components of the first harmonic of the velocity. In addition, the calculation gave results for the skin friction phase angle that had expected physical behavior for large distances downstream of the inflow boundary. For the three-dimensional case, the two-dimensional data of Karlsson was considered, but in a coordinate system skewed at 45 deg to the free stream direction. The results of the calculations were in excellent agreement with the data and the two-dimensional computations.

  19. S825 and S826 Airfoils: 1994--1995

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A family of airfoils, the S825 and S826, for 20- to 40-meter, variable-speed and variable-pitch (toward feather), horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moments and the airfoil thicknesses have been satisfied. The airfoils should exhibit docile stalls.

  20. High Reynolds number tests of the CAST-10-2/DOA 2 transonic airfoil at ambient and cryogenic temper ature conditions

    NASA Technical Reports Server (NTRS)

    Stanewsky, E.; Demurie, F.; Ray, Edward J.; Johnson, C. B.

    1989-01-01

    The transonic airfoil CAST 10-2/DOA 2 was investigated in several major transonic wind tunnels at Reynolds numbers ranging from Re=1.3 x 10(exp 6) to 45 x 10(exp 6) at ambient and cryogenic temperature conditions. The main objective was to study the degree and extent of the effects of Reynolds number on both the airfoil aerodynamic characteristics and the interference effects of various model-wind-tunnel systems. The initial analysis of the CAST 10-2 airfoil results revealed appreciable real Reynolds number effects on this airfoil and showed that wall interference can be significantly affected by changes in Reynolds number thus appearing as true Reynolds number effects.

  1. An approach to constrained aerodynamic design with application to airfoils

    NASA Technical Reports Server (NTRS)

    Campbell, Richard L.

    1992-01-01

    An approach was developed for incorporating flow and geometric constraints into the Direct Iterative Surface Curvature (DISC) design method. In this approach, an initial target pressure distribution is developed using a set of control points. The chordwise locations and pressure levels of these points are initially estimated either from empirical relationships and observed characteristics of pressure distributions for a given class of airfoils or by fitting the points to an existing pressure distribution. These values are then automatically adjusted during the design process to satisfy the flow and geometric constraints. The flow constraints currently available are lift, wave drag, pitching moment, pressure gradient, and local pressure levels. The geometric constraint options include maximum thickness, local thickness, leading-edge radius, and a 'glove' constraint involving inner and outer bounding surfaces. This design method was also extended to include the successive constraint release (SCR) approach to constrained minimization.

  2. Dynamic stall experiments on the NACA 0012 airfoil

    NASA Technical Reports Server (NTRS)

    Mcalister, K. W.; Carr, L. W.; Mccroskey, W. J.

    1978-01-01

    The flow over a NACA 0012 airfoil undergoing large oscillations in pitch was experimentally studied at a Reynolds number of and over a range of frequencies and amplitudes. Hot-wire probes and surface-pressure transducers were used to clarify the role of the laminar separation bubble, to delineate the growth and shedding of the stall vortex, and to quantify the resultant aerodynamic loads. In addition to the pressure distributions and normal force and pitching moment data that have often been obtained in previous investigations, estimates of the unsteady drag force during dynamic stall have been derived from the surface pressure measurements. Special characteristics of the pressure response, which are symptomatic of the occurrence and relative severity of moment stall, have also been examined.

  3. Sensitivity analysis applied to stalled airfoil wake and steady control

    NASA Astrophysics Data System (ADS)

    Patino, Gustavo; Gioria, Rafael; Meneghini, Julio

    2014-11-01

    The sensitivity of an eigenvalue to base flow modifications induced by an external force is applied to the global unstable modes associated to the onset of vortex shedding in the wake of a stalled airfoil. In this work, the flow regime is close to the first instability of the system and its associated eigenvalue/eigenmode is determined. The sensitivity analysis to a general punctual external force allows establishing the regions where control devices must be in order to stabilize the global modes. Different types of steady control devices, passive and active, are used in the regions predicted by the sensitivity analysis to check the vortex shedding suppression, i.e. the primary instability bifurcation is delayed. The new eigenvalue, modified by the action of the device, is also calculated. Finally the spectral finite element method is employed to determine flow characteristics before and after of the bifurcation in order to cross check the results.

  4. Wind Tunnel Tests of Ailerons at Various Speeds I : Ailerons of 0.20 Airfoil Chord and True Contour with 0.35 Aileron-chord Extreme Blunt Nose Balance on the NACA 66,2-216 Airfoil

    NASA Technical Reports Server (NTRS)

    Letko, W; Denaci, H. G.; Freed, C

    1943-01-01

    Hinge-moment, lift, and pressure-distribution measurements were made in the two-dimensional test section of the NACA stability tunnel on a blunt-nose balance-type aileron on an NACA 66,2-216 airfoil at speeds up to 360 miles per hour corresponding to a Mach number of 0.475. The tests were made primarily to determine the effect of speed on the action of this type of aileron. The balance-nose radii of the aileron were varied from 0 to 0.02 of the airfoil chord and the gap width was varied from 0.0005 to 0.0107 of the airfoil chord. Tests were also made with the gap sealed.

  5. Effect of initial acceleration on the development of the flow field of an airfoil pitching at constant rate

    NASA Technical Reports Server (NTRS)

    Koochesfahani, M. M.; Smiljanovski, V.; Brown, T. A.

    1992-01-01

    We present results from a series of experiments where an airfoil is pitched at constant rate from 0 to 60 degrees angle of attack. It is well documented that the dynamic stall behavior of such an airfoil strongly depends on the nondimensional pitch rate K = dot-alpha C/(2U(sub infinity)), where C is the chord, dot-alpha the constant pitch rate, and U(sub infinity) the free stream speed. In reality, the actual motion of the airfoil deviates from the ideal ramp due to the finite acceleration and deceleration periods imposed by the damping of drive system and response characteristics of the airfoil. It is possible that the pitch rate alone may not suffice in describing the flow and that the details of the motion trajectory before achieving a desired constant pitch rate may also affect the processes involved in the dynamic stall phenomenon. The effects of acceleration and deceleration periods are investigated by systematically varing the acceleration magnitude and its duration through the initial acceleration phase to constant pitch rate. The magnitude and duration of deceleration needed to bring the airfoil motion to rest is similarly controlled.

  6. Turbine airfoil with ambient cooling system

    DOEpatents

    Campbell, Jr, Christian X.; Marra, John J.; Marsh, Jan H.

    2016-06-07

    A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

  7. Analysis of airfoil transitional separation bubbles

    NASA Technical Reports Server (NTRS)

    Davis, R. L.; Carter, J. E.

