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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. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. An Experimental Study of Airfoil Icing Characteristics

    NASA Technical Reports Server (NTRS)

    Shaw, R. J.; Sotos, R. G.; Solano, F. R.

    1982-01-01

    A full scale general aviation wing with a NACA 63 sub 2 A415 airfoil section was tested to determine icing characteristics for representative rime and glaze icing conditions. Measurements were made of ice accretion shapes and resultant wing section drag coefficient levels. It was found that the NACA 63 sub 2 A415 wing section was less sensitive to rime and glaze icing encounters for climb conditions.

  12. Aerodynamic Characteristics of Airfoils at High Speeds

    NASA Technical Reports Server (NTRS)

    Briggs, L J; Hull, G F; Dryden, H L

    1925-01-01

    This report deals with an experimental investigation of the aerodynamical characteristics of airfoils at high speeds. Lift, drag, and center of pressure measurements were made on six airfoils of the type used by the air service in propeller design, at speeds ranging from 550 to 1,000 feet per second. The results show a definite limit to the speed at which airfoils may efficiently be used to produce lift, the lift coefficient decreasing and the drag coefficient increasing as the speed approaches the speed of sound. The change in lift coefficient is large for thick airfoil sections (camber ratio 0.14 to 0.20) and for high angles of attack. The change is not marked for thin sections (camber ratio 0.10) at low angles of attack, for the speed range employed. At high speeds the center of pressure moves back toward the trailing edge of the airfoil as the speed increases. The results indicate that the use of tip speeds approaching the speed of sound for propellers of customary design involves a serious loss in efficiency.

  13. Theoretical supersonic lift and drag characteristics of symmetrical wedge-shape-airfoil sections as affected by sweepback outside the Mach cone

    NASA Technical Reports Server (NTRS)

    Ivey, H Reese; Bowen, Edward N , Jr

    1947-01-01

    The theoretical supersonic section lift and drag characteristics of thin wedge-shape, untapered airfoils with sweepback are presented. The results apply to those parts of the wing in two dimensional flow and are not applicable to wings swept back within the Mach cone of the center section. The results may also be applied to swept-forward wings if the angle of sweep is not enough to put the wing within the Mach cone from the tips.

  14. 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.

  15. 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.

  16. Scale and Turbulence Effects on the Lift and Drag Characteristics of the NACA 65(Sub 3)-418, A=1.0 Airfoil Section

    NASA Technical Reports Server (NTRS)

    Tucker, Warren A.; Quinn, John H.

    1944-01-01

    Wind-tunnel tests, investigating low drag wing performance in small-scale tests, showed a large increase in minimum drag coefficient, and a decrease of maximum lift coefficient occurred with decreasing Reynolds Number above certain designated values. The lift-curve slope varied up to 6% between high and low turbulence levels. Low Reynolds Number test data are unreliable for low drag airfoils either to estimate full-scale characteristics or to determine merits of airfoils for higher Reynolds numbers.

  17. 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.

  18. 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).

  19. 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.

  20. 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.

  1. A Systematic Investigation of Pressure Distributions at High Speeds over Five Representative NACA Low-Drag and Conventional Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Graham, Donald J; Nitzberg, Gerald E; Olson, Robert N

    1945-01-01

    Pressure distributions determined from high-speed wind-tunnel tests are presented for five NACA airfoil sections representative of both low-drag and conventional types. Section characteristics of lift, drag, and quarter-chord pitching moment are presented along with the measured pressure distributions for the NACA 65sub2-215 (a=0.5), 66sub2-215 (a=0.6), 0015, 23015, and 4415 airfoils for Mach numbers up to approximately 0.85. A critical study is made of the airfoil pressure distributions in an attempt to formulate a set of general criteria for defining the character of high speed flows over typical airfoil shapes. Comparisons are made of the relative characteristics of the low-drag and conventional airfoils investigated insofar as they would influence the high-speed performance and the high-speed stability and control characteristics of airplanes employing these wing sections.

  2. Aerodynamic characteristics of airfoils with ice accretions

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.; Gregorek, G. M.

    1982-01-01

    Results of a wind tunnel test to evaluate the performance of an airfoil with simulated rime ice are presented with theoretical comparisons. A NACA 65A413 airfoil was tested in the OSU 6 x 22 inch Transonic Airfoil Wind Tunnel at a Reynolds number near three million and Mach numbers from 0.20 to 0.80. The model was tested in four configurations to determine the aero-dynamic effects of the roughness and shape of a rime ice accretion. The simulated rime ice shape was obtained analytically using a time-stepping dry ice accretion computer code. Lift, drag, moment coefficients, and pressure distributions for the clean and simulated rime ice cases are reported. The measured degradation in airfoil performance is compared to an analytical method which uses existing airfoil analysis computer codes with empirical corrections for the surface roughness. A discussion of the empirical surface roughness correction and uses of other airfoil computer methods is included.

  3. 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.

  4. 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.

  5. Airfoil shape for flight at subsonic speeds. [design analysis and aerodynamic characteristics of the GAW-1 airfoil

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T. (Inventor)

    1976-01-01

    An airfoil is examined that has 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. Diagrams illustrating supersonic flow and shock waves over the airfoil are shown.

  6. 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.

  7. Acoustic radiation and surface pressure characteristics of an airfoil due to incident turbulence

    NASA Technical Reports Server (NTRS)

    Paterson, R. W.

    1976-01-01

    A theoretical and experimental investigation of the noise and unsteady surface pressure characteristics of an isolated airfoil in a uniform mean velocity, homogeneous, nearly-isotropic turbulence field was conducted. Wind tunnel experiments were performed with a 23 cm chord, two dimensional NACA 0012 airfoil over a free stream Mach number range of 0.1 to 0.5. Far-field noise spectra and directivity were measured in an anechoic chamber that surrounded the tunnel open jet test section. Spanwise and chordwise distribution of unsteady airfoil surface pressure spectra and surface pressure cross-spectra were obtained. Incident turbulence intensities, length scales, spectra, and spanwise cross-spectra, required in the calculation of far-field noise and surface pressure characteristics were also measured.

  8. Low-speed aerodynamic characteristics of a 13.1-percent-thick, high-lift airfoil

    NASA Technical Reports Server (NTRS)

    Sivier, K. R.; Ormsbee, A. I.; Awker, R. W.

    1974-01-01

    Experimental study of the low-speed, sectional characteristics of a high-lift airfoil, and comparison of these characteristics with the predictions of the theoretical methods used in the airfoil's design. The 13.1% thick UI-1720 airfoil was found to achieve the predicted maximum lift coefficient of nearly 2.0. No upper-surface flow separation was found below the stall angle of attack of 16 deg; it appeared that stall was due to an abrupt leading-edge flow separation.

  9. Aerodynamic characteristics of an oscillating airfoil

    NASA Astrophysics Data System (ADS)

    Wickens, R. H.

    1986-03-01

    Results are reported from wind tunnel tests to study the effects of dynamic aerodynamics on the efficiency of a NACA 0018 airfoil used on a Darreius vertical axis wind turbine (VAWT). The topic is of interest because of uncontrolled pitching which occurs during operation and which produces stall, turbulence and separation effects that reduce efficiency. Present stream-tube theory and axial momentum models are not applicable in the unstable regimes. The wind tunnel tests were conducted with a 45 m/sec flow with an Re of 1.5 million. The situation mimicked typical wind turbine operational conditions. The airfoil was mounted on a hydraulic actuator to allow it to rotate about its quarter-chord location and to control the extent and frequency of oscillations. Data were also gathered on the performance in a steady flow for comparative purposes. Summary data are provided on the static and total pressures over a complete cycle of oscillation, and related to the angles of attack, time of onset of stall, and the lift and drag coefficients. The limitations of the study with regard to the absence of consideration of the flow acceleration experienced by an advancing blade are noted.

  10. Aerodynamic Characteristics of Twenty-Four Airfoils at High Speeds

    NASA Technical Reports Server (NTRS)

    Brigg, L J; Dryden, H L

    1930-01-01

    The aerodynamic characteristics of 24 airfoils are given for speeds of 0.5, 0.65, 0.8, 0.95, and 1.08 times the speed of sound, as measured in an open-jet air stream 2 inches in diameter, using models of 1-inch chord. The 24 airfoils belong to four general groups. The first is the standard R. A. F. family in general use by the Army and Navy for propeller design, the members of the family differing only in thickness. This family is represented by nine members ranging in thickness from 0.04 to 0.20 inch. The second group consists of five members of the Clark Y family, the members of the family again differing only in thickness. The third group, comprising six members, is a second R. A. F. Family in which the position of the maximum ordinate is varied. Combined with two members of the first R.A.F. family, this group represents a variation of maximum ordinate position from 30 to 60 percent of the chord in two camber ratios, 0.08 and 0.16. The fourth group consists of three geometrical forms, a flat plate, a wedge, and a segment of a right circular cylinder. In addition one section used in the reed metal propeller was included. These measurements form a part of a general program outlined at a Conference on Propeller Research organized by the National Advisory Committee for Aeronautics and the work was carried out with the financial assistance of the committee (author)

  11. Aerodynamic characteristics of three helicopter rotor airfoil sections at Reynolds number from model scale to full scale at Mach numbers from 0.35 to 0.90. [conducted in Langley 6 by 28 inch transonic tunnel

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

    An investigation was conducted in the Langely 6 by 28 inch transonic tunnel to determine the two dimensional aerodynamic characteristics of three helicopter rotor airfoils at Reynolds numbers from typical model scale to full scale at Mach numbers from about 0.35 to 0.90. The model scale Reynolds numbers ranged from about 700,00 to 1,500,000 and the full scale Reynolds numbers ranged from about 3,000,000 to 6,600,000. The airfoils tested were the NACA 0012 (0 deg Tab), the SC 1095 R8, and the SC 1095. Both the SC 1095 and the SC 1095 R8 airfoils had trailing edge tabs. The results of this investigation indicate that Reynolds number effects can be significant on the maximum normal force coefficient and all drag related parameters; namely, drag at zero normal force, maximum normal force drag ratio, and drag divergence Mach number. The increments in these parameters at a given Mach number owing to the model scale to full scale Reynolds number change are different for each of the airfoils.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Pressure distribution over an airfoil section with a flap and tab

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J

    1937-01-01

    Report presents the results of wind tunnel tests made in the NACA 7 by 10-foot wind tunnel of a Clark Y airfoil with a flap and an inset tab. The pressures were measured on both the upper and lower surfaces at one chord section. Calculations were made of the normal-force and pitching-moment coefficients of the airfoil section with flap section with tab, and the normal-force and hinge-moments coefficients of the tab alone. In addition, comparisons were made of the theoretical and experimental values for an airfoil with a multiply hinged flap system.

  17. Calculation of real-gas effects on airfoil aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Park, Chul; Yoon, Seokkwan

    1990-01-01

    The effects of high temperature thermochemical phenomena on the aerodynamic characteristics at hypersonic speeds are calculated for two-dimensional airfoils in air. The calculations are performed on an airfoil similar to that used for the Space Shuttle Orbiter, and ellipses of thickness ratios varying between 5 and 15 percent. For the airfoil, one flight condition is considered. For the ellipses, the calculations are carried out over a range of chord lengths, flight velocities, flight altitudes, and angles of attack. It is shown that the lift and drag coefficients are consistently reduced by the thermochemical phenomena, and that the behavior can be represented by a specific heat ratio value less than 1.4. The center of pressure shifts forward due to the thermochemical phenomena, but its extent is sensitively affected by the geometry and angle of attack and cannot be represented by a fixed specific heat ratio. The calculated results are in qualitative agreement with the data obtained during the entry flights of the Space Shuttle vehicle.

  18. 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.

  19. Increasing prototype airfoil fabrication efficiency through the use of sectional molds

    NASA Astrophysics Data System (ADS)

    Karges, Adam T.

    Airfoil development has always been important in the aeronautics industry. Current airfoil development techniques are being applied to design larger and more efficient wind turbine blades. To verify simulation results, a prototype blade must be built and tested. Current wing or blade structures are fabricated using traditional molding techniques. These large molds, particularly those used for wind turbine blades, can be fabricated from composite materials formed over a master shape. This process can be time and material intensive. This project develops techniques and methodology to build cavity molds using sectional pieces directly fabricated by computer numerically controlled (CNC) milling. A mold cavity was machined into tooling foam using CNC milling. This process allowed for mold creation without fabricating a master airfoil. Employment of several mold sections makes the machining process much easier and allows machine shops to produce larger, previously unfeasible, airfoil molds using limited machining length.

  20. 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.

  1. Detailed transonic flow field measurements about a supercritical airfoil section

    NASA Technical Reports Server (NTRS)

    Hurley, F. X.; Spaid, F. W.; Roos, F. W.; Stivers, L. S., Jr.; Bandettini, A.

    1975-01-01

    The transonic flow field about a Whitcomb-type supercritical airfoil profile was measured in detail. In addition to the usual surface pressure distributions and wake surveys, schlieren photographs were taken and velocity vector profiles were determined in the upper surface boundary layer and in the near wake. Spanwise variations in the measured pressures were also determined. The data are analyzed with the aid of an inviscid transonic finite-difference computer program as well as with boundary layer modeling and calculation schemes.

  2. 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

  3. Comparison of Pitching Moments Produced by Plain Flaps and by Spoilers and Some Aerodynamic Characteristics of an NACA 23012 Airfoil with Various Types of Aileron

    NASA Technical Reports Server (NTRS)

    Purser, Paul E.; Mckinney, Elizabeth G.

    1945-01-01

    Sectional characteristics of airfoil having retractable slotted flap with plain, slot-lip, or retractable ailerons are presented for a large range of aileron deflections. The analysis indicated that pitching moments produced by spoilers were less positive than those produced by plain flaps of equal effectiveness, also that pitching moments created by the spoiler increased less with the Mach number than similar moments produced by plain flaps. Positive values of pitching moment decreased as devices were located nearer airfoil leading edge.

  4. Investigation in the Langley 19-foot Pressure Tunnel of Two Wings of NACA 65-210 and 64-210 Airfoil Sections with Various Type Flaps

    NASA Technical Reports Server (NTRS)

    Sivells, James C; Spooner, Stanley H

    1949-01-01

    Report presents the results of an investigation conducted in the Langley 19-foot pressure tunnel to determine the maximum lift and stalling characteristics of two thin wings equipped with several types of flaps. Split, single slotted, and double slotted flaps were tested on one wing which had NACA 65-210 airfoil sections and split and double slotted flaps were tested on the other, which had NACA 64-210 airfoil sections. Both wings were zero sweep, an aspect ratio of 9, and a taper ratio of 0.4.

  5. Characteristics of the NACA 23012 Airfoil from Tests in the Full-Scale and Variable-Density Tunnels

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Clay, William C

    1936-01-01

    This report gives the results of tests in the NACA full-scale and variable-density tunnels of a new wing section, the NACA 23012, which is one of the more promising of an extended series of related airfoils recently developed. The tests were made at several values of the Reynolds number between 1,000,000 and 8,000,000. The new airfoil develops a reasonably high maximum lift and a low profile drag, which results in an unusually high value of the speed-range index. In addition, the pitching-moment coefficient is very small. The superiority of the new section over well-known and commonly used sections of small camber and moderate thickness is indicated by making a direct comparison with variable-density tests of the NACA 2212, the well-known NACA family airfoil that most nearly resembles it. The superiority is further indicated by comparing the characteristics with those obtained from full-scale-tunnel tests of the Clark y airfoil.

  6. 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.

  7. 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.

  8. Flow characteristics over NACA4412 airfoil at low Reynolds number

    NASA Astrophysics Data System (ADS)

    Genç, Mustafa Serdar; Koca, Kemal; Hakan Açıkel, Halil; Özkan, Gökhan; Sadık Kırış, Mehmet; Yıldız, Rahime

    2016-03-01

    In this study, the flow phenomena over NACA4412 were experimentally observed at various angle of attack and Reynolds number of 25000, 50000 and 75000, respectively. NACA4412 airfoil was manufactured at 3D printer and each tips of the wing were closed by using plexiglas to obtain two-dimensional airfoil. The experiments were conducted at low speed wind tunnel. The force measurement and hot-wire experiments were conducted to obtain data so that the flow phenomenon at the both top and bottom of the airfoil such as the flow separation and vortex shedding were observed. Also, smoke-wire experiment was carried out to visualize the surface flow pattern. After obtaining graphics from both force measurement experiment and hot-wire experiment compared with smoke wire experiment, it was noticed that there is a good coherence among the experiments. It was concluded that as Re number increased, the stall angle increased. And the separation bubble moved towards leading edge over the airfoil as the angle of attack increased.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. The variation with Reynolds number of pressure distribution over an airfoil section

    NASA Technical Reports Server (NTRS)

    Pinkerton, Robert M

    1938-01-01

    Pressures were simultaneously measured at 54 orifices distributed over the midspan section of a 5 by 30-inch rectangular model of the NACA 4412 airfoil in the variable-density tunnel. These measurements were made at 17 angles of attack from -20 degrees to 30 degrees for eight values of the effective Reynolds number form approximately 100,000 to 8,200,000. Accurate data were thus obtained for studying the variation of pressure distribution with Reynolds number. These results on the NACA 4412 section indicated that the pressure distribution is practically unaffected by changes in Reynolds number except where separation is involved.

  15. 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.

  16. Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils

    NASA Technical Reports Server (NTRS)

    Bousman, William G.

    2003-01-01

    Two airfoils are used on the main rotor blade of the UH-60A helicopter, the SC1095 and the SC1094 R8. Measurements of the section lift, drag, and pitching moment have been obtained in ten wind tunnel tests for the SC1095 airfoil, and in five of these tests, measurements have also been obtained for the SC1094 R8. The ten wind tunnel tests are characterized and described in the present study. A number of fundamental parameters measured in these tests are compared and an assessment is made of the adequacy of the test data for use in look-up tables required by lifting-line calculation methods.

  17. Desirable airfoil characteristics for large variable-speed horizontal axis wind turbines

    SciTech Connect

    Giguere, P.; Selig, M.S.

    1997-08-01

    In an effort to define the desirable airfoil characteristics for large variable-speed wind turbines, a systematic study was performed using a series of airfoils designed to have similar aerodynamic properties, except for the amount of lift, which varied over a wide range. For several airfoil combinations, blade shapes were designed for a 750-kW wind turbine with a 48.8-m diameter rotor using the optimization code PROPGA together with PROPID, which is an inverse design method for horizontal-axis wind turbines. Roughness effects, including the consideration of dirty-blade performance in the blade-shape optimization process, were also considered and are discussed. The results and conclusions reveal practical design implications that should aid in the aerodynamic blade design of not only large but also other sizes of variable-speed wind turbines.

  18. Wind Turbine Post-Stall Airfoil Performance Characteristics Guidelines for Blade-Element Momentum Methods: Preprint

    SciTech Connect

    Tangler, J. L.; Kocurek, J. D.

    2004-10-01

    The objective of this study was to provide post-stall airfoil data input guidelines for the prediction of peak and post-peak rotor power when using blade-element momentum theory. A steady-state data set from the Unsteady Aerodynamic Experiment (UAE) rotor test was used to provide guidelines for the development of a global post-stall method for the prediction of post-stall 3-D airfoil characteristics to be used with 2-D airfoil data. Based on these UAE data, methods to emulate the 3-D aerodynamics in the post-stall region were explored. Also suggested are experimental tests needed to better understand the 3-D flow physics and to quantify needed theory or empirical factors for a global post-stall approach to support blade-element momentum methods.

  19. 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

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Flow Observations with Tufts and Lampblack of the Stalling of Four Typical Airfoil Sections in the NACA Variable-density Tunnel

    NASA Technical Reports Server (NTRS)

    Abbott, Ira H; Sherman, Albert

    1938-01-01

    A preliminary investigation of the stalling processes of four typical airfoil sections was made over the critical range of the Reynolds Number. Motion pictures were taken of the movements of small silk tufts on the airfoil surface as the angle of attack increased through a range of angles including the stall. The boundary-layer flow also at certain angles of attack was indicated by the patterns formed by a suspension of lampblack in oil brushed onto the airfoil surface. These observations were analyzed together with corresponding force-test measurements to derive a picture of the stalling processes of airfoils.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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).

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Parametric study of separation and transition characteristics over an airfoil at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Boutilier, Michael S. H.; Yarusevych, Serhiy

    2012-06-01

    Time-resolved surface pressure measurements are used to experimentally investigate characteristics of separation and transition over a NACA 0018 airfoil for the relatively wide range of chord Reynolds numbers from 50,000 to 250,000 and angles of attack from 0° to 21°. The results provide a comprehensive data set of characteristic parameters for separated shear layer development and reveal important dependencies of these quantities on flow conditions. Mean surface pressure measurements are used to explore the variation in separation bubble position, edge velocity in the separated shear layer, and lift coefficients with angle of attack and Reynolds number. Consistent with previous studies, the separation bubble is found to move upstream and decrease in length as the Reynolds number and angle of attack increase. Above a certain angle of attack, the proximity of the separation bubble to the location of the suction peak results in a reduced lift slope compared to that observed at lower angles. Simultaneous measurements of the time-varying component of surface pressure at various spatial locations on the model are used to estimate the frequency of shear layer instability, maximum root-mean-square (RMS) surface pressure, spatial amplification rates of RMS surface pressure, and convection speeds of the pressure fluctuations in the separation bubble. A power-law correlation between the shear layer instability frequency and Reynolds number is shown to provide an order of magnitude estimate of the central frequency of disturbance amplification for various airfoil geometries at low Reynolds numbers. Maximum RMS surface pressures are found to agree with values measured in separation bubbles over geometries other than airfoils, when normalized by the dynamic pressure based on edge velocity. Spatial amplification rates in the separation bubble increase with both Reynolds number and angle of attack, causing the accompanying decrease in separation bubble length. Values of the

  17. 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.

  18. Method of making an airfoil

    NASA Technical Reports Server (NTRS)

    Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)

    1984-01-01

    An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.

  19. Method of making an airfoil

    NASA Technical Reports Server (NTRS)

    Moracz, Donald J. (Inventor); Cook, Charles R. (Inventor); Toth, Istvan J. (Inventor)

    1986-01-01

    An improved method of making an airfoil includes stacking plies in two groups. A separator ply is positioned between the two groups of plies. The groups of plies and the separator ply are interconnected to form an airfoil blank. The airfoil blank is shaped, by forging or other methods, to have a desired configuration. The material of the separator ply is then dissolved or otherwise removed from between the two sections of the airfoil blank to provide access to the interior of the airfoil blank. Material is removed from inner sides of the two separated sections to form core receiving cavities. After cores have been placed in the cavities, the two sections of the airfoil blank are interconnected and the shaping of the airfoil is completed. The cores are subsequently removed from the completed airfoil.

  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. Aerodynamics Characteristics of Multi-Element Airfoils at -90 Degrees Incidence

    NASA Technical Reports Server (NTRS)

    Stremel, Paul M.; Schmitz, Fredric H. (Technical Monitor)

    1994-01-01

    A developed method has been applied to calculate accurately the viscous flow about airfoils normal to the free-stream flow. This method has special application to the analysis of tilt rotor aircraft in the evaluation of download. In particular, the flow about an XV-15 airfoil with and without deflected leading and trailing edge flaps at -90 degrees incidence is evaluated. The multi-element aspect of the method provides for the evaluation of slotted flap configurations which may lead to decreased drag. The method solves for turbulent flow at flight Reynolds numbers. The flow about the XV-15 airfoil with and without flap deflections has been calculated and compared with experimental data at a Reynolds number of one million. The comparison between the calculated and measured pressure distributions are very good, thereby, verifying the method. The aerodynamic evaluation of multielement airfoils will be conducted to determine airfoil/flap configurations for reduced airfoil drag. Comparisons between the calculated lift, drag and pitching moment on the airfoil and the airfoil surface pressure will also be presented.

  2. 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.

  3. Experimental ice shape and performance characteristics for a multi-element airfoil in the NASA Lewis Icing Research Tunnel

    NASA Technical Reports Server (NTRS)

    Berkowitz, Brian M.; Potapczuk, Mark G.; Namdar, Bahman S.; Langhals, Tammy J.

    1991-01-01

    A study of the ice accretion patterns and performance of characteristics of a multi-element airfoil was undertaken at the NASA-Lewis Icing Research Tunnel. Several configurations were examined to determine the ice shape and performance characteristics. The testing included glaze, rime, and mixed icing regimes. Tunnel cloud conditions were set to correspond to those typical of the operating environment for commercial transport aircraft. Measurements acquired included ice profile tracings and aerodynamic forces both during the accretion process and in a post-accretion evaluation over a range of angle of attack. Substantial ice accretions developed on the main wing, flaps, and slat surfaces. Force measurements indicate severe performance degradation, especially near CL max, for both light and heavy ice accretion. Frost was seen on the lower surface of the airfoil which was found to contribute significantly to the force components.

  4. 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.

  5. 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.

  6. The Effect of Rivet Heads on the Characteristics of a 6 by 36 Foot Clark Y Metal Airfoil

    NASA Technical Reports Server (NTRS)

    Dearborn, Clinton H

    1933-01-01

    An investigation was conducted in the N.A.C.A. full-scale wind tunnel to determine the effects of exposed rivet heads on the aerodynamic characteristics of a metal-covered 6 by 36 foot Clasky airfoil. Lead punching simulating 1/8inch rivet heads were attached in full-span rows at a pitch of 1 inch at various chord positions. Tests were made at velocities varying from 40 to 120 miles per hour to investigate the scale effect. Rivets at the 5 percent chord position the upper surface of the airfoil produced the greatest increase in drag for a single row. Nine rows of rivets on both surfaces, simulating rivet spacing of multispan construction, increased the drag coefficients by a constant amount at velocities between 100 and 120 miles per hour. Accordingly, if rivets spaced the same as those on the test airfoil were used on a Clark Y wing of 300 square feet area and operated at 200 miles per hour, the drag would be increased over that for the smooth wing by 55 pounds and the power required would be increased by 29 horsepower.

  7. 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.

  8. The impact of unilateral vibrations on aerodynamic characteristics of airfoils in transonic flow

    NASA Astrophysics Data System (ADS)

    Zamuraev, V.; Kalinina, A.

    2016-06-01

    The work is devoted to the mathematical modeling of the influence of forced vibrations of a surface element on one side of the airfoil on the shock-wave structure of transonic flow around. The influence of parameters of oscillations on the airfoil wave drag and the lift force were qualitatively and quantitatively investigated for constant maximum velocity amplitude, which is close in magnitude to the sound velocity in the incoming flow, and for a wide range of frequencies. The arising of additional lift force is shown.

  9. Influence of unilateral oscillation on the aerodynamic characteristics of airfoils at transonic flow

    NASA Astrophysics Data System (ADS)

    Zamuraev, V. P.; Kalinina, A. P.

    2016-10-01

    The work is devoted to the mathematical modelling of the influence of forced vibrations of a surface element on one side of the airfoil on the shock-wave structure of transonic flow around. The influence of parameters of oscillations on the airfoil wave drag and the lift force were qualitatively and quantitatively investigated for constant maximum velocity amplitude, which is close in magnitude to the sound velocity in the oncoming flow, and for a wide range of frequencies. The additional lift force arising is shown.

  10. Airfoil structure

    SciTech Connect

    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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. Transonic Airfoil Development

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T.

    1983-01-01

    This lecture consists of three parts, in which discussions are presented of the current state of development of transonic or supercritical airfoils designed for fully turbulent boundary layers on the surfaces, previous research on subcritical airfoils designed to achieve laminar boundary layers on all or parts of the surfaces, and current research on supercritical airfoils designed to achieve laminar boundary layers. In the first part the use of available two dimensional computer codes in the development of supercritical airfoils and the general trends in the design of such airfoils with turbulent boundary layers are discussed. The second part provides the necessary background on laminar boundary layer phenomena. The last part, which constitutes the major portion of the lecture, covers research by NASA on supercritical airfoils utilizing both decreasing pressure gradients and surface suction for stabilizing the laminar boundary layer. An investigation of the former has been recently conducted in fight using gloves on the wing panels of the U.S. Air Force F111 TACT airplane, research on the later is currently being conducted in a transonic wind tunnel which has been modified to greatly reduce the stream turbulence and noise levels in the tests section.

  16. Nonlinear Behavior of a Typical Airfoil Section with Control Surface Freeplay: A Numerical and Experimental Study

    NASA Technical Reports Server (NTRS)

    Conner, M. D.; Tang, D. M.; Dowell, E. H.; Virgin, L. N.

    1997-01-01

    A three degree-of-freedom aeroelastic typical section with control surface freeplay is modeled theoretically as a system of piecewise linear state-space models. The system response is determined by time marching of the governing equations using a standard Runge-Kutta algorithm in conjunction with Henon's method for integrating a system of equations to a prescribed surface of phase space section. Henon's method is used to locate the "switching points" accurately and efficiently as the system moves from one linear region into another. An experimental model which closely approximates the three degree-of-freedom, typical section in two-dimensional, incompressible flow has been created to validate the theoretical model. Consideration is given to modeling realistically the structural damping present in the experimental system. The effect of the freeplay on the system response is examined numerically and experimentally. The development of the state-space model offers a low-order, computationally efficient means of modeling fully the freeplay nonlinearity and may offer advantages in future research which will investigate the effects of freeplay on the control of flutter in the typical section.

  17. 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.

  18. Experimental determination of the laminar separation bubble characteristics on an airfoil at low Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Omeara, M. M.; Mueller, T. J.

    1986-01-01

    An experimental investigation was conducted in order to document the structure and behavior of laminar separation bubbles at low Reynolds numbers. Data of this type is necessary if the currently insufficient analytical and numerical models are to be improved. The laminar separation bubble which forms on a NACA 66(3)-018 airfoil model was surveyed at chord Reynolds numbers ranging from 50,000 to 200,000 at angles of attack from 8 to 12 degrees. The effects of the various testing conditions on the separation bubble were isolated, and the data was analyzed in relation to existing separation bubble correlations in order to test their low Reynolds number applicability. This analysis indicated that the chord Reynolds number and the disturbance environment strongly influence the experimental pressure distributions. These effects must be included in any analytic prediction technique applied to the low Reynolds number flight regime.

  19. Transonic airfoil and axial flow rotary machine

    SciTech Connect

    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.

  20. The Surface Pressure Response of a NACA 0015 Airfoil Immersed in Grid Turbulence. Volume 1; Characteristics of the Turbulence

    NASA Technical Reports Server (NTRS)

    Bereketab, Semere; Wang, Hong-Wei; Mish, Patrick; Devenport, William J.

    2000-01-01

    Two grids have been developed for the Virginia Tech 6 ft x 6 ft Stability wind tunnel for the purpose of generating homogeneous isotropic turbulent flows for the study of unsteady airfoil response. The first, a square bi-planar grid with a 12" mesh size and an open area ratio of 69.4%, was mounted in the wind tunnel contraction. The second grid, a metal weave with a 1.2 in. mesh size and an open area ratio of 68.2% was mounted in the tunnel test section. Detailed statistical and spectral measurements of the turbulence generated by the two grids are presented for wind tunnel free stream speeds of 10, 20, 30 and 40 m/s. These measurements show the flows to be closely homogeneous and isotropic. Both grids produce flows with a turbulence intensity of about 4% at the location planned for the airfoil leading edge. Turbulence produced by the large grid has an integral scale of some 3.2 inches here. Turbulence produced by the small grid is an order of magnitude smaller. For wavenumbers below the upper limit of the inertial subrange, the spectra and correlations measured with both grids at all speeds can be represented using the von Karman interpolation formula with a single velocity and length scale. The spectra maybe accurately represented over the entire wavenumber range by a modification of the von Karman interpolation formula that includes the effects of dissipation. These models are most accurate at the higher speeds (30 and 40 m/s).

  1. 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

  2. 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)

  3. 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.

  4. The effect of undulating leading-edge modifications on NACA 0021 airfoil characteristics

    NASA Astrophysics Data System (ADS)

    Rostamzadeh, N.; Kelso, R. M.; Dally, B. B.; Hansen, K. L.

    2013-11-01

    In spite of its mammoth physical size, the humpback whale's manoeuvrability in hunting has captured the attention of biologists as well as fluid mechanists. It has now been established that the protrusions on the leading-edges of the humpback's pectoral flippers, known as tubercles, account for this species' agility and manoeuvrability. In the present work, Prandtl's nonlinear lifting-line theory was employed to propose a hypothesis that the favourable traits observed in the performance of tubercled lifting bodies are not exclusive to this form of leading-edge configuration. Accordingly, a novel alternative to tubercles was introduced and incorporated into the design of four airfoils that underwent wind tunnel force and pressure measurement tests in the transitional flow regime. In addition, a Computation Fluid Dynamics study was performed using the Shear Stress Transport transitional model in the context of unsteady Reynolds-Averaged Navier-Stokes at several attack angles. The results from the numerical investigation are in reasonable agreement with those of the experiments, and suggest the presence of features that are also observed in flows over tubercled foils, most notably a distinct pair of streamwise vortices for each wavelength of the tubercle-like feature.

  5. 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.

  6. Experimental studies of flow separation and stalling on a two-dimensional airfoil at low speeds

    NASA Technical Reports Server (NTRS)

    Seetharam, H. C.; Wentz, W. H., Jr.

    1975-01-01

    Detailed measurements of flow fields associated with low-speed turbulent boundary layers were made for the 17% thick GA(W)-1 airfoil section at nominal angles of attack of 10 deg, 14 deg, and 18 deg, Reynolds number 2.2 x 10, and Mach number 0.13. The data include pressure and velocity surveys of the pre- and post-separated regions on the airfoil and the associated wake. The boundary layer characteristics including regions of separation on the airfoil are also presented.

  7. 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.

  8. 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.

  9. Airfoil profile in a nonuniform flow

    NASA Technical Reports Server (NTRS)

    Polasek, J.

    1978-01-01

    A theory of airfoil section past two dimensional nonuniform flow is developed. The theory is based on representation of airfoil section by vortex and source distributions and it can be used for calculation of aircraft wings in homogeneous and inhomogeneous flow, as well as for calculation of straight and radial blade and vane-cascades.

  10. Preliminary Investigation on Boundary Layer Control by Means of Suction and Pressure with the U.S.A. 27 Airfoil

    NASA Technical Reports Server (NTRS)

    Reid, E G; Bamber, M J

    1928-01-01

    The tests described in this report constitute a preliminary investigation of airfoil boundary layer control, as carried out in the atmospheric wind tunnel of the Langley Memorial Aeronautical Laboratory, from February to August, 1927. Tests were made on a U.S.A. 27 airfoil section with various slot shapes and combinations, and at various amounts of pressure or suction on the slots. The lift of airfoils can be increased by removing or by accelerating the boundary layer. Removing the boundary layer by suction is more economical than to accelerate it by jet action. Gauze-covered suction slots apparently give the best results. When not in operation, all suction slots tested had a detrimental effect upon the aerodynamic characteristics of the airfoil which was not apparent with the backward-opening pressure slots. Thick, blunt-nose airfoils would seem to give best results with boundary layer control.

  11. Hybrid airfoil design methods for full-scale ice accretion simulation

    NASA Astrophysics Data System (ADS)

    Saeed, Farooq

    The objective of this thesis is to develop a design method together with a design philosophy that allows the design of "subscale" or "hybrid" airfoils that simulate fullscale ice accretions. These subscale or hybrid airfoils have full-scale leading edges and redesigned aft-sections. A preliminary study to help develop a design philosophy for the design of hybrid airfoils showed that hybrid airfoils could be designed to simulate full-scale airfoil droplet-impingement characteristics and, therefore, ice accretion. The study showed that the primary objective in such a design should be to determine the aft section profile that provides the circulation necessary for simulating full-scale airfoil droplet-impingement characteristics. The outcome of the study, therefore, reveals circulation control as the main design variable. To best utilize this fact, this thesis describes two innovative airfoil design methods for the design of hybrid airfoils. Of the two design methods, one uses a conventional flap system while the other only suggests the use of boundary-layer control through slot-suction on the airfoil upper surface as a possible alternative for circulation control. The formulation of each of the two design methods is described in detail, and the results from each method are validated using wind-tunnel test data. The thesis demonstrates the capabilities of each method with the help of specific design examples highlighting their application potential. In particular, the flap-system based hybrid airfoil design method is used to demonstrate the design of a half-scale hybrid model of a full-scale airfoil that simulates full-scale ice accretion at both the design and off-design conditions. The full-scale airfoil used is representative of a scaled modern business-jet main wing section. The study suggests some useful advantages of using hybrid airfoils as opposed to full-scale airfoils for a better understanding of the ice accretion process and the related issues. Results

  12. 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.

  13. Composite airfoil assembly

    SciTech Connect

    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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Charts for Determining the Characteristics of Sharp-nose Airfoils in Two-dimensional Flow at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Ivey, H Reese; Stickle, George W; Schuettler, Alberta

    1947-01-01

    Solutions of the Hugoniot shock equations and Meyer expansion equations are plotted in such a manner as to permit the pressure distribution, the local Mach number, and the angles of shock waves on arbitrary sharp-nose airfoils at supersonic speeds to be obtained directly. (author)

  19. 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.

  20. 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.

  1. Wind-Tunnel Investigation of an NACA 23021 Airfoil with a 0.32-Airfoil-Chord Double Slotted Flap

    NASA Technical Reports Server (NTRS)

    Fischel, Jack; Riebe, John M

    1944-01-01

    An investigation was made in the LMAL 7- by 10-foot wind tunnel of a NACA 23021 airfoil with a double slotted flap having a chord 32 percent of the airfoil chord (0.32c) to determine the aerodynamic section characteristics with the flaps deflected at various positions. The effects of moving the fore flap and rear flap as a unit and of deflecting or removing the lower lip of the slot were also determined. Three positions were selected for the fore flap and at each position the maximum lift of the airfoil was obtained with the rear flap at the maximum deflection used at that fore-flap position. The section lift of the airfoil increased as the fore flap was extended and maximum lift was obtained with the fore flap deflected 30 deg in the most extended position. This arrangement provided a maximum section lift coefficient of 3.31, which was higher than the value obtained with either a 0.2566c or a 0.40c single-slotted-flap arrangement and 0.25 less than the value obtained with a 0.4c double-slotted-flap arrangement on the same airfoil. The values of the profile-drag coefficient obtained with the 0.32c double slotted flap were larger than those for the 0.2566c or 0.40c single slotted flaps for section lift coefficients between 1.0 and approximately 2.7. At all values of the section lift coefficient above 1.0, the 0.40c double slotted flap had a lower profile drag than the 0.32c double slotted flap. At various values of the maximum section lift coefficient produced by various flap defections, the 0.32c double slotted flap gave negative section pitching-moment coefficients that were higher than those of other slotted flaps on the same airfoil. The 0.32c double slotted flap gave approximately the same maximum section lift coefficient as, but higher profile-drag coefficients over the entire lift range than, a similar arrangement of a 0.30c double slotted flap on an NACA 23012 airfoil.