    1984-01-01

    A previously developed local inviscid-viscous interaction technique for the analysis of airfoil transitional separation bubbles, ALESEP (Airfoil Leading Edge Separation) has been modified to utilize a more accurate windward finite difference procedure in the reversed flow region, and a natural transition/turbulence model has been incorporated for the prediction of transition within the separation bubble. Numerous calculations and experimental comparisons are presented to demonstrate the effects of the windward differencing scheme and the natural transition/turbulence model. Grid sensitivity and convergence capabilities of this inviscid-viscous interaction technique are briefly addressed. A major contribution of this report is that with the use of windward differencing, a second, counter-rotating eddy has been found to exist in the wall layer of the primary separation bubble.

  8. Turbine engine airfoil and platform assembly

    DOEpatents

    Campbell, Christian X [Oviedo, FL; James, Allister W [Chuluota, FL; Morrison, Jay A [Oviedo, FL

    2012-07-31

    A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.

  9. Turbomachinery Airfoil Design Optimization Using Differential Evolution

    NASA Technical Reports Server (NTRS)

    Madavan, Nateri K.; Biegel, Bryan (Technical Monitor)

    2002-01-01

    An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine and compared to earlier methods. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.

  10. Turbomachinery Airfoil Design Optimization Using Differential Evolution

    NASA Technical Reports Server (NTRS)

    Madavan, Nateri K.; Biegel, Bryan A. (Technical Monitor)

    2002-01-01

    An aerodynamic design optimization procedure that is based on a evolutionary algorithm known at Differential Evolution is described. Differential Evolution is a simple, fast, and robust evolutionary strategy that has been proven effective in determining the global optimum for several difficult optimization problems, including highly nonlinear systems with discontinuities and multiple local optima. The method is combined with a Navier-Stokes solver that evaluates the various intermediate designs and provides inputs to the optimization procedure. An efficient constraint handling mechanism is also incorporated. Results are presented for the inverse design of a turbine airfoil from a modern jet engine. The capability of the method to search large design spaces and obtain the optimal airfoils in an automatic fashion is demonstrated. Substantial reductions in the overall computing time requirements are achieved by using the algorithm in conjunction with neural networks.

  11. Improved methods of vibration analysis of pretwisted, airfoil blades

    NASA Technical Reports Server (NTRS)

    Subrahmanyam, K. B.; Kaza, K. R. V.

    1984-01-01

    Vibration analysis of pretwisted blades of asymmetric airfoil cross section is performed by using two mixed variational approaches. Numerical results obtained from these two methods are compared to those obtained from an improved finite difference method and also to those given by the ordinary finite difference method. The relative merits, convergence properties and accuracies of all four methods are studied and discussed. The effects of asymmetry and pretwist on natural frequencies and mode shapes are investigated. The improved finite difference method is shown to be far superior to the conventional finite difference method in several respects. Close lower bound solutions are provided by the improved finite difference method for untwisted blades with a relatively coarse mesh while the mixed methods have not indicated any specific bound.

  12. Damping element for reducing the vibration of an airfoil

    SciTech Connect

    Campbell, Christian X; Marra, John J

    2013-11-12

    An airfoil (10) is provided with a tip (12) having an opening (14) to a center channel (24). A damping element (16) is inserted within the opening of the center channel, to reduce an induced vibration of the airfoil. The mass of the damping element, a spring constant of the damping element within the center channel, and/or a mounting location (58) of the damping element within the center channel may be adjustably varied, to shift a resonance frequency of the airfoil outside a natural operating frequency of the airfoil.

  13. Transonic airfoil analysis and design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1975-01-01

    An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

  14. Numerical investigation of acoustic radiation from vortex-airfoil interaction

    NASA Astrophysics Data System (ADS)

    Legault, Anne; Ji, Minsuk; Wang, Meng

    2012-11-01

    Numerical simulations of vortices interacting with a NACA 0012 airfoil and a flat-plate airfoil at zero angle of attack are carried out to assess the applicability and accuracy of classical theories. Unsteady lift and sound are computed and compared with the predictions by theories of Sears and Amiet, which assume a thin-plate airfoil in an inviscid flow. A Navier-Stokes solver is used in the simulations, and therefore viscous effects are taken into consideration. For the thin-plate airfoil, the effect of viscosity is negligible. For a NACA 0012 airfoil, the viscous contribution to the unsteady lift and sound mainly comes from coherent vortex shedding in the wake of the airfoil and the interaction of the incoming vortices with the airfoil wake, which become stronger at higher Reynolds numbers for a 2-D laminar flow. When the flow is turbulent at chord Reynolds number of 4 . 8 ×105 , however, the viscous contribution becomes negligible as coherent vortex shedding is not present. Sound radiation from vortex-airfoil interaction at turbulent Reynolds numbers is computed numerically via Lighthill's theory and the result is compared with the predictions of Amiet and Curle. The effect of the airfoil thickness is also examined. Supported by ONR Grant N00014-09-1-1088.

  15. Investigation of the Boundary Layer Behavior on Turbine Airfoils.

    DTIC Science & Technology

    1979-08-01

    turbine airfoil cascade . The airfoil profile was based on a turbine blade design used by Lander ’’4 and employed in previous wake studies by Cox and...simulate the wake from upstream turning vanes or blades , a circular cylinder was placed upstream of the centra l or test airfoil . The displacement of this...of turbine airfoil cascade model s by Cox and Han 15 are very much evident in the graph . It might be noted that the blade stag- nation points are at

  16. The conformal transformation of an airfoil into a straight line and its application to the inverse problem of airfoil theory

    NASA Technical Reports Server (NTRS)

    Mutterperl, William

    1944-01-01

    A method of conformal transformation is developed that maps an airfoil into a straight line, the line being chosen as the extended chord line of the airfoil. The mapping is accomplished by operating directly with the airfoil ordinates. The absence of any preliminary transformation is found to shorten the work substantially over that of previous methods. Use is made of the superposition of solutions to obtain a rigorous counterpart of the approximate methods of thin-airfoils theory. The method is applied to the solution of the direct and inverse problems for arbitrary airfoils and pressure distributions. Numerical examples are given. Applications to more general types of regions, in particular to biplanes and to cascades of airfoils, are indicated. (author)

  17. Streamwise Oscillation of Airfoils into Reverse Flow

    NASA Astrophysics Data System (ADS)

    Granlund, Kenneth; Jones, Anya; Ol, Michael

    2015-11-01

    An airfoil in freestream is oscillated in streamwise direction to cyclically enter reverse flow. Measured lift is compared to analytical blade element theories. Advance ratio, reduced frequency and angle of attack is varied within those typical for helicopters. Experimental results reveal that lift does not become negative in the flow reversal part, contradicting one theory and supported by another. Flow visualization reveal the leading edge vortex advecting against the freestream to a point in front of the leading edge.