  2. An airfoil parameterization method for the representation and optimization of wind turbine special airfoil

    NASA Astrophysics Data System (ADS)

    Liu, Yixiong; Yang, Ce; Song, Xiancheng

    2015-04-01

    A new airfoil shape parameterization method is developed, which extended the Bezier curve to the generalized form with adjustable shape parameters. The local control parameters at airfoil leading and trailing edge regions are enhanced, where have significant effect on the aerodynamic performance of wind turbine. The results show this improved parameterization method has advantages in the fitting characteristics of geometry shape and aerodynamic performance comparing with other three common airfoil parameterization methods. The new parameterization method is then applied to airfoil shape optimization for wind turbine using Genetic Algorithm (GA), and the wind turbine special airfoil, DU93-W-210, is optimized to achieve the favorable Cl/Cd at specified flow conditions. The aerodynamic characteristic of the optimum airfoil is obtained by solving the RANS equations in computational fluid dynamics (CFD) method, and the optimization convergence curves show that the new parameterization method has good convergence rate in less number of generations comparing with other methods. It is concluded that the new method not only has well controllability and completeness in airfoil shape representation and provides more flexibility in expressing the airfoil geometry shape, but also is capable to find efficient and optimal wind turbine airfoil. Additionally, it is shown that a suitable parameterization method is helpful for improving the convergence rate of the optimization algorithm.

  3. 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.

  4. Darrieus wind-turbine airfoil configurations

    SciTech Connect

    Migliore, P.G.; Fritschen, J.R.

    1982-06-01

    The purpose of this study was to determine what aerodynamic performance improvement, if any, could be achieved by judiciously choosing the airfoil sections for Darrieus wind turbine blades. Analysis was limited to machines using two blades of infinite aspect ratio, having rotor solidites from seven to twenty-one percent, and operating at maximum Reynolds numbers of approximately three million. Ten different airfoils, having thickness to chord ratios of twelve, fifteen and eighteen percent, were investigated. Performance calculations indicated that the NACA 6-series airfoils yield peak power coefficients at least as great as the NACA four-digit airfoils which have historically been chosen for Darrieus turbines. Furthermore, the power coefficient-tip speed ratio curves were broader and flatter for the 6-series airfoils. Sample calculations for an NACA 63/sub 2/-015 airfoil showed an annual energy output increase of 17 to 27% depending upon rotor solidity, compared to an NACA 0015 airfoil. An attempt was made to account for the flow curvature effects associated with Darrieus turbines by transforming the NACA 63/sub 2/-015 airfoil to an appropriate shape.

  5. 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.

  6. 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.

  7. Wind-Tunnel Investigation of an NACA 23012 Airfoil with Various Arrangements of Slotted Flaps

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Harris , Thomas A

    1939-01-01

    An investigation was made in the 7 by 10-foot wind tunnel and in the variable-density wind tunnel of the NACA 23012 airfoil with various slotted-flap arrangements. The purpose of the investigation in the 7 by 10-foot wind tunnel was to determine the airfoil section aerodynamic characteristics as affected by flap shape, slot shape, and flap location. The flap position for maximum lift; polars for arrangements favorable for take-off and climb; and complete lift, drag, and pitching-moment characteristics for selected optimum arrangements were determined. The best arrangements were tested in the variable-density tunnel at an effective Reynolds number of 8,000,000. In addition, data from both wind tunnels are included for plain, split, external-airfoil, and Fowler flaps for purposes of comparison.

  8. Preparing and Analyzing Iced Airfoils

    NASA Technical Reports Server (NTRS)

    Vickerman, Mary B.; Baez, Marivell; Braun, Donald C.; Cotton, Barbara J.; Choo, Yung K.; Coroneos, Rula M.; Pennline, James A.; Hackenberg, Anthony W.; Schilling, Herbert W.; Slater, John W.; Burke, Kevin M.; Nolan, Gerald J.; Brown, Dennis

    2004-01-01

    SmaggIce version 1.2 is a computer program for preparing and analyzing iced airfoils. It includes interactive tools for (1) measuring ice-shape characteristics, (2) controlled smoothing of ice shapes, (3) curve discretization, (4) generation of artificial ice shapes, and (5) detection and correction of input errors. Measurements of ice shapes are essential for establishing relationships between characteristics of ice and effects of ice on airfoil performance. The shape-smoothing tool helps prepare ice shapes for use with already available grid-generation and computational-fluid-dynamics software for studying the aerodynamic effects of smoothed ice on airfoils. The artificial ice-shape generation tool supports parametric studies since ice-shape parameters can easily be controlled with the artificial ice. In such studies, artificial shapes generated by this program can supplement simulated ice obtained from icing research tunnels and real ice obtained from flight test under icing weather condition. SmaggIce also automatically detects geometry errors such as tangles or duplicate points in the boundary which may be introduced by digitization and provides tools to correct these. By use of interactive tools included in SmaggIce version 1.2, one can easily characterize ice shapes and prepare iced airfoils for grid generation and flow simulations.

  9. 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.

  10. 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

  11. Performance predictions of VAWTs with NLF airfoil blades

    SciTech Connect

    Masson, C.; Leclerc, C.; Paraschivoiu, I.

    1997-02-01

    The successful design of an efficient Vertical Axis Wind Turbine (VAWT) can be obtained only when appropriate airfoil sections have been selected. Most VAWTs currently operating worldwide use blades of symmetrical NACA airfoil series. As these blades were designed for aviation applications, Sandia National Laboratories developed a family of airfoils specifically designed for VAWTs in order to decrease the Cost of Energy (COE) of the VAWT (Berg, 1990). Objectives formulated for the blade profile were: modest values of maximum lift coefficient, low drag at low angle of attack, high drag at high angle of attack, sharp stall, and low thickness-to-chord ratio. These features are similar to those of Natural Laminar Flow airfoils (NLF) and gave birth to the SNLA airfoil series. This technical brief illustrates the benefits and losses resulting from using NLF airfoils on VAWT blades. To achieve this goal, the streamtube model of Paraschivoiu (1988) is used to predict the performance of VAWTs equipped with blades of various airfoil shapes. The airfoil shapes considered are the conventional airfoils NACA 0018 and NACA 0021, and the SNLA 0018/50 airfoil designed at Sandia. Furthermore, the potential benefit of reducing the airfoil drag is clearly illustrated by the presentation of the individual contributions of lift and drag to power.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Computation of airfoil buffet boundaries

    NASA Technical Reports Server (NTRS)

    Levy, L. L., Jr.; Bailey, H. E.

    1981-01-01

    The ILLIAC IV computer has been programmed with an implicit, finite-difference code for solving the thin layer compressible Navier-Stokes equation. Results presented for the case of the buffet boundaries of a conventional and a supercritical airfoil section at high Reynolds numbers are found to be in agreement with experimentally determined buffet boundaries, especially at the higher freestream Mach numbers and lower lift coefficients where the onset of unsteady flows is associated with shock wave-induced boundary layer separation.

  1. The effect of a cavity on airfoil tones

    NASA Astrophysics Data System (ADS)

    Schumacher, Karn L.; Doolan, Con J.; Kelso, Richard M.

    2014-03-01

    The presence of a cavity in the pressure surface of an airfoil has been found via experiment to play a role in the production of airfoil tones, which was attributed to the presence of an acoustic feedback loop. The cavity length was sufficiently small that cavity oscillation modes did not occur for most of the investigated chord-based Reynolds number range of 70,000-320,000. The airfoil tonal noise frequencies varied as the position of the cavity was moved along a parallel section at the airfoil's maximum thickness: specifically, for a given velocity, the frequency spacing of the tones was inversely proportional to the geometric distance between the cavity and the trailing edge. The boundary layer instability waves considered responsible for the airfoil tones were only detected downstream of the cavity. This may be the first experimental verification of these aspects of the feedback loop model for airfoil tonal noise.

  2. 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.

  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. Shapes for rotating airfoils

    NASA Technical Reports Server (NTRS)

    Bingham, G. J. (Inventor)

    1984-01-01

    An airfoil which has particular application to the blade or blades of rotor aircraft and aircraft propellers is presented. The airfoil thickness distribution, camber and leading edge radius are shaped to locate the airfoil crest at a more aft position along the chord, and to increase the freestream Mach number at which sonic flow is attained at the airfoil crest. The reduced slope of the airfoil causes a reduction in velocity at the airfoil crest at lift coefficients from zero to the maximum lift coefficient. The leading edge radius is adjusted so that the maximum local Mach number at 1.25 percent chord and at the designed maximum lift coefficient is limited to about 0.48 when the Mach number normal to the leading edge is approximately 0.20. The lower surface leading edge radius is shaped so that the maximum local Mach number at the leading edge is limited to about 0.29 when the Mach number normal to the leading edge is approximately 0.20. The drag divergence Mach number associated with the airfoil is moved to a higher Mach number over a range of lift coefficients resulting in superior aircraft performance.

  6. 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.

  7. 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.

  8. Internal-external flow integration for a thin ejector-flapped wing section

    NASA Technical Reports Server (NTRS)

    Woolard, H. W.

    1979-01-01

    Thin airfoil theories of an ejector flapped wing section are reviewed. The global matching of the external airfoil flow with the ejector internal flow and the overall ejector flapped wing section aerodynamic performance are examined. Mathematical models of the external and internal flows are presented. The delineation of the suction flow coefficient characteristics are discussed. The idealized lift performance of an ejector flapped wing relative to a jet augmented flapped wing are compared.

  9. Summary of Drag Characteristics of Practical-Construction Wing Sections

    NASA Technical Reports Server (NTRS)

    Quinn, John H , Jr

    1948-01-01

    The effect of several parameters on the drag characteristics of practical-construction wing sections have been considered and evaluated. The effects considered were those of surface roughness, surface waviness, compressive load, and de-icers. The data were obtained from a number of tests in the Langley two-dimensional low-turbulence tunnels.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Turbine airfoil with controlled area cooling arrangement

    DOEpatents

    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.

  3. 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

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. Oscillatory blowing airfoil

    NASA Technical Reports Server (NTRS)

    1997-01-01

    10' NACA 0015 with 30% chord trailing edge flap deflected 20 degrees. Used in 0.3 Meter Transonic Cryogenic Tunnel, this airfoil has a 0.44 mm slot at 70% chord. Oscillatory blowing out of slot used for separation control. Howard Price appears in side view shot, in building 1145, Studio.

  9. Turbine airfoil manufacturing technology

    SciTech Connect

    Kortovich, C.

    1995-12-31

    The specific goal of this program is to define manufacturing methods that will allow single crystal technology to be applied to complex-cored airfoils components for power generation applications. Tasks addressed include: alloy melt practice to reduce the sulfur content; improvement of casting process; core materials design; and grain orientation control.

  10. Airfoil Tonal Noise Generation in Resonant Environments

    NASA Astrophysics Data System (ADS)

    Atobe, Takashi; Tuinstra, Marthijn; Takagi, Shohei

    To clarify tonal noise generation, an experimental study on airfoil tonal noise was undertaken using a conventional wind tunnel, which allows acoustic reflection on test section walls. A two-dimensional wing model with the NACA0015 cross-section was used at 5 degrees angle of attack. Most previous experiments conducted in anechoic environments commonly show that the tonal noise frequency is selected in an overall trend of U1.5 (U is uniform velocity) locally consisting of a step-like structure, and Tollmien-Schlichting disturbances are rapidly amplified in the backflow region near the trailing edge of the pressure surface. The present experiments in an acoustically resonant environment show that the tonal noise emanates in accordance with the aforementioned features. However, the ladder-like structure has a different local slope from that observed in anechoic flow. These characteristics suggest that acoustic resonance does not play a fundamental role in tonal noise generation. Observation by hot-wire and smoke visualization techniques shows that unsteady disturbances rather than Tollmien-Schlichting waves are rapidly magnified by the Kelvin-Helmholtz instability in the backflow region. The frequency selection mechanism at tonal noise generation still remains unsolved.

  11. 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.

  12. 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.

  13. 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.

  14. On bimodal flutter behavior of a flexible airfoil

    NASA Astrophysics Data System (ADS)

    Drazumeric, Radovan; Gjerek, Bojan; Kosel, Franc; Marzocca, Pier

    2014-02-01

    The dynamic aeroelastic behavior of an elastically supported airfoil is studied in order to investigate the possibilities of increasing critical flutter speed by exploiting its chord-wise flexibility. The flexible airfoil concept is implemented using a rigid airfoil-shaped leading edge, and a flexible thin laminated composite plate conformally attached to its trailing edge. The flutter behavior is studied in terms of the number of laminate plies used in the composite plate for a given aeroelastic system configuration. The flutter behavior is predicted by using an eigenfunction expansion approach which is also used to design a laminated plate in order to attain superior flutter characteristics. Such an airfoil is characterized by two types of flutter responses, the classical airfoil flutter and the plate flutter. Analysis shows that a significant increase in the critical flutter speed can be achieved with high plunge and low pitch stiffness in the region where the aeroelastic system exhibits a bimodal flutter behavior, e.g., where the airfoil flutter and the plate flutter occur simultaneously. The predicted flutter behavior of a flexible airfoil is experimentally verified by conducting a series of systematic aeroelastic system configurations wind tunnel flutter campaigns. The experimental investigations provide, for each type of flutter, a measured flutter response, including the one with indicated bimodal behavior.

  15. Buffeting of NACA 0012 airfoil at high angle of attack

    NASA Astrophysics Data System (ADS)

    Zhou, Tong; Dowell, Earl

    2014-11-01

    Buffeting is a fluid instability caused by flow separation or shock wave oscillations in the flow around a bluff body. Typically there is a dominant frequency of these flow oscillations called Strouhal or buffeting frequency. In prior work several researchers at Duke University have noted the analogy between the classic Von Karman Vortex Street behind a bluff body and the flow oscillations that occur for flow around a NACA 0012 airfoil at sufficiently large angle of attack. Lock-in is found for certain combinations of airfoil oscillation (pitching motion) frequencies and amplitudes when the frequency of the airfoil motion is sufficiently close to the buffeting frequency. The goal of this paper is to explore the flow around a static and an oscillating airfoil at high angle of attack by developing a method for computing buffet response. Simulation results are compared with experimental data. Conditions for the onset of buffeting and lock-in of a NACA 0012 airfoil at high angle of attack are determined. Effects of several parameters on lift coefficient and flow response frequency are studied including Reynolds number, angle of attack and blockage ratio of the airfoil size to the wind tunnel dimensions. Also more detailed flow field characteristics are determined. For a static airfoil, a universal Strouhal number scaling has been found for angles of attack from 30° to 90°, where the flow around airfoil is fully separated. For an oscillating airfoil, conditions for lock-in are discussed. Differences between the lock-in case and the unlocked case are also studied. The second affiliation: Duke University.

  16. 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.

  17. The Role of Separation Bubbles on the Aerodynamic Characteristics of Airfoils, Including Stall and Post-Stall, at Low Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Chen, Hsun H.; Cebeci, Tuncer

    2007-01-01

    Airfoils at high Reynolds numbers, in general, have small separation bubbles that are usually confined to the leading edge. Since the Reynolds number is large, the turbulence model for the transition region between the laminar and turbulent flow is not important. Furthermore, the onset of transition occurs either at separation or prior to separation and can be predicted satisfactorily by empirical correlations when the incident angle is small and can be assumed to correspond to laminar separation when the correlations do not apply, i.e., at high incidence angles.

  18. 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.

  19. 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.

  20. Steady and Unsteady Aerodynamics of Thin Airfoils with Porosity Gradients

    NASA Astrophysics Data System (ADS)

    Hajian, Rozhin; Jaworski, Justin W.

    2015-11-01

    Porous treatments have been shown in previous studies to reduce turbulence noise generation from the edges of wings and blades. However, this acoustical benefit can come at the cost of aerodynamic performance that is degraded by seepage flow through the wing. To better understand the trade-off between acoustic stealth and the desired airfoil performance, the aerodynamic loads of a thin airfoil in uniform flow with a prescribed porosity distribution are determined analytically in closed form, provided that the distribution is Hölder-continuous. The theoretical model is extended to include unsteady heaving and pitching motions of the airfoil section, which has applications to the performance estimation of biologically-inspired swimmers and fliers and to the future assessment of vortex noise production from porous airfoils.

  1. 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.

  2. Tracking of raindrops in flow over an airfoil

    NASA Technical Reports Server (NTRS)

    Valentine, James R.; Decker, Rand

    1993-01-01

    The splashback that occurs when raindrops impact an airfoil results in an 'ejecta fog' of small droplets around the leading edge. Acceleration of these droplets by the air flow field is a momentum sink for the air flow and has been hypothesized to contribute to the degradation of airfoil performance in heavy rain. Presented here is a one-way coupled Eulerian-Lagrangian particle tracking scheme to evaluate droplet number densities and momentum sink terms around a NACA 64-210 airfoil section for three rainfall rates. A laminar air flow field is determined with a standard CFD code and input to the particle tracking algorithm. Raindrops are assumed to be non-interacting, non-deforming, non-evaporating, and non-spinning spheres and are tracked through the curvilinear grid used by the air flow code. A simple model is used to simulate impacts and the resulting splashback on the airfoil surface.

  3. Simulation of flow over double-element airfoil and wind tunnel test for use in vertical axis wind turbine

    NASA Astrophysics Data System (ADS)

    Chougule, Prasad; Nielsen, Søren R. K.

    2014-06-01

    Nowadays, small vertical axis wind turbines are receiving more attention due to their suitability in micro-electricity generation. There are few vertical axis wind turbine designs with good power curve. However, the efficiency of power extraction has not been improved. Therefore, an attempt has been made to utilize high lift technology for vertical axis wind turbines in order to improve power efficiency. High lift is obtained by double-element airfoil mainly used in aeroplane wing design. In this current work a low Reynolds number airfoil is selected to design a double-element airfoil blade for use in vertical axis wind turbine to improve the power efficiency. Double-element airfoil blade design consists of a main airfoil and a slat airfoil. Orientation of slat airfoil is a parameter of investigation in this paper and air flow simulation over double-element airfoil. With primary wind tunnel test an orientation parameter for the slat airfoil is initially obtained. Further a computational fluid dynamics (CFD) has been used to obtain the aerodynamic characteristics of double-element airfoil. The CFD simulations were carried out using ANSYS CFX software. It is observed that there is an increase in the lift coefficient by 26% for single-element airfoil at analysed conditions. The CFD simulation results were validated with wind tunnel tests. It is also observe that by selecting proper airfoil configuration and blade sizes an increase in lift coefficient can further be achieved.

  4. 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.

  5. NREL airfoil families for HAWTs

    NASA Astrophysics Data System (ADS)

    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.

  6. 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.

  7. Subsonic and transonic low-Reynolds-number airfoils with reduced pitching moments

    NASA Technical Reports Server (NTRS)

    Van Dam, C. P.; Hicks, R.; Reuther, J.

    1990-01-01

    A subsonic and a transonic airfoil are presented for application in a high-altitude long-endurance aircraft and a very-high-altitude aircraft, respectively. The subsonic airfoil is designed for a lift coefficient c(l) = 1.4 at a chord Reynolds number Re = 700,000 and a very low Mach number. The transonic airfoil is designed for c(l) = 1.0 at Re = 500,000 and a transonic Mach number M = 0.7. Both airfoils are developed to perform as well or better than previously designed airfoils. However, the present airfoils are developed for a constrained pitching moment to reduce aircraft trim drag and to relieve, to some extent, the torsional loads in the typically high-aspect-ratio wings. The beneficial effects of a cruise flap and of boundary-layer transition control on the off-design performance characteristics are illustrated.

  8. Synthesized airfoil data method for prediction of dynamic stall and unsteady airloads

    NASA Technical Reports Server (NTRS)

    Gangwani, S. T.

    1983-01-01

    A detailed analysis of dynamic stall experiments has led to a set of relatively compact analytical expressions, called synthesized unsteady airfoil data, which accurately describe in the time-domain the unsteady aerodynamic characteristics of stalled airfoils. An analytical research program was conducted to expand and improve this synthesized unsteady airfoil data method using additional available sets of unsteady airfoil data. The primary objectives were to reduce these data to synthesized form for use in rotor airload prediction analyses and to generalize the results. Unsteady drag data were synthesized which provided the basis for successful expansion of the formulation to include computation of the unsteady pressure drag of airfoils and rotor blades. Also, an improved prediction model for airfoil flow reattachment was incorporated in the method. Application of this improved unsteady aerodynamics model has resulted in an improved correlation between analytic predictions and measured full scale helicopter blade loads and stress data.