  18. Linearized propulsion theory of flapping airfoils revisited

    NASA Astrophysics Data System (ADS)

    Fernandez-Feria, R.

    2016-12-01

    A vortical impulse theory is used to compute the thrust force of a plunging and pitching airfoil in forward flight at high Reynolds numbers within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick, which considered only two effects, the leading-edge suction and the projection in the flight direction of the pressure force on the airfoil. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains, in addition to the pressure force projection term, a new term that generalizes the leading-edge suction term in Garrick's theory. This term depends on Theodorsen function C (k ) and on a new complex function C1(k ) of the reduced frequency k . The main qualitative difference with Garrick's theory is that the propulsive efficiency, or ratio of the mean thrust power and the mean input power required to drive the airfoil, tends to zero as the reduced frequency increases to infinity (as k-1), in contrast to Garrick's propulsive efficiency that tends to a constant (1 /2 ). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k →∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining nondimensional parameters. The present analytical results are in good agreement, for small amplitude oscillations, with numerical results from unsteady panel methods, and with experimental data and numerical results from the Navier-Stokes equations, except for small reduced frequencies where viscous effects are obviously important.

  19. Turbulent Flow over Rough Turbine Airfoils.

    DTIC Science & Technology

    1985-08-01

    SUBJECT TERMS (Continue on reverse if necessary and identify by block number) FIELD GROUP SUB. GR. Turbine blades ’ vanes ; surface roughness...turbulent boundary layer over rough turbine vanes or blades is developed. A new formulation of the mixing length model, expressed in the velocity-space...A-163 005 TURBULENT FLOW OVER ROUGH TURBINE AIRFOILS (U) OHIO 1/ STATE UNIV RESEARCH FOUNDATION COLUMBUS L S HAN AUG B5 OSURF-76357/?i4467 AFWL-TR-95

  20. Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise

    NASA Technical Reports Server (NTRS)

    2010-01-01

    Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.

  1. Two-dimensional separated wake modeling and its use to predict maximum section lift coefficient

    NASA Technical Reports Server (NTRS)

    Henderson, M. L.

    1978-01-01

    A technique for computing the lift of separating multielement airfoils in incompressible flow is presented. The procedure employs repeated application of a panel method to solve for the separated wake displacement surface using entirely inviscid boundary conditions. Results are presented that compare computed pressure distributions with those measured in the wind tunnel for airfoils with one, two, and four elements with separation on each element. A method employing this technique is presented which shows promise in predicting airfoil section lift through stall.

  2. Design, Response Rates, and Population Characteristics of a Cross-Sectional Study in Zanzibar, Tanzania

    PubMed Central

    Nyangasa, Maria Adam; Kelm, Soerge; Sheikh, Mohammed Ali

    2016-01-01

    Background Data on nutritional status and correlates of noncommunicable diseases are scarce for resource-poor settings in sub-Saharan countries. With the scope of a project, “Access to Food and Nutrition Status of the Zanzibari Population,” data for investigating public health questions were collected using proven measurement and laboratory standards. Objective The present study aims at providing a descriptive overview of recruitment approaches, standardization, quality control measures, and data collection, with special attention to the design, responses, and participant characteristics of the overall project. Methods A cross-sectional study across 80 randomly selected Shehias (wards) was conducted in 2013 in Unguja Island, Zanzibar. Examinations included all members living in 1 household, face-to-face interviews and anthropometric measurements (weight, height, mid-upper arm circumference, waist and hip circumference, and body composition) were assessed for all household members, blood pressure was taken from participants older than 2 years, and biosamples (urine and blood) from eligible household members were collected. Data collected from the core sample included sociodemographic data, nutritional status, and medical history (hypertension). Physical activity data was collected from a subsample of children between 3 and 16 years of age. Results A total of 1314 participants (mean age 23.6 ± 18.9 years, 54.54% female) completed all anthropometric measurements and were included in the analysis. Out of which, 98.40% (1293/1314) completed the household member’s questionnaire, 93.32% (1229/1314) participants older than 2 years completed blood pressure measurements, and 64.31% (845/1314) blood samples were collected from participants older than 5 years. Underweight prevalence for the total study population was 36.53% (480/1314) with the highest prevalence in children under 14 years. Overweight and obesity was highest among females with the prevalence of 7

  3. LES tests on airfoil trailing edge serration

    NASA Astrophysics Data System (ADS)

    Zhu, Wei Jun; Shen, Wen Zhong

    2016-09-01

    In the present study, a large number of acoustic simulations are carried out for a low noise airfoil with different Trailing Edge Serrations (TES). The Ffowcs Williams-Hawkings (FWH) acoustic analogy is used for noise prediction at trailing edge. The acoustic solver is running on the platform of our in-house incompressible flow solver EllipSys3D. The flow solution is first obtained from the Large Eddy Simulation (LES), the acoustic part is then carried out based on the instantaneous hydrodynamic pressure and velocity field. To obtain the time history data of sound pressure, the flow quantities are integrated around the airfoil surface through the FWH approach. For all the simulations, the chord based Reynolds number is around 1.5x106. In the test matrix, the effects from angle of attack, the TE flap angle, the length/width of the TES are investigated. Even though the airfoil under investigation is already optimized for low noise emission, most numerical simulations and wind tunnel experiments show that the noise level is further decreased by adding the TES device.

  4. Wake structure of a deformable Joukowski airfoil

    NASA Astrophysics Data System (ADS)

    Ysasi, Adam; Kanso, Eva; Newton, Paul K.

    2011-10-01

    We examine the vortical wake structure shed from a deformable Joukowski airfoil in an unbounded volume of inviscid and incompressible fluid. The deformable airfoil is considered to model a flapping fish. The vortex shedding is accounted for using an unsteady point vortex model commonly referred to as the Brown-Michael model. The airfoil’s deformations and rotations are prescribed in terms of a Jacobi elliptic function which exhibits, depending on a dimensionless parameter m, a range of periodic behaviors from sinusoidal to a more impulsive type flapping. Depending on the parameter m and the Strouhal number, one can identify five distinct wake structures, ranging from arrays of isolated point vortices to vortex dipoles and tripoles shed into the wake with every half-cycle of the airfoil flapping motion. We describe these regimes in the context of other published works which categorize wake topologies, and speculate on the importance of these wake structures in terms of periodic swimming and transient maneuvers of fish.

  5. Navier-Stokes calculations and turbulence modeling in the trailing edge region of a circulation control airfoil

    NASA Technical Reports Server (NTRS)

    Viegas, John R.; Rubesin, Morris W.; Maccormack, Robert W.

    1987-01-01

    The accurate prediction of turbulent flows over curved surfaces in general and over the trailing edge region of circulation control airfoils in particular requires the coupled efforts of turbulence modelers, numerical analysts and experimentalists. The purpose of the research program in this area is described. Then, the influence on turbulence modeling of the flow characteristics over a typical circulation control wing is discussed. Next, the scope of this effort to study turbulence in the trailing edge region of a circulation control airfoil is presented. This is followed by a brief overview of the computation scheme, including the grid, governing equations, numerical method, boundary conditions and turbulence models applied to date. Then, examples of applications of two algebraic eddy viscosity models to the trailing edge region of a circulation control airfoil is presented. The results from the calculations is summarized, and conclusions drawn based on examples. Finally, the future directions of the program is outlined.

  6. Computation of full-coverage film-cooled airfoil temperatures by two methods and comparison with high heat flux data

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.; Yeh, F. C.; Austin, P. J., Jr.

    1987-01-01

    Two methods were used to calculate the heat flux to full-coverage film cooled airfoils and, subsequently, the airfoil wall temperatures. The calculated wall temperatures were compared to measured temperatures obtained in the Hot Section Facility operating at real engine conditions. Gas temperatures and pressures up to 1900 K and 18 atm with a Reynolds number up to 1.9 million were investigated. Heat flux was calculated by the convective heat transfer coefficient adiabatic wall method and by the superposition method which incorporates the film injection effects in the heat transfer coefficient. The results of the comparison indicate the first method can predict the experimental data reasonably well. However, superposition overpredicted the heat flux to the airfoil without a significant modification of the turbulent Prandtl number. The results suggest that additional research is required to model the physics of full-coverage film cooling where there is significant temperature/density differences between the gas and the coolant.

  7. Numerical Study of Ram Air Airfoils and Upper Surface Bleed-Air Control

    DTIC Science & Technology

    2014-06-16

    of ram -air parachute systems to complement the design and analysis of new and existing airdrop systems. In this paper an unsteady numerical study of...two-dimensional, rigid, ram -air sections with an array of upper surface bleed-air actuators is presented. Aerodynamic forces and lift-to-drag ratios of...a modified Clark-Y ram -air airfoil are calculated from unsteady Reynolds-Averaged Navier-Stokes (RANS) simulations, using the Kestrel and Cobalt flow

  8. Performance and test section flow characteristics of the National Full-Scale Aerodynamics Complex 80- by 120-Foot Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Zell, Peter T.

    1993-01-01

    Results from the performance and test section flow calibration of the 80- by 120-Foot Wind Tunnel are presented. Measurements indicating the 80- by 120-ft test section flow quality were obtained throughout the tunnel operational envelope and for atmospheric wind speeds up to approximately 20 knots. Tunnel performance characteristics and a dynamic pressure system calibration were also documented during the process of mapping the test section flow field. Experimental results indicate that the test section flow quality is relatively insensitive to dynamic pressure and the level of atmospheric winds experienced during the calibration. The dynamic pressure variation in the test section is within +/-75 percent of the average. The axial turbulence intensity is less than 0.5 percent up to the maximum test section speed of 100 knots, and the vertical and lateral flow angle variations are within +/-5 deg and +/-7 deg, respectively. Atmospheric winds were found to affect the pressure distribution in the test section only at high ratios of wind speed to test section speed.

  9. Correlations Among Ice Measurements, Impingement Rates Icing Conditions, and Drag Coefficients for Unswept NACA 65A004 Airfoil

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H.

    1958-01-01

    An empirical relation has been obtained by which the change in drag coefficient caused by ice formations on an unswept NACA 65AO04 airfoil section can be determined from the following icing and operating conditions: icing time, airspeed, air total temperature, liquid-water content, cloud droplet impingement efficiencies, airfoil chord length, and angles of attack. The correlation was obtained by use of measured ice heights and ice angles. These measurements were obtained from a variety of ice formations, which were carefully photographed, cross-sectioned, and weighed. Ice weights increased at a constant rate with icing time in a rime icing condition and at progressively increasing rates in glaze icing conditions. Initial rates of ice collection agreed reasonably well with values predicted from droplet impingement data. Experimental droplet impingement rates obtained on this airfoil section agreed with previous theoretical calculations for angles of attack of 40 or less. Disagreement at higher angles of attack was attributed to flow separation from the upper surface of the experimental airfoil model.

  10. Toward large eddy simulation of turbulent flow over an airfoil

    NASA Technical Reports Server (NTRS)

    Choi, Haecheon

    1993-01-01

    The flow field over an airfoil contains several distinct flow characteristics, e.g. laminar, transitional, turbulent boundary layer flow, flow separation, unstable free shear layers, and a wake. This diversity of flow regimes taxes the presently available Reynolds averaged turbulence models. Such models are generally tuned to predict a particular flow regime, and adjustments are necessary for the prediction of a different flow regime. Similar difficulties are likely to emerge when the large eddy simulation technique is applied with the widely used Smagorinsky model. This model has not been successful in correctly representing different turbulent flow fields with a single universal constant and has an incorrect near-wall behavior. Germano et al. (1991) and Ghosal, Lund & Moin have developed a new subgrid-scale model, the dynamic model, which is very promising in alleviating many of the persistent inadequacies of the Smagorinsky model: the model coefficient is computed dynamically as the calculation progresses rather than input a priori. The model has been remarkably successful in prediction of several turbulent and transitional flows. We plan to simulate turbulent flow over a '2D' airfoil using the large eddy simulation technique. Our primary objective is to assess the performance of the newly developed dynamic subgrid-scale model for computation of complex flows about aircraft components and to compare the results with those obtained using the Reynolds average approach and experiments. The present computation represents the first application of large eddy simulation to a flow of aeronautical interest and a key demonstration of the capabilities of the large eddy simulation technique.

  11. Sealing apparatus for airfoils of gas turbine engines

    DOEpatents

    Jones, R.B.

    1998-05-19

    An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed. 17 figs.

  12. Sealing apparatus for airfoils of gas turbine engines

    DOEpatents

    Jones, Russell B.

    1998-01-01

    An improved airfoil tip sealing apparatus is disclosed wherein brush seals are attached to airfoil tips with the distal ends of the brush seal fibers sealingly contacting opposing wall surfaces. Embodiments for variable vanes, stators and both cooled and uncooled turbine blade applications are disclosed.

  13. Numerical Airfoil Optimization Using a Reduced Number of Design Coordinates

    NASA Technical Reports Server (NTRS)

    Vanderplaats, G. N.; Hicks, R. M.