  9. Impingement of Water Droplets on NACA 65A004 Airfoil at 8 deg Angle of Attack

    NASA Technical Reports Server (NTRS)

    Brun, R. J.; Gallagher, H. M.; Vogt, D. E.

    1954-01-01

    The trajectories of droplets in the air flowing past an NACA 65AO04 airfoil at an angle of attack of 8 deg were determined.. The amount of water in droplet form impinging on the airfoil, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface were calculated from the trajectories and presented to cover a large range of flight and atmospheric conditions. These impingement characteristics are compared briefly with those previously reported for the same airfoil at an angle of attack of 4 deg.

  10. 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.

  11. 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.

  12. Aerodynamic Flow Control of a Maneuvering Airfoil

    NASA Astrophysics Data System (ADS)

    Brzozowski, Daniel P.; Culp, John; Glezer, Ari

    2010-11-01

    The unsteady aerodynamic forces and moments on a maneuvering, free-moving airfoil are varied in wind tunnel experiments by controlling vorticity generation/accumulation near the surface using hybrid synthetic jet actuators. The dynamic characteristics of the airfoil that is mounted on a 2-DOF traverse are controlled using position and attitude feedback loops that are actuated by servo motors. Bi-directional changes in the pitching moment are induced using controllable trapped vorticity concentrations on the suction and pressure surfaces near the trailing edge. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The time scales associated with the actuation process is determined from PIV measurements of vorticity flux downstream of the trailing edge. Circulation time history shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales commensurate with the flow's convective time scale, and that the maneuver response is only limited by the inertia of the platform. Supported by AFSOR.

  13. 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.

  14. 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.

  15. Flow visualization around a double wedge airfoil model with focusing schlieren system

    NASA Astrophysics Data System (ADS)

    Kashitani, Masashi; Yamaguchi, Yutaka

    2006-03-01

    In the present study, aerodynamic characteristics of the double wedge airfoil model were investigated in a transonic flow by using the shock tube as an intermittent wind tunnel. The driver and driven gases of the shock tube are dry air. The airfoil model of double wedge has the span of 58 mm, chord length c = 75 mm and its maximum thickness is 7.5 mm. The apex of the double wedge airfoil model is located on the 35% chord length from the leading edge. The range of hot gas Mach numbers are from 0.80 to 0.88, and the Reynolds numbers based on chord length are 3.11 × 105 ˜ 3.49 × 105, respectively. The flow visualizations were performed by the sharp focusing schlieren method which can visualize the three dimensional flow fields. The results show that the present system can visualize the transonic flowfield clearer than the previous system, and the shock wave profiles of the center of span in the test section are visualized

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Tests of the NACA 0025 and 0035 Airfoils in the Full-Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Bullivant, W Kenneth

    1941-01-01

    This report presents the results of an investigation conducted in the NACA full-scale wind tunnel to determine the aerodynamic characteristics of the 6 by 36-foot rectangular NACA 0025 and 0035 airfoils. The aerodynamic characteristics of the plain airfoils with rounded and square tips were determined by force tests through a complete angle-of-attack range, in addition, the profile drag was determined by the momentum method. The transition points on the airfoils were located by boundary-layer determinations with small total-head and static tubes. Each airfoil was also tested with a 0.20c full-span split flap. Tuft surveys were included to show the progressive breakdown of flow with increasing angles of attack. Previously published data from tests of the NACA 0009, 0012, and 0018 airfoils in the full-scale tunnel have been included in the summary curves.

  1. 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.

  2. 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.

  3. 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

  4. Wind-Tunnel Investigation of Wings with Ordinary Ailerons and Full-Span External-Airfoil Flaps

    NASA Technical Reports Server (NTRS)

    Platt, Robert C; Shortal, Joseph A

    1937-01-01

    Report presents an investigation carried out in the NACA 7- by 10-foot wind tunnel of an NACA 23012 airfoil equipped, first, with a full-span NACA 23012 external-airfoil flap having a chord 0.20 of the main airfoil chord and with a full-span aileron with a chord 0.12 of the main airfoil chord on the trailing edge of the main airfoil and equipped second, with a 0.30-chord full-span NACA 23012 external-airfoil flap and a 0.13-chord full-span aileron. The results are arranged in three groups, the first two of which deal with the airfoil characteristics of the two airfoil-flap combinations and with the internal-control characteristics of the airfoil-flap-aileron combinations. The third group of tests deals with several means for balancing ailerons mounted on a special large-chord NACA 23012 external-airfoil flap. The tests included an ordinary aileron, a curtained-nose balance, a frise balance, and a tab.

  5. 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.

  6. 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

  7. High Reynolds number transonic tests on a NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.; Hill, S. Acquilla

    1987-01-01

    Tests were conducted in the two-dimensional test section of the Langley 0.3-m Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Plots of the spanwise variation of drag coefficient as a function of normal force coefficient and the variation of the basic aerodynamic characteristics with angle of attack are shown. The data are presented uncorrected for wall interference effects and without analysis.

  8. Recent work on airfoil theory

    NASA Technical Reports Server (NTRS)

    Prandtl, L

    1940-01-01

    The basic ideas of a new method for treating the problem of the airfoil are presented, and a review is given of the problems thus far computed for incompressible and supersonic flows. Test results are reported for the airfoil of circular plan form and the results are shown to agree well with the theory. As a supplement, a theory based on the older methods is presented for the rectangular of small aspect ratio.

  9. Reduction of airfoil trailing edge noise by trailing edge blowing

    NASA Astrophysics Data System (ADS)

    Gerhard, T.; Erbslöh, S.; Carolus, T.

    2014-06-01

    The paper deals with airfoil trailing edge noise and its reduction by trailing edge blowing. A Somers S834 airfoil section which originally was designed for small wind turbines is investigated. To mimic realistic Reynolds numbers the boundary layer is tripped on pressure and suction side. The chordwise position of the blowing slot is varied. The acoustic sources, i.e. the unsteady flow quantities in the turbulent boundary layer in the vicinity of the trailing edge, are quantified for the airfoil without and with trailing edge blowing by means of a large eddy simulation and complementary measurements. Eventually the far field airfoil noise is measured by a two-microphone filtering and correlation and a 40 microphone array technique. Both, LES-prediction and measurements showed that a suitable blowing jet on the airfoil suction side is able to reduce significantly the turbulence intensity and the induced surface pressure fluctuations in the trailing edge region. As a consequence, trailing edge noise associated with a spectral hump around 500 Hz could be reduced by 3 dB. For that a jet velocity of 50% of the free field velocity was sufficient. The most favourable slot position was at 90% chord length.

  10. Development of a Shape-controlled airfoil by use of SMA

    NASA Astrophysics Data System (ADS)

    Tsukamoto, Hiroshi; Tanaka, Kazuhiro; Matsunaga, Shigenori; Tanaka, Hiroki

    1992-08-01

    A Shape-controlled airfoil was developed by use of shape memory alloys (SMA). Two-way change in the blade shape was realized by use of a differential two-way element in which the two different shapes were memorized. The developed airfoil was tested in the wind tunnel in order to check the effect of the shape change on the characteristics of the airfoil. Flow visualization experiments in a smoke tunnel as well as the traverse of the wake behind the airfoil showed that the shape change by electrically heated SMA gives a marked change in flow around the airfoil near the stall angle of the original shape. As the result of this study, it was found that the developed SMA actuator is effective for the control of flow separation from the blade surface.

  11. Design and analytical study of a rotor airfoil

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.

    1978-01-01

    An airfoil section for use on helicopter rotor blades was defined and analyzed by means of potential flow/boundary layer interaction and viscous transonic flow methods to meet as closely as possible a set of advanced airfoil design objectives. The design efforts showed that the first priority objectives, including selected low speed pitching moment, maximum lift and drag divergence requirements can be met, though marginally. The maximum lift requirement at M = 0.5 and most of the profile drag objectives cannot be met without some compromise of at least one of the higher order priorities.

  12. Vortex scale of unsteady separation on a pitching airfoil.

    PubMed

    Fuchiwaki, Masaki; Tanaka, Kazuhiro

    2002-10-01

    The streaklines of unsteady separation on two kinds of pitching airfoils, the NACA65-0910 and a blunt trailing edge airfoil, were studied by dye flow visualization and by the Schlieren method. The latter visualized the discrete vortices shed from the leading edge. The results of these visualization studies allow a comparison between the dynamic behavior of the streakline of unsteady separation and that of the discrete vortices shed from the leading edge. The influence of the airfoil configuration on the flow characteristics was also examined. Furthermore, the scale of a discrete vortex forming the recirculation region was investigated. The non-dimensional pitching rate was k = 0.377, the angle of attack alpha(m) = 16 degrees and the pitching amplitude was fixed to A = +/-6 degrees for Re = 4.0 x 10(3) in this experiment.

  13. 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.

  14. Tables of properties of airfoil polynomials

    NASA Technical Reports Server (NTRS)

    Desmarais, Robert N.; Bland, Samuel R.

    1995-01-01

    This monograph provides an extensive list of formulas for airfoil polynomials. These polynomials provide convenient expansion functions for the description of the downwash and pressure distributions of linear theory for airfoils in both steady and unsteady subsonic flow.

  15. 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.

  16. 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. PMID:19151103

  17. Experimental aerodynamic characteristics of missiles with square cross sections

    NASA Astrophysics Data System (ADS)

    Zollars, G. J.; Yechout, R. T.; Daniel, D. C.; Lijewski, L. E.

    1983-05-01

    A series of quantitative and qualitative tests were conducted to expand the aerodynamic data base of square cross-section missiles. Quantitative tests included measuring forces and moments acting on square missiles, and measuring flowfield pressures on the leeward side of square missiles at various configurations and orientations in a subsonic wind tunnel. Force and moment data is presented showing the effects of variation in body corner radius, nose and fin shapes, pitch angle, and roll angle. Flowfield pressure and crossflow velocity data are presented for missiles of various pitch angles, roll angles, body corner radii, and fineness ratios (length to width ratio). In addition, flowfield data is shown along the axial length of a square missile. Qualitative tests included photographic tuft grids in the flowfield and photographing oil shear stress patterns on the surface of various missiles. These qualitative photographs are presented for various missile configurations and orientations.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Shape optimization of corrugated airfoils

    NASA Astrophysics Data System (ADS)

    Jain, Sambhav; Bhatt, Varun Dhananjay; Mittal, Sanjay

    2015-12-01

    The effect of corrugations on the aerodynamic performance of a Mueller C4 airfoil, placed at a 5° angle of attack and Re=10{,}000, is investigated. A stabilized finite element method is employed to solve the incompressible flow equations in two dimensions. A novel parameterization scheme is proposed that enables representation of corrugations on the surface of the airfoil, and their spontaneous appearance in the shape optimization loop, if indeed they improve aerodynamic performance. Computations are carried out for different location and number of corrugations, while holding their height fixed. The first corrugation causes an increase in lift and drag. Each of the later corrugations leads to a reduction in drag. Shape optimization of the Mueller C4 airfoil is carried out using various objective functions and optimization strategies, based on controlling airfoil thickness and camber. One of the optimal shapes leads to 50 % increase in lift coefficient and 23 % increase in aerodynamic efficiency compared to the Mueller C4 airfoil.

  3. 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.

  4. 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.

  5. 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

  6. Advanced Airfoils Boost Helicopter Performance

    NASA Technical Reports Server (NTRS)

    2007-01-01

    Carson Helicopters Inc. licensed the Langley RC4 series of airfoils in 1993 to develop a replacement main rotor blade for their Sikorsky S-61 helicopters. The company's fleet of S-61 helicopters has been rebuilt to include Langley's patented airfoil design, and the helicopters are now able to carry heavier loads and fly faster and farther, and the main rotor blades have twice the previous service life. In aerial firefighting, the performance-boosting airfoils have helped the U.S. Department of Agriculture's Forest Service control the spread of wildfires. In 2003, Carson Helicopters signed a contract with Ducommun AeroStructures Inc., to manufacture the composite blades for Carson Helicopters to sell

  7. Pressure distribution over an NACA 23012 airfoil with a slotted and a plain flap

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Delano, James B

    1938-01-01

    Report presents the results of pressure-distribution of an NACA 23012 airfoil equipped with a slotted flap and with a plain flap conducted in the 7 by 10-foot wind tunnel. A test installation was used in which the 7-foot-span airfoil was mounted vertically between the upper and lower sides of the closed test section so that two-dimensional flow was approximated. The pressures were measured on the upper and lower surfaces at one chord section both on the main airfoil and on the flaps for several different flap deflections and at several angles of attack. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated section coefficients for the airfoil-and-flap combinations and also for the flaps alone. The results are useful for application to rib and flap structural design; in addition, the plain-flap data furnish considerable information applicable to the structural design of plain ailerons.

  8. An Investigation of the Aerodynamic Characteristics of an Airplane Equipped with Several Different Sets of Wings

    NASA Technical Reports Server (NTRS)

    Crowley, J W , Jr; Green, M W

    1929-01-01

    This investigation was conducted by the National Advisory Committee for Aeronautics at Langley Field, Va., at the request of the Army Air Corps, for the purpose of comparing the full scale lift and drag characteristics of an airplane equipped with several sets of wings of commonly used airfoil sections. A Sperry Messenger Airplane with wings of R.A.F.-15, U.S.A.-5, U.S.A.-27, and Gottingen 387 airfoil sections was flown and the lift and drag characteristics of the airplane with each set of wings were determined by means of glide tests. The results are presented in tabular and curve form. (author)

  9. Hook nozzle arrangement for supporting airfoil vanes

    SciTech Connect

    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.

  10. 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.

  11. Hook nozzle arrangement for supporting airfoil vanes

    DOEpatents

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

    1996-02-20

    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. 8 figs.

  12. Pressure Distribution on a Slotted R.A.F. 31 Airfoil in the Variable Density Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1929-01-01

    Measurements were made in the variable density wind tunnel to determine the pressure distribution over one section of a R.A.F. 31 airfoil with a leading edge slot fully open. To provide data for the study of scale effect on this type of airfoil, the tests were conducted with air densities of approximately one and twenty atmospheres.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Calculation of the transient motion of elastic airfoils forced by control surface motion and gusts

    NASA Technical Reports Server (NTRS)

    Balakrishnan, A. V.; Edwards, J. W.

    1980-01-01

    The time-domain equations of motion of elastic airfoil sections forced by control surface motions and gusts were developed for the case of incompressible flow. Extensive use was made of special functions related to the inverse transform of Theodorsen's function. Approximations for the special cases of zero stream velocity, small time, large and time are given. A numerical solution technique for the solution of the general case is given. Examples of the exact transient response of an airfoil are presented.

  19. Self-sustained shock oscillations on airfoils at transonic speeds

    NASA Astrophysics Data System (ADS)

    Lee, B. H. K.

    2001-02-01

    Self-sustained shock wave oscillations on airfoils at transonic flow conditions are associated with the phenomenon of buffeting. The physical mechanisms of the periodic shock motion are not yet fully understood even though experiments performed over fifty years ago have demonstrated the presence of oscillatory shock waves on the airfoil surfaces at high subsonic speeds. The unsteady pressure fluctuations generated by the low-frequency large-amplitude shock motions are highly undesirable from the structural integrity and aircraft maneuverability point of view. For modern supercritical wing design with thick profiles, the shock-induced fluctuations are particularly severe and methods to reduce the shock wave amplitudes to lower values or even to delay the oscillations to higher Mach numbers or incidence angles will result in expanding the buffet boundary of the airfoil. This review begins with a recapitulation of the classical work on shock-induced bubble separation and trailing edge separation of a turbulent boundary layer. The characteristics of the unsteady pressure fluctuations are used to classify the types of shock-boundary layer interaction. The various modes of shock wave motion for different flow conditions and airfoil configurations are described. The buffet boundaries obtained using the standard trailing edge pressure divergence technique and an alternative approach of measuring the divergence of normal fluctuating forces are compared to show the equivalence. The mechanisms of self-sustained shock oscillations are discussed for symmetrical circular-arc airfoils at zero incidence and for supercritical airfoils at high incidence angles with fully separated flows. The properties of disturbances in the wake are examined from linear stability analysis of two-dimensional compressible flows. The advances in high-speed computing make predictions of buffeting flows possible. Navier-Stokes solvers and approximate boundary layer-inviscid flow interaction methods are

  20. 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.

  1. Wind-tunnel test of the S814 thick root airfoil

    SciTech Connect

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

    1995-01-01

    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 (c{sub 1,max} 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 it a high lift coefficient. To accomplish the objective, a two-dimensional wind-tunnel test of the S814 thick root airfog was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory. Data were obtained for transition-free and transition-fixed conditions at Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds numbers of 1.5 {times} l0{sup 6}, the transition-free c{sub 1,max} is 1.3 which satisfies the design specification. However, this value is significantly lower than the predicted c{sub 1,max} of almost l.6. With transition-fixed at the is 1.2. The difference in c{sub 1,max} 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 c{sub 1,max} tip-region airfoils for rotor blades 10 to 15 meters in length.

  2. 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.

  3. Airfoil data sensitivity analysis for actuator disc simulations used in wind turbine applications

    NASA Astrophysics Data System (ADS)

    Nilsson, Karl; Breton, Simon-Philippe; Sørensen, Jens N.; Ivanell, Stefan

    2014-06-01

    To analyse the sensitivity of blade geometry and airfoil characteristics on the prediction of performance characteristics of wind farms, large-eddy simulations using an actuator disc (ACD) method are performed for three different blade/airfoil configurations. The aim of the study is to determine how the mean characteristics of wake flow, mean power production and thrust depend on the choice of airfoil data and blade geometry. In order to simulate realistic conditions, pre-generated turbulence and wind shear are imposed in the computational domain. Using three different turbulence intensities and varying the spacing between the turbines, the flow around 4-8 aligned turbines is simulated. The analysis is based on normalized mean streamwise velocity, turbulence intensity, relative mean power production and thrust. From the computations it can be concluded that the actual airfoil characteristics and blade geometry only are of importance at very low inflow turbulence. At realistic turbulence conditions for an atmospheric boundary layer the specific blade characteristics play an minor role on power performance and the resulting wake characteristics. The results therefore give a hint that the choice of airfoil data in ACD simulations is not crucial if the intention of the simulations is to compute mean wake characteristics using a turbulent inflow.

  4. 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.

  5. 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.

  6. Airfoil seal system for gas turbine engine

    SciTech Connect

    Diakunchak, Ihor S.

    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.

  7. 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.

  8. Turbulent separated flow over and downstream of a two-element airfoil

    NASA Technical Reports Server (NTRS)

    Adair, D.; Horne, W. C.

    1989-01-01

    Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analyzed. 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 airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 x 10 to the 6th in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. 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.

  9. 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.

  10. Analysis of the development of dynamic stall based on oscillating airfoil experiments

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    The effects of dynamic stall on airfoils oscillating in pitch were investigated by experimentally determining the viscous and inviscid characteristics of the airflow on the NACA 0012 airfoil and on several leading-edge modifications. The test parameters included a wide range of frequencies, Reynolds numbers, and amplitudes-of-oscillation. Three distinct types of separation development were observed within the boundary layer, each leading to classical dynamic stall. The NACA 0012 airfoil is shown to stall by the mechanism of abrupt turbulent leading-edge separation. A detailed step-by-step analysis of the events leading to dynamic stall, and of the results of the stall process, is presented for each of these three types of stall. Techniques for flow analysis in the dynamic stall environment are discussed. A method is presented that reduces most of the oscillating airfoil normal force and pitching-moment data to a single curve, independent of frequency or Reynolds number.

  11. 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.

  12. 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.

  13. Experimental Droplet Impingement on Several Two-Dimensional Airfoils with Thickness Ratios of 6 to 16 Percent

    NASA Technical Reports Server (NTRS)

    Gelder, Thomas F.; Smyers, William H., Jr.; VonGlahn, Uwe

    1956-01-01

    The rate and area of cloud droplet impingement on several two-dimensional swept and unswept airfoils were obtained experimentally in the NACA Lewis icing tunnel with a dye-tracer technique. Airfoil thickness ratios of 6 to 16 percent; angles of attack from 0 deg to 12 deg, and chord sizes from 13 to 96 inches were included in the study. The data were obtained at 152 knots and are extended to other conditions by dimensionless impingement parameters. In general, the data show that the total and local collection efficiencies and impingement limits are primary functions of the modified inertia parameter (in which airspeed, droplet size, and body size are the most significant variables) and the airfoil thickness ratio. Local collection efficiencies and impingement limits also depend on angle of attack. Secondary factors affecting impingement characteristics are airfoil shape, camber, and sweep angle. The impingement characteristics obtained experimentally for the airfoils were within +/-10 percent on the average of the characteristics calculated from theoretical trajectories. Over the range of conditions studied, the experimental data demonstrate that a specific method can be used to predict the impingement characteristics of swept airfoils with large aspect ratios from the data for unswept airfoils of the same series.

  14. 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.

  15. 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.

  16. 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)

  17. 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.

  18. 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.

  19. 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

  20. 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…

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Aerodynamic effects of leading-edge serrations on a two-dimensional airfoil

    NASA Technical Reports Server (NTRS)

    Soderman, P. T.

    1972-01-01

    An investigation was conducted to determine the flow field and aerodynamic effects of leading-edge serrations on a two-dimensional airfoil at a Mach number of 0.13. The model was a NACA 66-012 airfoil section with a 0.76 m (30 in.) chord, 1.02 m (40 in.) span, and floor and end plates. It was mounted in the Ames 7- by 10-Foot Wind Tunnel. Serrated brass strips of various sizes and shapes were attached to the model in the region of the leading edge. Force and moment data, and photographs of tuft patterns and of oil flow patterns are presented. Results indicated that the smaller serrations, when properly placed on the airfoil, created vortices that increased maximum lift and angle of attack for maximum lift. The drag of the airfoil was not increased by these serrations at airfoil angles of attack near zero and was decreased at large angles of attack. Important parameters were serration size, position on the airfoil, and spacing between serrations.

  6. 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.

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

    NASA Technical Reports Server (NTRS)

    Nagamatsu, H. T.; Orozco, R. D.; Ling, D. C.

    1984-01-01

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

  8. On the birth of stall cells on airfoils

    NASA Astrophysics Data System (ADS)

    Rodríguez, Daniel; Theofilis, Vassilis

    2011-06-01

    Critical point theory asserts that two-dimensional topologies are defined as degeneracies and any three-dimensional disturbance of a two-dimensional flow will lead to a new three-dimensional flowfield topology, regardless of the disturbance amplitude. Here, the topology of the composite flowfields reconstructed by linear superposition of the two-dimensional flow around a stalled airfoil and the leading stationary three-dimensional global eigenmode has been studied. In the conditions monitored the two-dimensional flow is steady and laminar and is separated over a fraction of the suction side, while the amplitudes considered in the linear superposition are small enough for the linearization assumption to be valid. The multiple topological bifurcations resulting have been analysed in detail; the surface streamlines generated by the leading stationary global mode of the separated flow have been found to be strongly reminiscent of the characteristic stall cells, observed experimentally on airfoils just beyond stall in both laminar and turbulent flow.

  9. Performance measurements of an airfoil at low Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Mcghee, Robert J.; Walker, Betty S.

    1989-01-01

    Performance characteristics of an Eppler 387 airfoil using both direct (force) and indirect (pressure) measurement techniques have been obtained at Reynolds numbers from 60,000 to 460,000 in the Langley Low-Turbulence Pressure Tunnel. Lift, drag, and pitching-moment data were obtained from two internally-mounted strain-gage balances specifically designed for small aerodynamic loads. Comparisons of these results with data from a pressure model of an Eppler 387 airfoil are included. Drag data for both models using the wake traverse method are compared with the balance data. Oil flow visualization and surface mounted hot-film sensors were used to determine laminar-separation and turbulent-reattachment locations. Problems associated with obtaining accurate wind-tunnel data at low Reynolds numbers are discussed.

  10. 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.

  11. 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.

  12. 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.

  13. 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

  14. 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

  15. Icing Research Tunnel Test Section

    NASA Technical Reports Server (NTRS)

    1968-01-01

    Icing Research Tunnel Test Section NASA technician measuring ice deposits on an airfoil after completing a test at the Lewis Research Center. NASA Lewis is now known as John H. Glean Research Center at Lewis Field.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Aerodynamics Investigation of Faceted Airfoils at Low Reynolds Number

    NASA Astrophysics Data System (ADS)

    Napolillo, Zachary G.

    The desire and demand to fly farther and faster has progressively integrated the concept of optimization with airfoil design, resulting in increasingly complex numerical tools pursuing efficiency often at diminishing returns; while the costs and difficulty associated with fabrication increases with design complexity. Such efficiencies may often be necessary due to the power density limitations of certain aircraft such as small unmanned aerial vehicles (UAVs) and micro air vehicles (MAVs). This research, however, focuses on reducing the complexity of airfoils for applications where aerodynamic performance is less important than the efficiency of manufacturing; in this case a Hybrid Projectile. By employing faceted sections to approximate traditional contoured wing sections it may be possible to expedite manufacturing and reduce costs. We applied this method to the development of a low Reynolds number, disposable Hybrid Projectile requiring a 4.5:1 glide ratio, resulting in a series of airfoils which are geometric approximations to highly contoured cross-sections called ShopFoils. This series of airfoils both numerically and experimentally perform within a 10% margin of the SD6060 airfoil at low Re. Additionally, flow visualization has been conducted to qualitatively determine what mechanisms, if any, are responsible for the similarity in performance between the faceted ShopFoil sections and the SD6060. The data obtained by these experiments did not conclusively reveal how the faceted surfaces may influence low Re flow but did indicate that the ShopFoil s did not maintain flow attachment at higher angles of attack than the SD6060. Two reasons are provided for the unexpected performance of the ShopFoil: one is related to downwash effects, which are suspected of placing the outer portion of the span at an effective angle of attack where the ShopFoils outperform the SD6060; the other is the influence of the tip vortex on separation near the wing tips, which possibly

  1. 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.

  2. 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.

  3. The aerodynamic design of an advanced rotor airfoil

    NASA Technical Reports Server (NTRS)

    Blackwell, J. A., Jr.; Hinson, B. L.

    1978-01-01

    An advanced rotor airfoil, designed utilizing supercritical airfoil technology and advanced design and analysis methodology is described. The airfoil was designed subject to stringent aerodynamic design criteria for improving the performance over the entire rotor operating regime. The design criteria are discussed. The design was accomplished using a physical plane, viscous, transonic inverse design procedure, and a constrained function minimization technique for optimizing the airfoil leading edge shape. The aerodynamic performance objectives of the airfoil are discussed.

  4. 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.

  5. 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.

  6. 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.

  7. Effect of Intercycle Ice Accretions on Airfoil Performance

    NASA Technical Reports Server (NTRS)

    Broeren, Andy P.; Bragg, Michael B.; Addy, Harold E., Jr.

    2003-01-01

    This paper presents the results of an experimental study designed to characterize and evaluate the aerodynamic performance penalties of residual and intercycle ice accretions that result from the cyclic operation of a typical aircraft deicing system. Icing wind tunnel tests were carried out on a 36-inch chord NACA 23012 airfoil section equipped with a pneumatic deicer for several different FAR 25 Appendix C cloud conditions. Results from the icing tests showed that the intercycle ice accretions were much more severe in terms of size and shape than the residual ice accretions. Molds of selected intercycle ice shapes were made and converted to castings that were attached to the leading edge of a 36-inch chord NACA 23012 airfoil model for aerodynamic testing. The aerodynamic testing revealed that the intercycle ice shapes caused a significant performance degradation. Maximum lift coefficients were typically reduced about 60% from 1.8 (clean) to 0.7 (iced) and stall angles were reduced from 17 deg. (clean) to 9 deg. (iced). Changes in the Reynolds number (from 2.0 x 10(exp 6) to 10.5 x 10(exp 6) and Mach number (from 0.10 to 0.28) did not significantly affect the iced-airfoil performance.

  8. 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.

  9. Phenotypic characteristics of isolates of Aspergillus section Fumigati from different geographic origins and their relationships with genotypic characteristics

    PubMed Central

    2011-01-01

    Background Epidemiological studies worldwide have shown that A. fumigatus exhibits important phenotypic and genotypic diversity, and these findings have been of great importance in improving the diagnosis and treatment of diseases caused by this fungus. However, few studies have been carried out related to the epidemiology of this fungus in Latin America. This study´s aim is to report on the epidemiology of the fungus by analyzing the phenotypic variability of Aspergillus section Fumigati isolates from different Latin American countries and the relationship between this variability, the geographical origin and genotypic characteristics. Methods We analyzed the phenotypic characteristics (macro- and micromorphology, conidial size, vesicles size, antifungal susceptibility and thermotolerance at 28, 37 and 48°C) of A. section Fumigati isolates from Mexico (MX), Argentina (AR), Peru (PE) and France (FR). The results were analyzed using analysis of variance (ANOVA) and Tukey's multiple comparison test to detect significant differences. Two dendrograms among isolates were obtained with UPGMA using the Euclidean distance index. One was drawn for phenotypic data, and the other for phenotypic and genotypic data. A PCoA was done for shown isolates in a space of reduced dimensionality. In order to determine the degree of association between the phenotypic and genotypic characteristics AFLP, we calculated the correlation between parwise Euclidean distance matrices of both data sets with the nonparametric Mantel test. Results No variability was found in the macromorphology of the studied isolates; however, the micromorphology and growth rate showed that the PE isolates grew at a faster rate and exhibited the widest vesicles in comparison to the isolates from MX, AR and FR. The dendrogram constructed with phenotypic data showed three distinct groups. The group I and II were formed with isolates from PE and FR, respectively, while group III was formed with isolates from MX and

  10. 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.

  11. Aerodynamic Control of a Dynamically Pitching Airfoil using Transitory Pulsed Actuation

    NASA Astrophysics Data System (ADS)

    Tan, Yuehan; Crittenden, Thomas; Glezer, Ari

    2015-11-01

    Transitory control and regulation of trapped vorticity concentrations are exploited for control of the aerodynamic loads on an airfoil that is dynamically pitching beyond the dynamic stall margin in wind tunnel experiments. 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 on characteristic time scales that are 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. The aerodynamic loading can be significantly altered by a number of actuation programs using multiple actuation bursts during the pitch cycle. While actuation during the upstroke primarily affects the formation, evolution, and advection of the dynamic stall vortex, actuation during the downstroke accelerates flow attachment. Superposition of such actuation programs leads to enhancement of cycle lift and pitch stability, and reduced cycle hysteresis.

  12. 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.

  13. 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

  14. Effect of Fuselage Cross Sections on Aerodynamic Characteristics of Reusable Launch Vehicles in Subsonic Flow

    NASA Astrophysics Data System (ADS)

    Tadakuma, Kenji; Morita, Wataru; Aso, Shigeru; Tani, Yasuhiro

    An experimental study on aerodynamic effect of RLVs (Reusable Launch Vehicles) due to fuselage cross sections has been conducted in subsonic flow. Three fuselage models and two wing-body models have been considered. Fuselage models have a circular, a square and a triangular cross section. Wing-body models have a square and a triangular cross section with wings. Experiments have been conducted under test conditions of free-stream Mach number M∞=0.3 and Reynolds number Re=3.2×106. Aerodynamic forces are measured and flow fields are visualized by smoke-wire technique and oil-flow technique. Results show that fuselage cross sections have much effect on whole aerodynamic characteristics, the fuselage model with a triangular cross section has higher lift coefficient in high angle of attack region than that of the other fuselage models and the wing-body model with a triangular fuselage cross section does not stall till high angle of attack region compared with the “Square” fuselage wing-body model.

  15. 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.

  16. Airfoil for a gas turbine

    SciTech Connect

    Liang, George

    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.

  17. Turbine airfoil to shroud attachment method

    DOEpatents

    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.

  18. 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.

  19. 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.

  20. TAIR: A transonic airfoil analysis computer code

    NASA Technical Reports Server (NTRS)

    Dougherty, F. C.; Holst, T. L.; Grundy, K. L.; Thomas, S. D.

    1981-01-01

    The operation of the TAIR (Transonic AIRfoil) computer code, which uses a fast, fully implicit algorithm to solve the conservative full-potential equation for transonic flow fields about arbitrary airfoils, is described on two levels of sophistication: simplified operation and detailed operation. The program organization and theory are elaborated to simplify modification of TAIR for new applications. Examples with input and output are given for a wide range of cases, including incompressible, subcritical compressible, and transonic calculations.

  1. Analysis of Non-symmetrical Flapping Airfoils

    NASA Astrophysics Data System (ADS)

    Beng Tay, Wee; Lim, Kah Bin

    2007-11-01

    Simulations have been done to assess the performance of different types of non-symmetrical airfoils on lift, thrust and propulsive efficiency under different flapping configurations at a Reynolds number of 10,000. The variables studied include the Stroudal number, reduced frequency, pitch angle and phase angle difference. In order to analyze the variables more efficiently, the Design of Experiments using the response surface methodology is applied. The simulation results show that besides the flapping configuration, airfoil shape also has a profound effect on the efficiency, thrust and lift production. The 4 factors have different levels of significance on the responses, indicating the shape of the airfoil plays a part as well. Thrust production depends more heavily on these parameters, rather than the shape of the airfoil. On the other hand, lift production is primarily dominated by its airfoil shape. Efficiency falls somewhere in between. Two-factor interactions among the variables also exist in efficiency and thrust production. Vorticity plots are analyzed to explain some of the results. Overall, the s1020 airfoil is able to provide relatively good efficiency and at the same time generate high thrust and lift force. These results can be used to help in the design of a better ornithopter's wing.

  2. 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.

  3. The distribution characteristics of pollutants released at different cross-sectional positions of a river.

    PubMed

    Huang, Heqing; Chen, Guang; Zhang, Qian-Feng

    2010-05-01

    The distribution characteristics of heavier or lighter pollutants released at different cross-sectional positions of a wide river is investigated with a well-tested three-dimensional numerical model of gravity flows based on Reynolds-Averaged Navier-Stokes equations and turbulence k-epsilon model. By focusing on investigating the influences of flow and buoyancy on pollutants, it is found that while carrying by the river flow downstream: i) a heavier pollutant released from the cross-sectional side position, forms transverse oscillation between two banks with decreased amplitude, i.e. forms kind of helical flow pattern along the straight part of channel bed; ii) a heavier pollutant released from the cross-sectional middle position, forms collapse oscillation in the middle of the straight channel part with reduced amplitude; iii) in the downstream sinuous channel, heavier pollutant is of higher concentration on the outer side of channel bends; iv) a light pollutant released from the cross-sectional side position, slips partly to the other side of the river, resulting in higher concentrations on two sides of the channel top; v) a light pollutant released from the cross-sectional middle position, splits into two parts symmetrically along two sides of the channel top; vi) in the downstream sinuous channel, light pollutant presents higher concentration on the inner side of channel bends. These findings may assist in cost-effective scientific countermeasures to be taken for accidental or planned pollutant releases into a river.

  4. Effect of pre-strain and excess length on unsteady fluid-structure interactions of membrane airfoils

    NASA Astrophysics Data System (ADS)

    Rojratsirikul, P.; Wang, Z.; Gursul, I.

    2010-04-01

    Aerodynamic characteristics of two-dimensional membrane airfoils were experimentally investigated in a wind tunnel. The effects of the membrane pre-strain and excess length on the unsteady aspects of the fluid-structure interaction were studied. The deformation of the membrane as a function of angle of attack and free-stream velocity was measured using a high-speed camera. These measurements were complemented by the measurements of unsteady velocity field with a high frame-rate Particle Image Velocimetry (PIV) system as well as smoke visualization. Membrane airfoils with excess length exhibit higher vibration modes, earlier roll-up of vortices, and smaller separated flow regions, whereas the membranes with pre-strain generally behave more similarly to a rigid airfoil. Measured frequencies of the membrane vibrations suggest a possible coupling with the wake instabilities at high incidences for all airfoils.

  5. 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.

  6. Detailed measurements of the flowfield in the vicinity of an airfoil with glaze ice

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.; Coirier, W. J.

    1985-01-01

    An experimental study has been conducted in the OSU subsonic tunnel to measure the characteristics of the separation bubble on an airfoil with glaze ice. A measured glaze ice accretion on a NACA 0012 airfoil was simulated in wood for this dry tunnel test. The 21 inch chord model was pressure belted and the ice shape internally tapped to obtain surface pressures, lift and moment coefficients. A wake survey probe was used to obtain airfoil drag. The separation bubble was explored by measuring the time averaged velocities using a split film probe. The probe was positioned using a computer controlled two-dimensional traversing system. In this paper, airfoil lift, drag, and moment coefficient data are compared for the airfoil with and without glaze ice. Velocity profiles in the separation bubble are presented for several chordwise stations at three angles of attack. The ice shape caused a severe lift and drag penalty. The velocity profiles show clearly the large bubble geometry, regions of reversed flow, and bubble reattachment.

  7. Experimental Study of the Power Profile Airfoil Equipped with Plasma Flow Control

    NASA Astrophysics Data System (ADS)

    Daniel, Libin; Jacob, Jamey

    2013-11-01

    This presentation discusses results from an experimental study of the power profile airfoil at low Reynolds number. The power profile airfoil was developed by AMO Smith and consists of a blunt trailing edge shape with two wall jets near the trailing edge. The replacement of streamlining with properly designed blowing is used to prevent flow separation and additionally offers potential applications as a powered high-lift system, propulsive system, or low inertia control device. The 2D wind-tunnel model consists of the 22.5% thick power profile airfoil equipped with a movable trailing edge plug to direct flow along the trailing edge streamline. Compressed air was passed into the model via a plenum with flow conditioning devices to create pressure backdrop to allow uniform blowing at the trailing edge. The effects of varying jet momentum coefficient and trailing edge positioning on the aerodynamic characteristics are observed with both wake surveys and PIV. The impact of plasma synthetic jet actuators (PSJA) placed along the trailing edge of the power profile airfoil is also discussed. PSJA operation is compared to the baseline power profile airfoil both alone and working with the blowing to provide additional control authority.

  8. Numerical bifurcation analysis of static stall of airfoil and dynamic stall under unsteady perturbation

    NASA Astrophysics Data System (ADS)

    Liu, Yan; Li, Kailun; Zhang, Jiazhong; Wang, Hang; Liu, Liguang

    2012-08-01

    By the finite element method combined with Arbitrary-Lagrangian-Eulerian (ALE) frame and explicit Characteristic Based Split Scheme (CBS), the complex flows around stationary and sinusoidal pitching airfoil are studied numerically. In particular, the static and dynamic stalls are analyzed in detail, and the natures of the static stall of NACA0012 airfoil are given from viewpoint of bifurcations. Following the bifurcation in Map, the static stall is proved to be the result from saddle-node bifurcation which involves both the hysteresis and jumping phenomena, by introducing a Map and its Floquet multiplier, which is constructed in the numerical simulation of flow field and related to the lift of the airfoil. Further, because the saddle-node bifurcation is sensitive to imperfection or perturbation, the airfoil is then subjected to a perturbation which is a kind of sinusoidal pitching oscillation, and the flow structure and aerodynamic performance are studied numerically. The results show that the large-scale flow separation at the static stall on the airfoil surface can be removed or delayed feasibly, and the ensuing lift could be enhanced significantly and also the stalling incidence could be delayed effectively. As a conclusion, it can be drawn that the proper external excitation can be considered as a powerful control strategy for the stall. As an unsteady aerodynamic behavior of high angle of attack, the dynamic stall can be investigated from viewpoint of nonlinear dynamics, and there exists a rich variety of nonlinear phenomena, which are related to the lift enhancement and drag reduction.