    1976-01-01

    A method is presented for numerical airfoil optimization whereby a reduced number of design coordinates are used to define the airfoil shape. The approach is to define the airfoil as a linear combination of shapes. These basic shapes may be analytically or numerically defined, allowing the designer to use his insight to propose candidate designs. The design problem becomes one of determining the participation of each such function in defining the optimum airfoil. Examples are presented for two-dimensional airfoil design and are compared with previous results based on a polynomial representation of the airfoil shape. Four existing NACA airfoils are used as basic shapes. Solutions equivalent to previous results are achieved with a factor of more than 3 improvements in efficiency, while superior designs are demonstrated with an efficiency greater than 2 over previous methods. With this shape definition, the optimization process is shown to exploit the simplifying assumptions in the inviscid aerodynamic analysis used here, thus demonstrating the need to use more advanced aerodynamics for airfoil optimization.

  14. TRANDES: A FORTRAN program for transonic airfoil analysis or design

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1977-01-01

    A program called TRANDES is presented that is used for the analysis of steady, irrotational transonic flow over specified two-dimensional airfoils in free air or for the design of airfoils having a prescribed pressure distribution, including the effects of weak viscous interaction. Instructions on program usage, listings of the program, and sample cases are given.

  15. High Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.

    1985-01-01

    A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

  16. Complex Flow Separation Pattern on Transonic Fan Airfoils Revealed by Flow Visualization

    NASA Technical Reports Server (NTRS)

    Lepicovsky, Jan

    2001-01-01

    Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. The latest view is that the blade oscillations are triggered by high-frequency changes in the extent of the partially separated area on the airfoil suction side. There is a lack of experimental data describing the separated flow characteristics of modern airfoils for transonic fans.

  17. Large-eddy simulation of flow around an airfoil on a structured mesh

    NASA Technical Reports Server (NTRS)

    Kaltenbach, Hans-Jakob; Choi, Haecheon

    1995-01-01

    The diversity of flow characteristics encountered in a flow over an airfoil near maximum lift taxes the presently available statistical turbulence models. This work describes our first attempt to apply the technique of large-eddy simulation to a flow of aeronautical interest. The challenge for this simulation comes from the high Reynolds number of the flow as well as the variety of flow regimes encountered, including a thin laminar boundary layer at the nose, transition, boundary layer growth under adverse pressure gradient, incipient separation near the trailing edge, and merging of two shear layers at the trailing edge. The flow configuration chosen is a NACA 4412 airfoil near maximum lift. The corresponding angle of attack was determined independently by Wadcock (1987) and Hastings & Williams (1984, 1987) to be close to 12 deg. The simulation matches the chord Reynolds number U(sub infinity)c/v = 1.64 x 10(exp 6) of Wadcock's experiment.

  18. Geometry Modeling and Grid Generation for Computational Aerodynamic Simulations Around Iced Airfoils and Wings

    NASA Technical Reports Server (NTRS)

    Choo, Yung K.; Slater, John W.; Vickerman, Mary B.; VanZante, Judith F.; Wadel, Mary F. (Technical Monitor)

    2002-01-01

    Issues associated with analysis of 'icing effects' on airfoil and wing performances are discussed, along with accomplishments and efforts to overcome difficulties with ice. Because of infinite variations of ice shapes and their high degree of complexity, computational 'icing effects' studies using available software tools must address many difficulties in geometry acquisition and modeling, grid generation, and flow simulation. The value of each technology component needs to be weighed from the perspective of the entire analysis process, from geometry to flow simulation. Even though CFD codes are yet to be validated for flows over iced airfoils and wings, numerical simulation, when considered together with wind tunnel tests, can provide valuable insights into 'icing effects' and advance our understanding of the relationship between ice characteristics and their effects on performance degradation.

  19. Numerical Simulation of the Processes of Icing on Airfoils with Formation of a "Barrier" Ice

    NASA Astrophysics Data System (ADS)

    Prikhod'ko, A. A.; Alekseenko, S. V.

    2014-05-01

    Software and methods allowing one to model the processes of formation of a "barrier" ice on the unprotected part of an airfoil have been developed with the use of the Reynolds-averaged Navier-Stokes equations for a compressible gas, which are closed with the aid of the Spalart-Allmaras model of turbulence. An inertial model is used to describe the motion of overcooled water droplets. In modeling the process of ice accretion, differential equations of mass, momentum, and energy conservation are used for each element of the surface. The initial equations are made discrete by means of the control volume approach. The influence of the height of ice accretions and of their location on the character of air-droplet flow past a NACA 0012 airfoil and on its aerodynamic characteristics has been analyzed.

  20. Transonic shock-induced dynamics of a flexible wing with a thick circular-arc airfoil

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Rivera, Jose A., Jr.

    1991-01-01

    Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.

  1. A systematic method for computer design of supercritical airfoils in cascade

    NASA Technical Reports Server (NTRS)

    Garabedian, P.; Korn, D.

    1976-01-01

    A computer code has been developed for the direct calculation of shockless transonic airfoils whose pressure distributions can be assigned within reasonable limits. The partial differential equations of two-dimensional inviscid gas dynamics are solved by analytic continuation into the domain of two independent complex characteristic coordinates. The domain of integration is mapped conformally onto the unit circle in the hodograph plane of one of these coordinates. It is possible to formulate a boundary value problem on this circle for the stream function that is well posed in the case of transonic flow. This enables the formulation of a procedure for the calculation of an airfoil on which the speed is prescribed as a function of the arc length

  2. Navier-Stokes analysis of airfoils with leading edge ice accretions

    NASA Technical Reports Server (NTRS)

    Potapczuk, Mark G.

    1993-01-01

    A numerical analysis of the flowfield characteristics and the performance degradation of an airfoil with leading edge ice accretions was performed. The important fluid dynamic processes were identified and calculated. Among these were the leading edge separation bubble at low angles of attack, complete separation on the low pressure surface resulting in premature shell, drag rise due to the ice shape, and the effects of angle of attack on the separated flow field. Comparisons to experimental results were conducted to confirm these calculations. A computer code which solves the Navier-Stokes equations in two dimensions, ARC2D, was used to perform the calculations. A Modified Mixing Length turbulence model was developed to produce grids for several ice shape and airfoil combinations. Results indicate that the ability to predict overall performance characteristics, such as lift and drag, at low angles of attack is excellent. Transition location is important for accurately determining separation bubble shape. Details of the flowfield in and downstream of the separated regions requires some modifications. Calculations for the stalled airfoil indicate periodic shedding of vorticity that was generated aft of the ice accretion. Time averaged pressure values produce results which compare favorably with experimental information. A turbulence model which accounts for the history effects in the flow may be justified.

  3. Controlled Aeroelastic Response and Airfoil Shaping Using Adaptive Materials and Integrated Systems

    NASA Technical Reports Server (NTRS)

    Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer

    1996-01-01

    This paper presents an overview of several activities of the Aeroelasticity Branch at the NASA Langley Research Center in the area of applying adaptive materials and integrated systems for controlling both aircraft aeroelastic response and airfoil shape. The experimental results of four programs are discussed: the Piezoelectric Aeroelastic Response Tailoring Investigation (PARTI); the Adaptive Neural Control of Aeroelastic Response (ANCAR) program; the Actively Controlled Response of Buffet Affected Tails (ACROBAT) program; and the Airfoil THUNDER Testing to Ascertain Characteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant rcductions in aeroelastic response at dynamic pressures below flutter using piezoelectric actuators. The ANCAR program seeks to demonstrate the effectiveness of using neural networks to schedule flutter suppression control laws. Th,e ACROBAT program studied the effectiveness of a number of candidate actuators, including a rudder and piezoelectric actuators, to alleviate vertical tail buffeting. In the ATTACH project, the feasibility of using Thin-Layer Composite-Uimorph Piezoelectric Driver and Sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.

  4. Characterization of dynamic stall on 9-15 % thick airfoils using experiment and computation

    NASA Astrophysics Data System (ADS)

    Davidson, Phillip B.

    In recent years, the blade geometry on wind turbines and helicopters has been optimized for a particular span location. Unsteady flow phenomena like dynamic stall limit these designs and need to be better understood and correctly simulated. Currently, empirical and computational fluid dynamics (CFD) methods are used to simulate rotating wind turbine or helicopter blades, but each of these methods has limitations in predicting unsteady separated flows. To address these needs, the present work investigated oscillating airfoils over a range of conditions with an approach that provided fast, low-cost unsteady pressure data combined with a highly resolved flow field to better understand the physics of dynamic stall. An additional objective was to show how such data may be used to assess CFD simulations. This research has yielded interesting results showing characteristics of thin airfoil stall, leading edge stall, and trailing edge stall that were sorted and classified. Classification of the oscillating airfoil behavior with or without dynamic stall was performed using previous definitions for stall regime, separation characteristics, and other qualitative differences in stall pattern. After classifying the unsteady flow for each of the cases, comparison of experimental results and results obtained using an unsteady Reynolds Averaged Navier-Stokes (URANS) solver was performed to assess the ability of the solver to produce the same unsteady effects. Although both experiment and computation produced similar flow features, the timing and magnitude of the features in the dynamic stall and re-attachment process of the pitching cycle exhibited some significant differences.

  5. Experimental Investigation of Wind-Tunnel Interference on the Downwash Behind an Airfoil

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe; Katzoff, S

    1937-01-01

    The interference of the wind-tunnel boundaries on the downwash behind an airfoil has been experimentally investigated and the results have been compared with the available theoretical results for open-throat wind tunnels. As in previous studies, the simplified theoretical treatment that assumes the test section to be an infinite free jet has been shown to be satisfactory at the lifting line. The experimental results, however, show that this assumption may lead to erroneous conclusions regarding the corrections to be applied to the downwash in the region behind the airfoil where the tail surfaces are normally located. The results of a theory based on the more accurate concept of the open-jet wind tunnel as a finite length of free jet provided with a closed exit passage are in good qualitative agreement with the experimental results.

  6. Application of numerical optimization to the design of supercritical airfoils without drag-creep

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Vanderplaats, G. N.

    1977-01-01

    Recent applications of numerical optimization to the design of advanced airfoils for transonic aircraft have shown that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while the drag at the design Mach number is the objective function. Such a procedure doubles the computation time over that for single design-point problems, but the final result is worth the increased cost of computation. The ability to treat such multiple design-point problems by numerical optimization has been enhanced by the development of improved airfoil shape functions. Such functions permit a considerable increase in the range of profiles attainable during the optimization process.

  7. Investigation of passive shock wave-boundary layer control for transonic airfoil drag reduction

    NASA Technical Reports Server (NTRS)

    Nagamatsu, H. T.; Brower, W. B., Jr.; Bahi, L.; Ross, J.

    1982-01-01

    The passive drag control concept, consisting of a porous surface with a cavity beneath it, was investigated with a 12-percent-thick circular arc and a 14-percent-thick supercritical airfoil mounted on the test section bottom wall. The porous surface was positioned in the shock wave/boundary layer interaction region. The flow circulating through the porous surface, from the downstream to the upstream of the terminating shock wave location, produced a lambda shock wave system and a pressure decrease in the downstream region minimizing the flow separation. The wake impact pressure data show an appreciably drag reduction with the porous surface at transonic speeds. To determine the optimum size of porosity and cavity, tunnel tests were conducted with different airfoil porosities, cavities and flow Mach numbers. A higher drag reduction was obtained by the 2.5 percent porosity and the 1/4-inch deep cavity.

  8. Statistical characteristics of gassy sediment sections in the northwest Gulf of Mexico

    SciTech Connect

    Bryant, W.R.; Anderson, A.L.; Orsi, T.H.

    1994-09-01

    Examination of seismic data from over 1000 MMS geohazard reports and core logs of 1670 foundation boreholes (drilled to an average subbottom depth of 125 m), has revealed that gassy sediment sections are most abundant near the Mississippi River delta, in buried stream channels eroded during the early and late Wisconsinan, and in Miocene and Pliocene-Pleistocene depocenters on the shelf and upper slope. Of all the boreholes that we examined, 1158 (68%) contained indications of gassy sediments. The {open_quotes}average{close_quotes} Mississippi River delta related gassy sediment section occurs at a water depth of 53 m; the top of the section lies at a subbottom depth of 10 m; its thickness is 12 m; and the sediment water content is 45%, with liquid and plastic limits of 73% and 24%, respectively. For the nondeltaic areas of the continental shelf (areas west of 90{degrees}W), the average gassy sediment section lies at a water depth of 37 m, its depth below the sea floor is 37 m, its thickness is 8 m, the water content is 40%, and the liquid and plastic limits are 68 and 27%. For the upper continental slope in the northwestern Gulf of Mexico, the average gassy sediment section is found at a water depth of 265 m, is 25 m below the sea floor, and has a thickness of 60 in, a water content of 40%, and a liquid and plastic limit of 62 and 32%. The median areal extent of gassy sediment sections within the entire shelf and upper slope of the northwestern Gulf area ranges from 238 to 546 m. Although large patches of gassy sediments exist, with some exceeding 10 km in size, most are less than 500 m.

  9. A CFD Database for Airfoils and Wings at Post-Stall Angles of Attack

    NASA Technical Reports Server (NTRS)

    Petrilli, Justin; Paul, Ryan; Gopalarathnam, Ashok; Frink, Neal T.