  9. 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.

  10. An empirically derived basis for calculating the area, rate, and distribution of water-drop impingement on airfoils

    NASA Technical Reports Server (NTRS)

    Bergrun, Norman R

    1952-01-01

    An empirically derived basis for predicting the area, rate, and distribution of water-drop impingement on airfoils of arbitrary section is presented. The concepts involved represent an initial step toward the development of a calculation technique which is generally applicable to the design of thermal ice-prevention equipment for airplane wing and tail surfaces. It is shown that sufficiently accurate estimates, for the purpose of heated-wing design, can be obtained by a few numerical computations once the velocity distribution over the airfoil has been determined. The calculation technique presented 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.

  11. 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.

  12. 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.

  13. Impingement cooling with film coolant extraction in the airfoil leading edge regions

    NASA Astrophysics Data System (ADS)

    Li, Liguo; Li, Zhaohui

    An extensive experimental study is conducted to determine the heat transfer characteristics of arrays of air jets impinging on perforated target surfaces in turbine blade leading edge regions by six large-scale models. The relations of pressure loss and Nusselt number to jet Reynolds number are obtained in a wide range of parameter combinations of interest in cooled airfoil practice for various models, respectively. These parameter combinations are covered in a test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge curvature radius-to-diameter ratio, jet pitch-to-diameter ratio, and jet impingement gap-to-diameter ratio.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. Wave propagation in beams with periodic arrays of airfoil-shaped resonating units

    NASA Astrophysics Data System (ADS)

    Casadei, Filippo; Bertoldi, Katia

    2014-12-01

    This paper presents an analytical and numerical study on the dispersion properties of an Euler-Bernoulli beam immersed in a steady fluid flow with periodic arrays of airfoil-shaped vibration absorbers attached to it. The resonance characteristics of the airfoils generate strong attenuation of flexural waves in the beam occurring at frequencies defined by the properties of the airfoils and the speed of the incident fluid. Analytical and numerical tools are developed to investigate the effects of the incident flow on the dispersion properties and the bandgaps of the system. Both steady and unsteady aerodynamic models are used to model the lift force and the pitching moment acting on the resonators and their effect on the dispersion relations of the system is evaluated. Finally, an effective medium description of the beam is developed to capture its behavior at long-wavelengths. In this regime, the system can be effectively considered as an acoustic metamaterial with adaptive dispersion properties.

  19. Effects of icing on the aerodynamic performance of high lift airfoils

    NASA Technical Reports Server (NTRS)

    Sankar, L. N.; Phaengsook, N.; Bangalore, A.

    1993-01-01

    A 2D compressible Navier-Stokes solver capable of analyzing multi-element airfoils is described. The flow field is divided into multiple zones. In each zone, the governing equations are solved using an implicit finite difference scheme. The flow solver is validated through a study of the aerodynamic characteristics of a GA(W)-1 configuration, for which good quality measured surface pressure data and load data are available. The solver is next applied to a study of the effects of icing on an iced 5-element airfoil configuration, experimentally studied at NASA Lewis Research Center. It is demonstrated that the formation of ice over the leading edge slat and the main airfoil can lead to significant flow separation, and a significant loss in lift, compared to clean configurations.

  20. Wind Tunnel Tests of Wind Turbine Airfoils at High Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Llorente, E.; Gorostidi, A.; Jacobs, M.; Timmer, W. A.; Munduate, X.; Pires, O.

    2014-06-01

    Wind tunnel tests have been performed to measure the two-dimensional aerodynamic characteristics of two different airfoil families at high Reynolds numbers (from 3 to 12 millions) in the DNW High Pressure Wind Tunnel in Gottingen (HDG), Germany. Also, tests at a Reynolds number of 3 millions have been performed in the Low-Speed Low- Turbulence Wind Tunnel of Delft University, The Netherlands. The airfoils tested belong to two wind turbine dedicated families: the TU-Delft DU family and the ACCIONA Windpower AWA family that was designed in collaboration with CENER. Reynolds number effects on airfoil performance have been obtained in the range of 3 to 12 millions. The availability of data from two different wind tunnels has brought the opportunity to cross compare the results from the two facilities.

  1. Calculation of boundary layers near the stagnation point of an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Cebeci, T.; Carr, L. W.

    1983-01-01

    The results of an investigation of boundary layers close to the stagnation point of an oscillating airfoil are reported. Two procedures for generating initial conditions, the characteristics box scheme and a quasi-static approach, were investigated, and the quasi-static approach was shown to be appropriate provided the initial region was far from any flow separation. With initial conditions generated in this way, the unsteady boundary layer equations were solved for the flow in the leading edge region of a NACA 0012 airfoil oscillating from 0 to 5 deg. Results were obtained for both laminar and turbulent flow, and, in the latter case, the effect of transition was assessed by specifying its occurrence at different locations. The results demonstrate the validity of the numerical scheme and suggest that the procedures should be applied to calculation of the entire flow around oscillating airfoils.

  2. Two-dimensional wind tunnel test of an oscillating rotor airfoil, volume 1

    NASA Technical Reports Server (NTRS)

    Dadone, L. U.

    1977-01-01

    A two dimensional wind tunnel test was conducted to obtain the quasisteady and unsteady characteristics of an advanced airfoil designed for helicopter rotor applications. Differential pressures were measured at 17 locations along the chord of the airfoil model. The airfoil motions were sinusoidal forced-pitch oscillations about the quarter chord at amplitudes varying from 2.5 to 10.0 degrees and at frequencies from 23 Hz to 90 Hz. The quasisteady tests were conducted at Mach numbers from 0.2 to 0.9, and the oscillatory tests between M = 0.2 and M = 0.7. At quasisteady conditions a limited number of drag measurements were made with a wake-traversing probe.

  3. Method for calculating wing characteristics by lifting-line theory using nonlinear section lift data

    NASA Technical Reports Server (NTRS)

    Sivells, James C; Neely, Robert H

    1947-01-01

    A method is presented for calculating wing characteristics by lifting-line theory using nonlinear section lift data. Material from various sources is combined with some original work into the single complete method described. Multhopp's systems of multipliers are employed to obtain the induced angle of attack directly from the spanwise lift distribution. Equations are developed for obtaining these multipliers for any even number of spanwise stations, and values are tabulated for 10 stations along the semispan for asymmetrical, symmetrical, and antisymmetrical lift distributions. In order to minimize the computing time and to illustrate the procedures involved, simplified computing forms containing detailed examples are given for symmetrical lift distributions. Similar forms for asymmetrical and antisymmetrical lift distributions, although not shown, can be readily constructed in the same manner as those given. The adaptation of the method for use with linear section lift data is also illustrated. The adaptation has been found to require less computing time than most existing methods.

  4. Experimental studies of flow separation of three airfoils at low speeds

    NASA Technical Reports Server (NTRS)

    Rodgers, E. J.; Seetharam, H. C.; Wentz, W. H., Jr.

    1982-01-01

    Wind tunnel tests were conducted on the NASA LS(1)-0421 Mod, NACA 2412 and NASA GA(W)-2 airfoil sections at a Reynolds number of 2.2 x 10(6) and a Mach number of 0.13. Detailed measurements of flow fields associated with turbulent boundary layers of these airfoils were obtained at pre-stall, near-stall, and post-stall angles of attack. Velocity and pressure survey results over the airfoil and in the associated wake, are presented for fully attached flow conditions through the stalled flow condition. Extensive force, pressure, tuft survey, hot-film survey, local skin friction and boundary layer data are also included.

  5. On the validation of a code and a turbulence model appropriate to circulation control airfoils

    NASA Technical Reports Server (NTRS)

    Viegas, J. R.; Rubesin, M. W.; Maccormack, R. W.

    1988-01-01

    A computer code for calculating flow about a circulation control airfoil within a wind tunnel test section has been developed. This code is being validated for eventual use as an aid to design such airfoils. The concept of code validation being used is explained. The initial stages of the process have been accomplished. The present code has been applied to a low-subsonic, 2-D flow about a circulation control airfoil for which extensive data exist. Two basic turbulence models and variants thereof have been successfully introduced into the algorithm, the Baldwin-Lomax algebraic and the Jones-Launder two-equation models of turbulence. The variants include adding a history of the jet development for the algebraic model and adding streamwise curvature effects for both models. Numerical difficulties and difficulties in the validation process are discussed. Turbulence model and code improvements to proceed with the validation process are also discussed.

  6. A visualization study of the interaction of a free vortex with the wake behind an airfoil

    NASA Astrophysics Data System (ADS)

    Świrydczuk, J.

    1990-06-01

    The two-dimensional interaction of a single vortex with a thin symmetrical airfoil and its vortex wake has been investigated in a low turbulence wind tunnel having velocity of about 2 m/s in the measuring section. The flow Reynolds number based on the airfoil chord length was 4.5 × 103. The investigation was carried out using a smoke-wire visualization technique with some support of standard hot-wire measurements. The experiment has proved that under certain conditions the vortex-airfoil-wake interaction leads to the formation of new vortices from the part of the wake positioned closely to the vortex. After the formation, the vortices rotate in the direction opposite to that of the incident vortex.

  7. 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.

  8. 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.

  9. 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.

  10. An experimental study of airfoil-spoiler aerodynamics

    NASA Technical Reports Server (NTRS)

    Mclachlan, B. G.; Karamcheti, K.

    1985-01-01

    The steady/unsteady flow field generated by a typical two dimensional airfoil with a statically deflected flap type spoiler was investigated. Subsonic wind tunnel tests were made over a range of parameters: spoiler deflection, angle of attack, and two Reynolds numbers; and comprehensive measurements of the mean and fluctuating surface pressures, velocities in the boundary layer, and velocities in the wake. Schlieren flow visualization of the near wake structure was performed. The mean lift, moment, and surface pressure characteristics are in agreement with previous investigations of spoiler aerodynamics. At large spoiler deflections, boundary layer character affects the static pressure distribution in the spoiler hingeline region; and, the wake mean velocity fields reveals a closed region of reversed flow aft of the spoiler. It is shown that the unsteady flow field characteristics are as follows: (1) the unsteady nature of the wake is characterized by vortex shedding; (2) the character of the vortex shedding changes with spoiler deflection; (3) the vortex shedding characteristics are in agreement with other bluff body investigations; and (4) the vortex shedding frequency component of the fluctuating surface pressure field is of appreciable magnitude at large spoiler deflections. The flow past an airfoil with deflected spoiler is a particular problem in bluff body aerodynamics is considered.

  11. 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.

  12. Reversed cowl flap inlet thrust augmentor. [with adjustable airfoil

    NASA Technical Reports Server (NTRS)

    Cheng, D. Y. (Inventor)

    1975-01-01

    An adjustable airfoil is described for varying the geometry of a jet inlet and an ejector inlet in a jet engine for providing thrust augmentation and noise reduction. The airfoil comprises essentially a plurality of segments which are extended radially outward and retracted relative to the longitudinal axis of the engine as a function of a change in the pressure differential between the upstream and downstream surfaces of the airfoil. A servo mechanism responsive to the change in the pressure differential is coupled to the airfoil to extend and retract the airfoil segments to maintain the pressure at a maximum on the downstream side of the airfoil relative to the pressure on the upstream side of the airfoil. At low speeds, such as at take-offs and landings, the airfoil is fully extended while at high speeds it is fully retracted.

  13. 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.

  14. Lift-Enhancing Tabs on Multielement Airfoils

    NASA Technical Reports Server (NTRS)

    Ross, James C.; Storms, Bruce L.; Carrannanto, Paul G.

    1995-01-01

    The use of flat-plate tabs (similar to Gurney flaps) to enhance the lift of multielement airfoils is extended here by placing them on the pressure side and near the trailing edge of the main element rather than just on the furthest downstream wing element. The tabs studied range in height from 0.125 to 1.25% of the airfoil reference chord. In practice, such tabs would be retracted when the high-lift system is stowed. The effectiveness of the concept was demonstrated experimentally and computationally on a two-dimensional NACA 63(sub 2)-215 Mod B airfoil with a single-slotted, 30%-chord flap. Both the experiments and computations showed that the tabs significantly increase the lift at a given angle of attack and the maximum lift coefficient of the airfoil. The computational results showed that the increased lift was a result of additional turning of the flow by the tab that reduced or eliminated now separation on the flap. The best configuration tested, a 0.5%-chord tab placed 0.5% chord upstream of the trailing edge of the main element, increased the maximum lift coefficient of the airfoil by 12% and the maximum lift-to-drag ratio by 40%.

  15. Numerical simulations of iced airfoils and wings

    NASA Astrophysics Data System (ADS)

    Pan, Jianping

    A numerical study was conducted to understand the effects of simulated ridge and leading-edge ice shapes on the aerodynamic performance of airfoils and wings. In the first part of this study, a range of Reynolds numbers and Mach numbers, as well as ice-shape sizes and ice-shape locations were examined for various airfoils with the Reynolds-Averaged Navier-Stokes approach. Comparisons between simulation results and experimental force data showed favorable comparison up to stall conditions. At and past stall condition, the aerodynamic forces were typically not predicted accurately for large upper-surface ice shapes. A lift-break (pseudo-stall) condition was then defined based on the lift curve slope change. The lift-break angles compared reasonably with experimental stall angles, and indicated that the critical ice-shape location tended to be near the location of minimum pressure and the location of the most adverse pressure gradient. With the aim of improving the predictive ability of the stall behavior for iced airfoils, simulations using the Detached Eddy Simulation (DES) approach were conducted in the second part of this numerical investigation. Three-dimensional DES computations were performed for a series of angles of attack around stall for the iced NACA 23012 and NLF 0414 airfoils. The simulations for both iced airfoils provided the maximum lift coefficients and stall behaviors qualitatively consistent with experiments.

  16. Influence of cross-sectional discontinuity on the damping characteristics of viscoelastically supported rectangular plates

    NASA Astrophysics Data System (ADS)

    Park, Jeongwon; Ahn, Sangkeun; Yoo, Ji Woo; Park, Junhong

    2016-09-01

    The geometry of a rectangular plate used in a structural application is an important design parameter that influences the vibrational response of the plate when it is subjected to an impact. In this study, the influence of a cross-sectional discontinuity on the vibration characteristics of viscoelastically supported plates was investigated. The discontinuity was induced at a specific location in the length-wise span. Experimental studies were performed to identify the effect of the discontinuity on the plate vibration response. The mode shapes and damping ratios of the plates with and without discontinuities in the cross-section were measured and compared. Forced vibration responses and modal properties were predicted using a numerical model. The variation in cross-sectional geometry was modeled to determine the changes in bending stiffness. The translational and rotational viscoelastic stiffnesses at the plate edges were used for modeling the vibration damping at the boundaries. This damping occurred at the contact surface between the plate and the fixtures. To investigate the effect of support stiffness on the vibration damping, flexural wave propagation analysis was performed with different boundary conditions. The ratio between the incident and reflected waves from the boundaries was predicted for flexural waves of different wavelengths. The predicted reflection ratios of the plate with and without the discontinuities were compared to the predicted loss factors using numerical analysis. The vibration energy dissipation at the viscoelastic supports was proportional to the measured modal damping.

  17. Preliminary Report on Laminar-Flow Airfoils and New Methods Adopted for Airfoil and Boundary-Layer Investigations

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N.

    1939-01-01

    Recent developments in airfoil-testing methods and fundamental air-flow investigations, as applied to airfoils, are discussed. Preliminary test results, obtained under conditions relatively free from stream turbulence and other disturbances, are presented. Suitable airfoils and airfoil-design principles were developed to take advantage of the unusually extensive laminar boundary layers that may be maintained under the improved testing conditions. The results are of interest mainly in range of below 6,000,000.

  18. 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.

  19. Airfoil noise in a uniform flow

    NASA Astrophysics Data System (ADS)

    Garcia, P.

    An experimental analysis was made of the noise radiated by a NACA 0012 airfoil in a uniform flow in the CEPRA 19 anechoic wind tunnel. The investigations concerned the estimate of the radiated noise from existing theories developed in particular by Chandiramani, Chase and Howe. They required experimental characterization of the pressure field induced by the turbulent boundary layer in the trailing edge region of the airfoil. This work is original in that it allows the noise to be predicted from wave number spectrum measurements made using a sensor array. The prediction is not limited to low frequencies as is the case for computations using the measured integral scales of Corcos. This approach was also applied to airfoils at an incidence.

  20. Feedback in separated flows over symmetric airfoils

    NASA Technical Reports Server (NTRS)

    Atassi, H. M.

    1984-01-01

    For a flow over an airfoil with laminar separation, a feedback cycle may exist whereby a Kelvin-Helmholtz instability wave emanating from the separation point on the airfoil surface grows along the shear layer and is diffracted as it interacts with the sharp trailing edge of the airfoil, causing acoustic radiation which, in turn, propagates upstream and regenerates the initial instability wave. The analysis is restricted to the high frequency limit. Solutions to the boundary-value problem are obtained using the slowly varying approximation and the method of matched asymptotic expansions. Resonant solutions exist for certain discrete values of the Reynolds and Strouhal numbers. The results are discussed and compared with available data.

  1. 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.

  2. Near-wall serpentine cooled turbine airfoil

    DOEpatents

    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. 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.

  4. 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.

  5. 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

  6. 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

  7. 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.

  8. 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.

  9. Community characteristics associated with where urgent care centers are located: a cross-sectional analysis

    PubMed Central

    Le, Sidney T; Hsia, Renee Y

    2016-01-01

    Objectives To determine the community characteristics associated with non-hospital-based urgent care centres wherever they are located. Design National cross-sectional study evaluating the association between non-hospital-based urgent care centers, and their demographic characteristics in a community, using descriptive statistics and multivariate logistic regressions. Setting Communities in the USA with non-hospital-based urgent care centers, as identified using a 2014 national database from the Urgent Care Association of America. Participants 31 022 communities encompassing 6898 urgent care centers across the USA. Primary and secondary outcome measures Presence of a non-hospital-based urgent care center within a community. Results Communities with non-hospital-based urgent care centers are urban (75.7% with vs 22.2% without; p<0.001 across rural urban commuting area levels), and are located in areas with higher income levels (38.6% in highest quartile with vs 22.3% without; p<0.001 across quartiles) and higher levels of private insurance (29.6% in highest quartile with vs 23.9% without; p<0.001 across quartiles). Conclusions While the growth of the urgent care industry may have other promising implications, policymakers should recognise that it may exacerbate disparities in access to acute care faced by poorer, uninsured patients, and may also have financial implications for providers that are providing overlapping services, such as emergency departments and primary care practices. PMID:27056591

  10. 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.

  11. 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

  12. Investigating Separated Shear Layer Development over an Airfoil with an Imbedded Microphone Array

    NASA Astrophysics Data System (ADS)

    Yarusevych, Serhiy; Gerakopulos, Ryan

    2010-11-01

    At low Reynolds numbers, laminar boundary layer separation on an airfoil often leads to deterioration in airfoil performance and noise emissions. The development of a separated shear layer is governed by laminar to turbulent transition, involving formation of coherent structures. This study highlights the design of a time-resolved surface pressure measurement system capable of estimating salient flow characteristics based on the analysis of surface pressure fluctuations. Wind tunnel experiments were performed for a symmetric NACA 0018 aluminum airfoil model equipped with a total of 95 static pressure taps and 24 microphones. Tests were performed for a range of angles of attack and Reynolds numbers to investigate two flow regimes common to airfoils operating at low Reynolds numbers, namely, flow separation without subsequent reattachment and separation bubble. Experimental results show that the microphones can be utilized to estimate the extent of the separation region and study the development of flow disturbances in the separated shear layer. Using hot wire measurements for validation, it is demonstrated that the microphones can detect the frequency signature of disturbances amplified in the separated shear layer. Further statistical analysis is employed to estimate such important characteristics of the disturbances and coherent structures as spanwise correlation, propagation speed, and phase.

  13. 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.

  14. 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.

  15. 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.

  16. Near-wall serpentine cooled turbine airfoil

    SciTech Connect

    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. Analysis of non-symmetrical flapping airfoils

    NASA Astrophysics Data System (ADS)

    Tay, W. B.; Lim, K. B.

    2009-08-01

    Simulations have been done to assess the lift, thrust and propulsive efficiency of different types of non-symmetrical airfoils under different flapping configurations. The variables involved are reduced frequency, Strouhal number, pitch amplitude and phase angle. In order to analyze the variables more efficiently, the design of experiments using the response surface methodology is applied. Results show that both the variables and shape of the airfoil have a profound effect on the lift, thrust, and efficiency. By using non-symmetrical airfoils, average lift coefficient as high as 2.23 can be obtained. The average thrust coefficient and efficiency also reach high values of 2.53 and 0.61, respectively. The lift production is highly dependent on the airfoil’s shape while thrust production is influenced more heavily by the variables. Efficiency falls somewhere in between. Two-factor interactions are found to exist among the variables. This shows that it is not sufficient to analyze each variable individually. Vorticity diagrams are analyzed to explain the results obtained. Overall, the S1020 airfoil is able to provide relatively good efficiency and at the same time generate high thrust and lift force. These results aid in the design of a better ornithopter’s wing.

  18. 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.

  19. Causal mechanisms in airfoil-circulation formation

    NASA Astrophysics Data System (ADS)

    Zhu, J. Y.; Liu, T. S.; Liu, L. Q.; Zou, S. F.; Wu, J. Z.

    2015-12-01

    In this paper, we trace the dynamic origin, rather than any kinematic interpretations, of lift in two-dimensional flow to the physical root of airfoil circulation. We show that the key causal process is the vorticity creation by tangent pressure gradient at the airfoil surface via no-slip condition, of which the theoretical basis has been given by Lighthill ["Introduction: Boundary layer theory," in Laminar Boundary Layers, edited by L. Rosenhead (Clarendon Press, 1963), pp. 46-113], which we further elaborate. This mechanism can be clearly revealed in terms of vorticity formulation but is hidden in conventional momentum formulation, and hence has long been missing in the history of one's efforts to understand lift. By a careful numerical simulation of the flow around a NACA-0012 airfoil, and using both Eulerian and Lagrangian descriptions, we illustrate the detailed transient process by which the airfoil gains its circulation and demonstrate the dominating role of relevant dynamical causal mechanisms at the boundary. In so doing, we find that the various statements for the establishment of Kutta condition in steady inviscid flow actually correspond to a sequence of events in unsteady viscous flow.

  20. Some experiments on autorotation of an airfoil

    NASA Technical Reports Server (NTRS)

    Ober, Shatswell

    1929-01-01

    These experiments show that the rate of auto rotation of a monoplane airfoil is reduced by sweepback, ceasing entirely when the sweepback is 30 degrees. In addition a very serious increase in rate and range of auto rotation with yaw is shown.

  1. Effects of surface roughness and vortex generators on the NACA 4415 airfoil

    SciTech Connect

    Reuss, R.L.; Hoffman, M.J.; Gregorek, G.M.

    1995-12-01

    Wind turbines in the field can be subjected to many and varying wind conditions, including high winds with rotor locked or with yaw excursions. In some cases the rotor blades may be subjected to unusually large angles of attack that possibly result in unexpected loads and deflections. To better understand loadings at unusual angles of attack, a wind tunnel test was performed. An 18-inch constant chord model of the NACA 4415 airfoil section was tested under two dimensional steady state conditions in the Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 7 x 10 Subsonic Wind Tunnel (7 x 10). The objective of these tests was to document section lift and moment characteristics under various model and air flow conditions. These included a normal angle of attack range of {minus}20{degree} to +40{degree}, an extended angle of attack range of {minus}60{degree} to +230{degree}, applications of leading edge grit roughness (LEGR), and use of vortex generators (VGs), all at chord Reynolds numbers as high as possible for the particular model configuration. To realistically satisfy these conditions the 7 x 10 offered a tunnel-height-to-model-chord ratio of 6.7, suggesting low interference effects even at the relatively high lift and drag conditions expected during the test. Significantly, it also provided chord Reynolds numbers up to 2.0 million. 167 figs., 13 tabs.

  2. Performance of advanced wind turbine airfoils with vortex generators

    SciTech Connect

    Wetzel, K.K.; Farokhi, S.

    1995-12-31

    The performance of the NREL S807 airfoil is experimentally determined via wind tunnel testing. The tests are conducted at Reynolds numbers of 0.5, 1.0, and 1.5{sm_bullet}10{sup 6}, with a clean surface, with two levels of leading edge surface roughness, and with surface roughness and large wishbone vortex generators. The results show that the S807 maximum lift coefficient drops with the application of leading edge surface roughness. The wishbone vortex generators are successful in restoring most of the loss in maximum lift coefficient at the cost of significant increase in profile drag at pre-stall angles of attack. The aerodynamic characteristics of the S807 with and without vortex generators are used as the input to the PROP93 and SEACC computer models to simulate the performance of an advanced wind turbine employing vortex generators. The results demonstrate that vortex generators could improve the performance of advanced wind turbines using the NREL airfoils by up to 4%.

  3. Closed-Loop Aerodynamic Flow Control of a Maneuvering Airfoil

    NASA Astrophysics Data System (ADS)

    Brzozowski, Daniel P.; Culp, John R.; Glezer, Ari

    2011-11-01

    The unsteady interaction between trailing edge aerodynamic flow control and airfoil motion in pitch and plunge is investigated in wind tunnel experiments using a 2-DOF traverse which enables application of time-dependent external torque and forces by servo motors. The global aerodynamic forces and moments are regulated by controlling vorticity generation and accumulation near the surface using hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and velocity measurements that are taken phase-locked to the commanded actuation waveform. The effect of the unsteady motion on the model-embedded flow control is assessed in unsteady several maneuvers. Circulation time history that is estimated from a PIV wake survey shows that the entire flow over the airfoil readjusts within about 1.5 TCONV, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales that are commensurate with the flow's convective time scale, and that the maneuver response is primarily limited by the inertia of the platform.

  4. 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.

  5. 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.

  6. Cyclists' helmet usage and characteristics in central and southern Malawi: a cross-sectional study.

    PubMed

    Kraemer, John D; Honermann, Brian J; Roffenbender, Jason S

    2012-01-01

    The objective of this study was to determine the prevalence of, and factors associated with, bicycle helmet usage in southern and central Malawi. This study was across-sectional observation of public behaviour. The urban and rural roadways in southern and central Malawi were studied during the dry season. In total, 1900 bicyclists were observed along the roadways of southern and central Malawi over a four-day period. Observer ascertainment of cyclists' helmet status, approximate age, sex and bicycle operator or passenger status were measured. Of the 1900 cyclists observed, no cyclist was identified as wearing a helmet (exact 95% CI: 0.0-0.2%). There was no variation by age, sex or operator/passenger status. Nearly, 91.5% of observed cyclists were males and 87.7% were operating the bicycle. The sizeable majority of male cyclists were classified as young adults from adolescence to 25 years old (47.2%) or adults over age 25 (44.9%); 7.9% of male cyclists were pre-adolescent. Passengers were more likely to be female than operators (39.1% versus 4.2%), though, even for passengers, a higher proportion were males than females (p < 0.001). In Malawi, helmet usage is so rare as to be non-existent. This suggests an opportunity for significant improvement. Based on the observed cyclists' characteristics, interventions should be targeted to adult and young adult males. PMID:22394127

  7. Cyclists' helmet usage and characteristics in central and southern Malawi: a cross-sectional study.

    PubMed

    Kraemer, John D; Honermann, Brian J; Roffenbender, Jason S

    2012-01-01

    The objective of this study was to determine the prevalence of, and factors associated with, bicycle helmet usage in southern and central Malawi. This study was across-sectional observation of public behaviour. The urban and rural roadways in southern and central Malawi were studied during the dry season. In total, 1900 bicyclists were observed along the roadways of southern and central Malawi over a four-day period. Observer ascertainment of cyclists' helmet status, approximate age, sex and bicycle operator or passenger status were measured. Of the 1900 cyclists observed, no cyclist was identified as wearing a helmet (exact 95% CI: 0.0-0.2%). There was no variation by age, sex or operator/passenger status. Nearly, 91.5% of observed cyclists were males and 87.7% were operating the bicycle. The sizeable majority of male cyclists were classified as young adults from adolescence to 25 years old (47.2%) or adults over age 25 (44.9%); 7.9% of male cyclists were pre-adolescent. Passengers were more likely to be female than operators (39.1% versus 4.2%), though, even for passengers, a higher proportion were males than females (p < 0.001). In Malawi, helmet usage is so rare as to be non-existent. This suggests an opportunity for significant improvement. Based on the observed cyclists' characteristics, interventions should be targeted to adult and young adult males.

  8. 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.

  9. 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.

  10. Comparative analysis of characteristic electron energy loss spectra and inelastic scattering cross-section spectra of Fe

    NASA Astrophysics Data System (ADS)

    Parshin, A. S.; Igumenov, A. Yu.; Mikhlin, Yu. L.; Pchelyakov, O. P.; Zhigalov, V. S.

    2016-05-01

    The inelastic electron scattering cross section spectra of Fe have been calculated based on experimental spectra of characteristic reflection electron energy loss as dependences of the product of the inelastic mean free path by the differential inelastic electron scattering cross section on the electron energy loss. It has been shown that the inelastic electron scattering cross-section spectra have certain advantages over the electron energy loss spectra in the analysis of the interaction of electrons with substance. The peaks of energy loss in the spectra of characteristic electron energy loss and inelastic electron scattering cross sections have been determined from the integral and differential spectra. It has been shown that the energy of the bulk plasmon is practically independent of the energy of primary electrons in the characteristic electron energy loss spectra and monotonically increases with increasing energy of primary electrons in the inelastic electron scattering cross-section spectra. The variation in the maximum energy of the inelastic electron scattering cross-section spectra is caused by the redistribution of intensities over the peaks of losses due to various excitations. The inelastic electron scattering cross-section spectra have been analyzed using the decomposition of the spectra into peaks of the energy loss. This method has been used for the quantitative estimation of the contributions from different energy loss processes to the inelastic electron scattering cross-section spectra of Fe and for the determination of the nature of the energy loss peaks.

  11. Validation of DYSTOOL for unsteady aerodynamic modeling of 2D airfoils

    NASA Astrophysics Data System (ADS)

    González, A.; Gomez-Iradi, S.; Munduate, X.

    2014-06-01

    From the point of view of wind turbine modeling, an important group of tools is based on blade element momentum (BEM) theory using 2D aerodynamic calculations on the blade elements. Due to the importance of this sectional computation of the blades, the National Renewable Wind Energy Center of Spain (CENER) developed DYSTOOL, an aerodynamic code for 2D airfoil modeling based on the Beddoes-Leishman model. The main focus here is related to the model parameters, whose values depend on the airfoil or the operating conditions. In this work, the values of the parameters are adjusted using available experimental or CFD data. The present document is mainly related to the validation of the results of DYSTOOL for 2D airfoils. The results of the computations have been compared with unsteady experimental data of the S809 and NACA0015 profiles. Some of the cases have also been modeled using the CFD code WMB (Wind Multi Block), within the framework of a collaboration with ACCIONA Windpower. The validation has been performed using pitch oscillations with different reduced frequencies, Reynolds numbers, amplitudes and mean angles of attack. The results have shown a good agreement using the methodology of adjustment for the value of the parameters. DYSTOOL have demonstrated to be a promising tool for 2D airfoil unsteady aerodynamic modeling.

  12. 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

  13. Oral polymorphonuclear neutrophil characteristics in relation to oral health: a cross-sectional, observational clinical study.

    PubMed

    Rijkschroeff, Patrick; Jansen, Ineke D C; van der Weijden, Fridus A; Keijser, Bart J F; Loos, Bruno G; Nicu, Elena A

    2016-01-01

    Polymorphonuclear neutrophils (PMNs) have a major role in the innate immune system. However, little is known about PMN contribution in relation to oral health. The objective of this study was to investigate the numbers and functional characteristics of oral PMNs (oPMNs) compared with circulatory PMNs (cPMNs). Oral rinse and venous blood samples were obtained from 268 systemically and orally healthy volunteers in a cross-sectional observational study. PMN counts, cell cycle analysis and cellular activation state were investigated. Also, reactive oxygen species (ROS) production was analyzed, with and without bacterial stimulation (Fusobacterium nucleatum). In males, 1.2 × 10(6)±1.0 × 10(6) oPMNs were collected, and showed a tendency to correlate with the levels of gingival bleeding (r=0.215, P=0.008). Comparable oPMNs counts were found among females (1.0 × 10(6)±0.7 × 10(6)). More late-stage apoptotic/necrotic cells were found among the oPMNs (53.1%) compared with the cPMNs (8.5%; P<0.001). Without additional stimulation, oPMNs were more activated than cPMNs, as indicated by higher expression of CD11b, CD63 and CD66b, and higher constitutive ROS levels (P<0.001). Notably, in response to bacterial stimulation, oPMNs released comparable ROS levels as cPMNs (P=0.042). In conclusion, this study provides data on viable oPMNs showing high levels of activation in orally and systemically healthy individuals, free of apparent caries lesions and periodontal disease. These data suggests that although the oPMNs are in a more mature stage of their life cycle compared with the cPMNs, oPMNs are still responsive to stimulation, which indicates their functional potential and possible contribution to a healthy oral ecosystem. PMID:27515277

  14. [Hydrochemical Characteristics and the Dissolved Inorganic Carbon Flux in Liuzhou Section of Liujiang Basin].

    PubMed

    Yuan, Ya-qiong; He, Shi-yi; Yu, Shi; Sun, Ping-an; Wang, Yan-xue; Wu, Zhao-yun; Li, Xin-gui; Xie, Ming-xian; Liu, Wen; Li, Rui; Zhang, Hua-sheng

    2015-07-01

    An important aspect of the current global change research is using river chemical composition to reveal the chemical weathering process and its effect of carbon sink. In this study, water samples were collected and analyzed 2 ~3 times per month from January to December in 2013. The hydrochemistry belonged to HCO3-Ca type. Ca+ and HCO3- were the main cation and anion, which reflected that the hydrochemical characteristics of river were mainly affected by the dissolution of carbonate rock. The concentration of main ions varied with the seasons, which reflected that the crest value occurred in winter, followed by those in autumn and spring, and the lowest value was observed in summer. Due to the interaction of effect of dilution and effect of C2, the seasonal variation of Ca2+ and HCO3- showed that the highest value was in autumn and the lowest value was in summer. The seasonal variation law of other ions should be attributed to the effect of dilution or agricultural activities or combined action of them. Both carbonic acid and sulfuric acid took part in the chemical weathering of carbonate rocks as evidenced by stoichiometric analysis. Besides, the δ34S of sulfate ion of the river waters (δ34S: from 7. 65 per thousand to 8. 55 per thousand) showed that SO2- was originated mainly from oxidation of sulfide minerals in ore deposits and acid rain. Chemical mass balance method was applied to estimate the proportion of HCO- coming from carbonate weathering by sulfuric acid. The result was 28. 26% . On this basis, the total carbon flux of carbon ( by CO2 calculation) in Liuzhou section calculated month by month was about 8. 95 x 10(5) t . a-1. What's more, the carbon flux showed a positive correlation with flow, which implied that the discharge of catchment was the main influencing factor of carbon flux rather than the HCO3- concentration. PMID:26489309

  15. Estimating HIV-1 Fitness Characteristics from Cross-Sectional Genotype Data

    PubMed Central

    Gopalakrishnan, Sathej; Montazeri, Hesam; Menz, Stephan; Beerenwinkel, Niko; Huisinga, Wilhelm

    2014-01-01

    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. PMID:25375675

  16. [Hydrochemical Characteristics and the Dissolved Inorganic Carbon Flux in Liuzhou Section of Liujiang Basin].

    PubMed

    Yuan, Ya-qiong; He, Shi-yi; Yu, Shi; Sun, Ping-an; Wang, Yan-xue; Wu, Zhao-yun; Li, Xin-gui; Xie, Ming-xian; Liu, Wen; Li, Rui; Zhang, Hua-sheng

    2015-07-01

    An important aspect of the current global change research is using river chemical composition to reveal the chemical weathering process and its effect of carbon sink. In this study, water samples were collected and analyzed 2 ~3 times per month from January to December in 2013. The hydrochemistry belonged to HCO3-Ca type. Ca+ and HCO3- were the main cation and anion, which reflected that the hydrochemical characteristics of river were mainly affected by the dissolution of carbonate rock. The concentration of main ions varied with the seasons, which reflected that the crest value occurred in winter, followed by those in autumn and spring, and the lowest value was observed in summer. Due to the interaction of effect of dilution and effect of C2, the seasonal variation of Ca2+ and HCO3- showed that the highest value was in autumn and the lowest value was in summer. The seasonal variation law of other ions should be attributed to the effect of dilution or agricultural activities or combined action of them. Both carbonic acid and sulfuric acid took part in the chemical weathering of carbonate rocks as evidenced by stoichiometric analysis. Besides, the δ34S of sulfate ion of the river waters (δ34S: from 7. 65 per thousand to 8. 55 per thousand) showed that SO2- was originated mainly from oxidation of sulfide minerals in ore deposits and acid rain. Chemical mass balance method was applied to estimate the proportion of HCO- coming from carbonate weathering by sulfuric acid. The result was 28. 26% . On this basis, the total carbon flux of carbon ( by CO2 calculation) in Liuzhou section calculated month by month was about 8. 95 x 10(5) t . a-1. What's more, the carbon flux showed a positive correlation with flow, which implied that the discharge of catchment was the main influencing factor of carbon flux rather than the HCO3- concentration.