    2013-01-01

    This paper presents selected results from an ongoing effort to develop an aerodynamic database from Reynolds-Averaged Navier-Stokes (RANS) computational analysis of airfoils and wings at stall and post-stall angles of attack. The data obtained from this effort will be used for validation and refinement of a low-order post-stall prediction method developed at NCSU, and to fill existing gaps in high angle of attack data in the literature. Such data could have potential applications in post-stall flight dynamics, helicopter aerodynamics and wind turbine aerodynamics. An overview of the NASA TetrUSS CFD package used for the RANS computational approach is presented. Detailed results for three airfoils are presented to compare their stall and post-stall behavior. The results for finite wings at stall and post-stall conditions focus on the effects of taper-ratio and sweep angle, with particular attention to whether the sectional flows can be approximated using two-dimensional flow over a stalled airfoil. While this approximation seems reasonable for unswept wings even at post-stall conditions, significant spanwise flow on stalled swept wings preclude the use of two-dimensional data to model sectional flows on swept wings. Thus, further effort is needed in low-order aerodynamic modeling of swept wings at stalled conditions.

  10. Effects of Small Angles of Sweep and Moderate Amounts of Dihedral on Stalling and Lateral Characteristics of a Wing-Fuselage Combination Equipped with Partial- and Full-Span Double Slotted Flaps

    NASA Technical Reports Server (NTRS)

    Teplitz, Jerome

    1944-01-01

    Tests of a wing-fuselage combinations incorporating NACA 65-series airfoil sections were conducted in the NACA 19-foot pressure tunnel. The investigation included the tests with flaps neutral and with partial- and full-span double slotted flaps deflected to determine the effects of (1) variations of wing sweep between -4 degrees and 8 degrees on stalling and lateral stability and control characteristics and (2) variations of dihedral between 0 degree and 6.75 degrees on lateral stability characteristics.

  11. Multiple element airfoils optimized for maximum lift coefficient.

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.; Chen, A. W.

    1972-01-01

    Optimum airfoils in the sense of maximum lift coefficient are obtained for incompressible fluid flow at large Reynolds number. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the airfoil upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution is a function of Reynolds number and the trailing edge velocity. Geometries of those airfoils which will generate these optimum pressure distributions are obtained using a direct-iterative method which is developed in this study. This method can be used to design airfoils consisting of any number of elements. Numerical examples of one- and two-element airfoils are given. The maximum lift coefficients obtained range from 2 to 2.5.

  12. Transition effects on compressible dynamic stall of transiently pitching airfoils

    NASA Technical Reports Server (NTRS)

    Wilder, M. C.; Chandrasekhara, M. S.; Carr, L. W.

    1993-01-01

    Experimental results and analysis of the effects of boundary layer tripping on dynamic stall of a transiently pitching airfoil are presented. At low Mach numbers, the tripped airfoil exhibits qualitative similarity with the behavior of the untripped airfoil. However, the local supersonic flow at Mach numbers greater than 0.3 is significantly modified by the trip leading to vastly different shock/boundary layer interactions, dynamic stall onset and vortex formation angles. The formation of the laminar separation bubble is found to have a favorable influence in delaying dynamic stall on the untripped airfoil flow. In both Mach number regimes, the tripped flow actually stalls at slightly lower angles of attack. Further experimentation with three trips on an oscillating airfoil showed that the dynamic stall process is very sensitive to the state of the turbulence in the boundary layer. This sensitivity points to a need for finer turbulence modeling techniques for use in dynamic stall computations.

  13. Effect of High-Fidelity Ice Accretion Simulations on the Performance of a Full-Scale Airfoil Model

    NASA Technical Reports Server (NTRS)

    Broeren, Andy P.; Bragg, Michael B.; Addy, Harold E., Jr.; Lee, Sam; Moens, Frederic; Guffond, Didier

    2010-01-01

    The simulation of ice accretion on a wing or other surface is often required for aerodynamic evaluation, particularly at small scale or low-Reynolds number. While there are commonly accepted practices for ice simulation, there are no established and validated guidelines. The purpose of this article is to report the results of an experimental study establishing a high-fidelity, full-scale, iced-airfoil aerodynamic performance database. This research was conducted as a part of a larger program with the goal of developing subscale aerodynamic simulation methods for iced airfoils. Airfoil performance testing was carried out at the ONERA F1 pressurized wind tunnel using a 72-in. (1828.8-mm) chord NACA 23012 airfoil over a Reynolds number range of 4.5x10(exp 6) to 16.0 10(exp 6) and a Mach number range of 0.10 to 0.28. The high-fidelity, ice-casting simulations had a significant impact on the aerodynamic performance. A spanwise-ridge ice shape resulted in a maximum lift coefficient of 0.56 compared to the clean value of 1.85 at Re = 15.9x10(exp 6) and M = 0.20. Two roughness and streamwise shapes yielded maximum lift values in the range of 1.09 to 1.28, which was a relatively small variation compared to the differences in the ice geometry. The stalling characteristics of the two roughness and one streamwise ice simulation maintained the abrupt leading-edge stall type of the clean NACA 23012 airfoil, despite the significant decrease in maximum lift. Changes in Reynolds and Mach number over the large range tested had little effect on the iced-airfoil performance.

  14. Characteristics and Dental Experiences of Autistic Children in Saudi Arabia: Cross-Sectional Study

    ERIC Educational Resources Information Center

    Murshid, Ebtissam Z.

    2011-01-01

    The purpose of this study is to report base line information about characteristics, and dental experiences of a group of autistic children in three major cities of Saudi Arabia. Most of the children (76.2%) included in the study were diagnosed with autism before the age of 5 years. More than half of the children (53.7%) had no previous dental…

  15. Nailfold capillary morphological characteristics of hand-arm vibration syndrome: a cross-sectional study

    PubMed Central

    Chen, QingSong; Chen, GuiPing; Xiao, Bin; Lin, HanSheng; Qu, HongYing; Zhang, DanYing; Shi, MaoGong; Lang, Li; Yang, Bei; Yan, MaoSheng

    2016-01-01

    Objective The purpose of this study was to investigate the characteristics of nailfold capillaroscopy associated with hand-arm vibration syndrome (HAVS). Methods In total, 113 male gold miners were recruited: 35 workers who were chronically exposed to vibration and developed vibration-induced white finger were defined as the HAVS group, 39 workers who were exposed to vibration but did not have HAVS were classified as the vibration-exposed controls (VEC) group, and 39 workers without vibration exposure were categorised as the non-VEC (NVEC) group. Video capillaroscopy was used to capture images of the 2nd, 3rd and 4th fingers of both hands. The following nailfold capillary characteristics were included: number of capillaries/mm, avascular areas, haemorrhages and enlarged capillaries. The experiments were carried out in the same winter. All characteristics were evaluated under blinded conditions. Results Significant differences in all morphological characteristics existed between the groups (p<0.05). Avascular areas in the HAVS, VEC and NVEC groups appeared in 74.3%, 43.6% and 25.0% of participants, respectively. A higher percentage of participants had haemorrhages in the HAVS group (65.7%) compared with the other groups (VEC: 7.7% and NVEC: 7.5%). The number of capillaries/mm, input limb width, output limb width, apical width, and ratio of output limb and input limb all had more than 70% sensitivity or specificity of their cut-off value. Conclusions Nailfold capillary characteristics, especially the number of capillaries/mm, avascular areas, haemorrhages, output limb width, input limb width and apical width alterations, revealed significant associations with HAVS. PMID:27888176

  16. Reversible airfoils for stopped rotors in high speed flight

    NASA Astrophysics Data System (ADS)

    Niemiec, Robert; Jacobellis, George; Gandhi, Farhan

    2014-10-01

    This study starts with the design of a reversible airfoil rib for stopped-rotor applications, where the sharp trailing-edge morphs into the rounded leading-edge, and vice-versa. A NACA0012 airfoil is approximated in a piecewise linear manner and straight, rigid outer profile links used to define the airfoil contour. The end points of the profile links connect to control links, each set on a central actuation rod via an offset. Chordwise motion of the actuation rod moves the control and the profile links and reverses the airfoil. The paper describes the design methodology and evolution of the final design, based on which two reversible airfoil ribs were fabricated and used to assemble a finite span reversible rotor/wing demonstrator. The profile links were connected by Aluminum strips running in the spanwise direction which provided stiffness as well as support for a pre-tensioned elastomeric skin. An inter-rib connector with a curved-front nose piece supports the leading-edge. The model functioned well and was able to reverse smoothly back-and-forth, on application and reversal of a voltage to the motor. Navier-Stokes CFD simulations (using the TURNS code) show that the drag coefficient of the reversible airfoil (which had a 13% maximum thickness due to the thickness of the profile links) was comparable to that of the NACA0013 airfoil. The drag of a 16% thick elliptical airfoil was, on average, about twice as large, while that of a NACA0012 in reverse flow was 4-5 times as large, even prior to stall. The maximum lift coefficient of the reversible airfoil was lower than the elliptical airfoil, but higher than the NACA0012 in reverse flow operation.

  17. Approximation concepts for numerical airfoil optimization

    NASA Technical Reports Server (NTRS)

    Vanderplaats, G. N.

    1979-01-01

    An efficient algorithm for airfoil optimization is presented. The algorithm utilizes approximation concepts to reduce the number of aerodynamic analyses required to reach the optimum design. Examples are presented and compared with previous results. Optimization efficiency improvements of more than a factor of 2 are demonstrated. Improvements in efficiency are demonstrated when analysis data obtained in previous designs are utilized. The method is a general optimization procedure and is not limited to this application. The method is intended for application to a wide range of engineering design problems.

  18. Direct Numerical Simulations of Plunging Airfoils

    DTIC Science & Technology

    2010-01-07

    Schmidt and E Turkel, Numerical Solutions of the Euler Equations by Finite Volume Methods Using Runge-Kutta Time-Stepping Schemes, AIAA paper 81-1259...Ω ( p ∂vj ∂xj − σij ∂v i ∂xj ) dV (4) Definition 1 A numerical scheme to solve the viscous Navier-Stokes equations is said to be Kinetic Energy...Direct Numerical Simulations of Plunging Airfoils Yves Allaneau∗ and Antony Jameson† Stanford University, Stanford, California, 94305, USA This paper

  19. Comparison of CDOM EEMs Characteristics along F and PN section in Eastern China Sea: significance for sources tracing

    NASA Astrophysics Data System (ADS)

    Du, Yong; Zhang, Xiaoyu; Jiang, Binbin; Huang, Dasong; Yao, Lingling

    2015-04-01

    In this paper, a total of 28 water samples were collected mainly from three sections(C section in the Yangtze river inner estuary, PN section and F section on the spindle of Changjiang diluted water influenced by different hydrodynamic processes),which taken on two cruises in spring and summer of 2011. Absorption and fluorescence spectroscopy were measured along with dissolved organic carbon(DOC) concentrations and temperature, salinity and another environmental parameters to characterize the material sources and environmental implications of dissolved organic matter(DOM). Two protein-like components(tyrosine-like peak B and tryptophan-like peak T1), and two humic-like components(marine humic-like peak M and ultraviolet region humic-like peak A ) were identified by PARAFAC. We discussed CDOM distribution characteristic, material composition, and influence factors during the slowly dilution process of Changjiang diluted water into the east China sea by comparing the correlation of the CDOM absorption, fluorescence intensity, and fluorescence peak with DOC, in order to provide the based biogeochemistry theory basis for building DOC implications using CDOM fluorescence properties. The results revealed that:1) the Yangtze river and its inner estuary (upstream of the river mouth) were detected a higher amount of humic-like components. With the rapid dilution (or settlement) at the inner estuary, the humic-like components would further spread and dilute slowly on PN section and F section. On PN section, the terrigenous material is the main source material, and the main mechanism of CDOM distribution characteristics is controlled by dilution diffusion. Affected by the water mass convergence, marine dissolved organic matter in local waters had obvious input. However, due to the complexed hydrodynamic environment on F section, the input of terrigenous material has many ways. The influence of marine dissolved organic matter increased with the offshore distance increases.2

  20. Effects of surface roughness and vortex generators on the LS(1)-0417MOD airfoil

    SciTech Connect

    Reuss, R.L.; Hoffman, M.J.; Gregorek, G.M.

    1995-12-01

    An 18-inch constant-chord model of the LS(l)-0417MOD airfoil section was tested under two dimensional steady state conditions ate University 7{times}10 Subsonic Wind Tunnel. The objective was to document section lift and moment characteristics model and air flow conditions. Surface pressure data was acquired at {minus}60{degrees} through + 230{degrees} geometric angles of attack, at a nominal 1 million Reynolds number. Cases with and without leading edge grit roughness were investigated. The leading edge mulated blade conditions in the field. Additionally, surface pressure data were acquired for Reynolds numbers of 1.5 and 2.0 million, with and without leading edge grit roughness; the angle of attack was limited to a {minus}20{degrees} to 40{degrees} range. In general, results showed lift curve slope sensitivities to Reynolds number and roughness. The maximum lift coefficient was reduced as much as 29% by leading edge roughness. Moment coefficient showed little sensitivity to roughness beyond 50{degrees} angle of attack, but the expected decambering effect of a thicker boundary layer with roughness did show at lower angles. Tests were also conducted with vortex generators located at the 30% chord location on the upper surface only, at 1 and 1.5 million Reynolds numbers, with and without leading edge grit roughness. In general, with leading edge grit roughness applied, the vortex generators restored 85 percent of the baseline level of maximum lift coefficient but with a more sudden stall break and at a higher angle of attack than the baseline.