  17. Exercise training characteristics in cardiac rehabilitation programmes: a cross-sectional survey of Australian practice

    PubMed Central

    Abell, Bridget; Glasziou, Paul; Briffa, Tom; Hoffmann, Tammy

    2016-01-01

    Introduction Exercise training is a core component of cardiac rehabilitation (CR), however, little information exists regarding the specific exercise interventions currently provided for coronary heart disease in Australian practice. We aimed to analyse the current status of exercise-based CR services across Australia. Design Cross-sectional survey. Methods Australian sites offering exercise-based CR were identified from publically available directories. All sites were invited by email to participate in an online Survey Monkey questionnaire between October 2014 and March 2015, with reminders via email and phone follow-up. Questions investigated the demographics and format of individual programmes, as well as specific exercise training characteristics. Results 297 eligible programmes were identified, with an 82% response rate. Most sites (82%) were based at hospital or outpatient centres, with home (15%), community (18%) or gym-based options (5%) less common. While CR was most often offered in a comprehensive format (72% of sites), the level of exercise intervention varied greatly among programmes. Most frequently, exercise was prescribed 1–2 times per week for 60 min over 7 weeks. Almost one-quarter (24%) had a sole practitioner supervising exercise, although the majority used a nurse/physiotherapist combination. Low to moderate exercise intensities were used in 60% of programmes, however, higher intensity prescriptions were not uncommon. Few sites (<6%) made use of technology, such as mobile phones or the internet, to deliver or support exercise training. Conclusions While advances have been made towards providing flexible and accessible exercise-based CR, much of Australia's service remains within traditional models of care. A continuing focus on service improvement and evidence-based care should, therefore, be considered a core aim of those providing exercise for CR in order to improve health service delivery and optimise outcomes for patients. PMID

  18. Patient characteristics and frequency of bodily distress syndrome in primary care: a cross-sectional study

    PubMed Central

    Budtz-Lilly, Anna; Vestergaard, Mogens; Fink, Per; Carlsen, Anders Helles; Rosendal, Marianne

    2015-01-01

    Background Bodily distress syndrome (BDS) is a newly proposed diagnosis of medically unexplained symptoms, which is based on empirical research in primary care. Aim To estimate the frequency of BDS in primary care and describe the characteristics of patients with BDS. Design and setting A cross-sectional study of primary care patients in urban and rural areas of Central Denmark Region. Method Data were obtained from GP one-page registration forms, patient questionnaires (including a checklist for BDS), and national registers. Results A total of 1356 primary care patients were included, of whom 230 patients (17.0%, 95% confidence intervals [CI] = 15.0 to 19.1) fulfilled the BDS criteria. BDS was more common among primary care patients aged 41–65 years (odds ratio [OR] = 1.9, 95% CI = 1.3 to 3.0) and was equally frequent among males and females (female sex, OR 0.9, 95% CI = 0.6 to 1.3). Patients with BDS were characterised by poor health-related quality of life (HRQOL) on the 12-item Short-Form Health Survey, that is, physical component summary scores <40 (OR 20.5, 95% CI = 12.9 to 32.4) and mental component summary scores <40 (OR 3.5, 95% CI = 2.2 to 5.6). Furthermore, patients with BDS were more likely to have high scores on the Symptom Checklist for anxiety (OR 2.2, 95% CI = 1.4 to 3.4) and depression (OR 5.1, 95% CI = 3.3 to 7.9), but regression analyses showed that mental morbidity did not account for the poor HRQOL. Conclusion BDS is common among primary care patients, and patients with BDS have a higher probability of poor HRQOL and mental health problems. PMID:26324499

  19. 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.

  20. 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.

  1. Turbine airfoil with laterally extending snubber having internal cooling system

    DOEpatents

    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.

  2. Quiet airfoils for small and large wind turbines

    DOEpatents

    Tangler, James L.; Somers, Dan L.

    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.

  3. 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.

  4. 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.

  5. S904 and S905 Airfoils: May 1998--January 1999

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A family of natural-laminar-flow airfoils, the S904 and S905, for cooling-tower fans has been designed and analyzed theoretically. The two primary objectives of high maximum lift, relatively insensitive to roughness, and low profile drag have been achieved. The constraint on the lift a zero angle of attack has not been satisfied. The constraints on the pitching moment and the airfoil thicknesses have essentially been satisfied. The airfoils should exhibit docile stalls.

  6. S829 Airfoil; Period of Performance: 1994--1995

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A 16%-thick, natural-laminar-flow airfoil, the S829, for the tip region of 20- to 40-meter-diameter, stall-regulated, horizontal-axis wind turbines has been designed and analyzed theoretically. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. The airfoil should exhibit a docile stall.

  7. Wind tunnel results of the low-speed NLF(1)-0414F airfoil

    NASA Technical Reports Server (NTRS)

    Murri, Daniel G.; Mcghee, Robert J.; Jordan, Frank L., Jr.; Davis, Patrick J.; Viken, Jeffrey K.

    1987-01-01

    The large performance gains predicted for the Natural Laminar Flow (NLF)(1)-0414F airfoil were demonstrated in two-dimensional airfoil tests and in wind tunnel tests conducted with a full scale modified Cessna 210. The performance gains result from maintaining extensive areas of natural laminar flow, and were verified by flight tests conducted with the modified Cessna. The lift, stability, and control characteristics of the Cessna were found to be essentially unchanged when boundary layer transition was fixed near the wing leading edge. These characteristics are very desirable from a safety and certification view where premature boundary layer transition (due to insect contamination, etc.) must be considered. The leading edge modifications were found to enhance the roll damping of the Cessna at the stall, and were therefore considered effective in improving the stall/departure resistance. Also, the modifications were found to be responsible for only minor performance penalties.

  8. 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.

  9. 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.

  10. 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.

  11. Saturation-free numerical scheme for computing the flow past a lattice of airfoils and the determination of separation points in a viscous fluid

    NASA Astrophysics Data System (ADS)

    Petrov, A. G.

    2011-07-01

    A numerical method for computing the potential flow past a lattice of airfoils is described. The problem is reduced to a linear integrodifferential equation on the lattice contour, which is then approximated by a linear system of equations with the help of specially derived quadrature formulas. The quadrature formulas exhibit exponential convergence in the number of points on an airfoil and have a simple analytical form. Due to its fast convergence and high accuracy, the method can be used to directly optimize the airfoils as based on any given integral characteristics. The shear stress distribution and the separation points are determined from the velocity distribution at the airfoil boundary calculated by solving the boundary layer equations. The method proposed is free of laborious grid generation procedures and does not involve difficulties associated with numerical viscosity at high Reynolds numbers.

  12. Turbine engine airfoil and platform assembly

    DOEpatents

    Campbell, Christian X.; James, Allister W.; Morrison, Jay A.

    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.

  13. 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.

  14. 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.

  15. 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.

  16. Experimental study of flow separation control on a low- Re airfoil using leading-edge protuberance method

    NASA Astrophysics Data System (ADS)

    Zhang, M. M.; Wang, G. F.; Xu, J. Z.

    2014-04-01

    An experimental study of flow separation control on a low- Re c airfoil was presently investigated using a newly developed leading-edge protuberance method, motivated by the improvement in the hydrodynamics of the giant humpback whale through its pectoral flippers. Deploying this method, the control effectiveness of the airfoil aerodynamics was fully evaluated using a three-component force balance, leading to an effectively impaired stall phenomenon and great improvement in the performances within the wide post-stall angle range (22°-80°). To understand the flow physics behind, the vorticity field, velocity field and boundary layer flow field over the airfoil suction side were examined using a particle image velocimetry and an oil-flow surface visualization system. It was found that the leading-edge protuberance method, more like low-profile vortex generator, effectively modified the flow pattern of the airfoil boundary layer through the chordwise and spanwise evolutions of the interacting streamwise vortices generated by protuberances, where the separation of the turbulent boundary layer dominated within the stall region and the rather strong attachment of the laminar boundary layer still existed within the post-stall region. The characteristics to manipulate the flow separation mode of the original airfoil indicated the possibility to further optimize the control performance by reasonably designing the layout of the protuberances.

  17. 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.

  18. Design of the LRP airfoil series using 2D CFD

    NASA Astrophysics Data System (ADS)

    Zahle, Frederik; Bak, Christian; Sørensen, Niels N.; Vronsky, Tomas; Gaudern, Nicholas

    2014-06-01

    This paper describes the design and wind tunnel testing of a high-Reynolds number, high lift airfoil series designed for wind turbines. The airfoils were designed using direct gradient- based numerical multi-point optimization based on a Bezier parameterization of the shape, coupled to the 2D Navier-Stokes flow solver EllipSys2D. The resulting airfoils, the LRP2-30 and LRP2-36, achieve both higher operational lift coefficients and higher lift to drag ratios compared to the equivalent FFA-W3 airfoils.

  19. Airfoil shape and thickness effects on transonic airloads and flutter

    NASA Technical Reports Server (NTRS)

    Bland, S. R.; Edwards, J. W.

    1983-01-01

    A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential equation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.

  20. Airfoil shape and thickness effects on transonic airloads and flutter

    NASA Technical Reports Server (NTRS)

    Bland, S. R.; Edwards, J. W.

    1983-01-01

    A transient pulse technique is used to obtain harmonic forces from a time-marching solution of the complete unsteady transonic small perturbation potential evaluation. The unsteady pressures and forces acting on a model of the NACA 64A010 conventional airfoil and the MBB A-3 supercritical airfoil over a range of Mach numbers are examined in detail. Flutter calculations at constant angle of attack show a similar flutter behavior for both airfoils, except for a boundary shift in Mach number associated with a corresponding Mach number shift in the unsteady aerodynamic forces. Differences in the static aeroelastic twist behavior for the two airfoils are significant.

  1. 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.

  2. 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.

  3. Numerical study of porous airfoils in transonic flow

    NASA Technical Reports Server (NTRS)

    Chen, C. L.; Chow, C. Y.; Holst, T. L.; Vandalsem, W. R.

    1985-01-01

    A numerical study was made to examine the effect of a porous surface on the aerodynamic performance of a transonic airfoil. The pressure jump across the normal shock wave on the upper surface of the airfoil was reduced by making the surface below the shock porous. The weakened shock is preceded by an oblique shock at the upstream end of the porous surface where air is blown out of the cavity. The lambda shock structure shown in the numerical result qualitatively agrees with that observed in the wind tunnel. According to the present analysis, the porous airfoil has a smaller drag and a higher lift than the solid airfoil.

  4. 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)

  5. 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.

  6. Control of Flow Separation Using Adaptive Airfoils

    NASA Technical Reports Server (NTRS)

    Chandrasekhara, M. S.; Wilder, M. C.; Carr, L. W.; Davis, Sanford S. (Technical Monitor)

    1996-01-01

    A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature. Interestingly, the flow over a nearly semicircular nosed airfoil was separated even at low angles.

  7. Profile design for an advanced-technology airfoil for general aviation aircraft

    NASA Technical Reports Server (NTRS)

    Welte, D.

    1978-01-01

    A profile from the NASA General Aviation Whitcomb series and NACA profiles are used as a starting point in designing an advanced airfoil for general aviation aircraft. Potential theory pressure distribution calculations, together with boundary layer calculations, permit a decrease in the null moment and an optimization of the lift characteristics of the wing. Trailing edge flap design is also improved. Wind tunnel tests are used to compare the conventional profiles, the NASA profile, and the improved design.

  8. Tests of N.A.C.A. airfoils in the variable density wind tunnel Series 44 and 64

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

    This note is one of a series covering an investigation of a number of related airfoils. It presents the results obtained from tests in the N.A.C.A. Variable Density Wind Tunnel of two groups of six airfoils each. One group, the 44 series, has a maximum mean camber of 4 percent of the chord at a position 0.4 of the chord behind the leading edge and the other group, the 64 series, has a maximum mean camber of 6 percent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being: 0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line form.

  9. Tests of N.A.C.A. airfoils in the variable-density wind tunnel Series 43 and 63

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

    This note is one of a series covering an investigation of a family of related airfoils. It gives in preliminary form the results obtained from tests in the N.A.C.A. Variable-Density Wind Tunnel of two groups of six airfoils each. One group, the 43 series, has a maximum mean camber of 4 per cent of the chord at a position 0.3 of the chord from the leading edge; the other group, the 63 series, has a maximum mean camber of 6 per cent of the chord at the same position. The members within each group differ only in maximum thickness, the maximum thickness/chord ratios being:0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results are analyzed with a view to indicating the variation of the aerodynamic characteristics with profile thickness for airfoils having a certain mean camber line.

  10. Self-sustained oscillations of a shock wave interacting with a boundary layer on a supercritical airfoil

    NASA Technical Reports Server (NTRS)

    Ventres, C. S.; Howe, M. S.

    1983-01-01

    A theory is proposed of the self-sustaining oscillations of a weak shock on an airfoil in steady, transonic flow. The interaction of the shock with the boundary layer on the airfoil produces displacement thickness fluctuations which convect downstream and generate sound by interaction with the trailing edge. A feedback loop is established when this sound impinges on the shock wave, resulting in the production of further fluctuations in the displacement thickness. The details are worked out for an idealized mean boundary layer velocity profile, but strong support for the basic hypotheses of the theory is provided by a comparison with recent experiments involving the generation of acoustic "tone bursts' by a supercritical airfoil section.

  11. 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.

  12. An experimental evaluation of the application of the Kirchhoff formulation for sound radiation from an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Brooks, T. F.

    1977-01-01

    The Kirchhoff integral formulation is evaluated for its effectiveness in quantitatively predicting the sound radiated from an oscillating airfoil whose chord length is comparable with the acoustic wavelength. A rigid airfoil section was oscillated at samll amplitude in a medium at rest to produce the sound field. Simultaneous amplitude and phase measurements were made of surface pressure and surface velocity distributions and the acoustic free field. Measured surface pressure and motion are used in applying the theory, and airfoil thickness and contour are taken into account. The result was that the theory overpredicted the sound pressure level by 2 to 5, depending on direction. Differences are also noted in the sound field phase behavior.

  13. 40 CFR Appendix Vi to Part 268 - Recommended Technologies To Achieve Deactivation of Characteristics in Section 268.42

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 28 2013-07-01 2013-07-01 false Recommended Technologies To Achieve Deactivation of Characteristics in Section 268.42 VI Appendix VI to Part 268 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) SOLID WASTES (CONTINUED) LAND DISPOSAL RESTRICTIONS Pt. 268, App....

  14. Characteristics and Challenges of Distance Learning Librarians: Report to the Distance Learning Section

    ERIC Educational Resources Information Center

    Brandt, John M.; Macaluso, Stephan J.

    2006-01-01

    Purpose: This study examined the professional and research needs of distance learning librarians in higher education in order to inform program development by the Association of College Libraries Distance Learning Section. Methodology: 270 members of the section completed a 23 item online questionnaire on distance learning librarians' work…

  15. 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.

  16. 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.

  17. Development of heat flux sensors in turbine airfoils

    NASA Technical Reports Server (NTRS)

    Atkinson, W. H.; Strange, R. R.

    1984-01-01

    The objective is to develop heat flux sensors suitable for use on turbine airfoils and to verify the operation of the heat flux measurement techniques through laboratory experiments. The requirements for a program to investigate the measurement of heat flux on airfoils in areas of strong non-one-dimensional flow were also identified.

  18. Sealing apparatus for airfoils of gas turbine engines

    SciTech Connect

    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.

  19. S822 and S823 Airfoils: October 1992--December 1993

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A family of thick airfoils for 3- to 10-meter, stall-regulated, horizontal-axis wind turbines, the S822 and S823, has been designed and analyzed theoretically. The primary objectives of restrained maximum lift, insensitive to roughness, and low profile have been achieved. The constraints on the pitching moments and airfoil thicknesses have been satisfied.

  20. 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.

  1. 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.

  2. URANS simulations of separated flow with stall cells over an NREL S826 airfoil

    NASA Astrophysics Data System (ADS)

    Sarlak, H.; Nishino, T.; Sørensen, J. N.

    2016-06-01

    A series of wind tunnel measurements and oil flow visualization was recently carried out at the Technical University of Denmark in order to investigate flow characteristics over a 14% thick NREL S826 airfoil at low Reynolds numbers. This paper aims at presenting numerical simulations of the same airfoil using unsteady Reynolds-averaged Navier-Stokes (URANS) approach. Results of the simulations are demonstrated in terms of mean flow velocity, lift and drag, as well as pressure distribution, and validated against available experimental data. The simulations are carried out with a wide computational domain (with a span-to-chord ratio of 5) and it is illustrated that the URANS approach is capable of predicting 3D spanwise structures, known as stall cells.

  3. 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.

  4. Numerical and experimental study of blowing jet on a high lift airfoil

    NASA Astrophysics Data System (ADS)

    Bobonea, A.; Pricop, M. V.

    2013-10-01

    Active manipulation of separated flows over airfoils at moderate and high angles of attack in order to improve efficiency or performance has been the focus of a number of numerical and experimental investigations for many years. One of the main methods used in active flow control is the usage of blowing devices with constant and pulsed blowing. Through CFD simulation over a 2D high-lift airfoil, this study is trying to highlight the impact of pulsed blowing over its aerodynamic characteristics. The available wind tunnel data from INCAS low speed facility are also beneficial for the validation of the numerical analysis. This study intends to analyze the impact of the blowing jet velocity and slot geometry on the efficiency of an active flow control.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. Dynamic Stall Measurements and Computations for a VR-12 Airfoil with a Variable Droop Leading Edge

    NASA Technical Reports Server (NTRS)

    Martin, P. B.; McAlister, K. W.; Chandrasekhara, M. S.; Geissler, W.

    2003-01-01

    High density-altitude operations of helicopters with advanced performance and maneuver capabilities have lead to fundamental research on active high-lift system concepts for rotor blades. The requirement for this type of system was to improve the sectional lift-to-drag ratio by alleviating dynamic stall on the retreating blade while simultaneously reducing the transonic drag rise of the advancing blade. Both measured and computational results showed that a Variable Droop Leading Edge (VDLE) airfoil is a viable concept for application to a rotor high-lift system. Results are presented for a series of 2D compressible dynamic stall wind tunnel tests with supporting CFD results for selected test cases. These measurements and computations show a dramatic decrease in the drag and pitching moment associated with severe dynamic stall when the VDLE concept is applied to the Boeing VR-12 airfoil. Test results also show an elimination of the negative pitch damping observed in the baseline moment hysteresis curves.

  11. Unsteady flow past an airfoil pitched at constant rate

    NASA Technical Reports Server (NTRS)

    Lourenco, L.; Vandommelen, L.; Shib, C.; Krothapalli, A.

    1992-01-01

    The unsteady flow past a NACA 0012 airfoil that is undertaking a constant-rate pitching up motion is investigated experimentally by the PIDV technique in a water towing tank. The Reynolds number is 5000, based upon the airfoil's chord and the free-stream velocity. The airfoil is pitching impulsively from 0 to 30 deg. with a dimensionless pitch rate alpha of 0.131. Instantaneous velocity and associated vorticity data have been acquired over the entire flow field. The primary vortex dominates the flow behavior after it separates from the leading edge of the airfoil. Complete stall emerges after this vortex detaches from the airfoil and triggers the shedding of a counter-rotating vortex near the trailing edge. A parallel computational study using the discrete vortex, random walk approximation has also been conducted. In general, the computational results agree very well with the experiment.

  12. 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.

  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. Time-domain inflow boundary condition for turbulence-airfoil interaction noise prediction using synthetic turbulence modeling

    NASA Astrophysics Data System (ADS)

    Kim, Daehwan; Heo, Seung; Cheong, Cheolung

    2015-03-01

    The present paper deals with development of the synthetic turbulence inflow boundary condition (STIBC) to predict inflow broadband noise generated by interaction between turbulence and an airfoil/a cascade of airfoils in the time-domain. The STIBC is derived by combining inflow boundary conditions that have been successfully applied in external and internal computational aeroacoustics (CAA) simulations with a synthetic turbulence model. The random particle mesh (RPM) method based on a digital filter is used as the synthetic turbulence model. Gaussian and Liepmann spectra are used to define the filters for turbulence energy spectra. The linearized Euler equations are used as governing equations to evaluate the suitability of the STIBC in time-domain CAA simulations. First, the velocity correlations and energy spectra of the synthesized turbulent velocities are compared with analytic ones. The comparison results reveal that the STIBC can reproduce a turbulent velocity field satisfying the required statistical characteristics of turbulence. Particularly, the Liepmann filter representing a non-Gaussian filter is shown to be effectively described by superposing the Gaussian filters. Each Gaussian filter has a different turbulent kinetic energy and integral length scale. Second, two inflow noise problems are numerically solved using the STIBC: the turbulence-airfoil interaction and the turbulence-a cascade of airfoils interaction problems. The power spectrum of noise due to an isolated flat plate airfoil interacting with incident turbulence is predicted, and its result is successfully validated against Amiet's analytic model (Amiet, 1975) [4]. The prediction results of the upstream and downstream acoustic power spectra from a cascade of flat plates are then compared with Cheong's analytic model (Cheong et al., 2006) [30]. These comparisons are also in excellent agreement. On the basis of these illustrative computation results, the STIBC is expected to be applied to

  15. 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…

  16. Tonal noise production from a wall-mounted finite airfoil

    NASA Astrophysics Data System (ADS)

    Moreau, Danielle J.; Doolan, Con J.

    2016-02-01

    This study is concerned with the flow-induced noise of a smooth wall-mounted finite airfoil with flat ended tip and natural boundary layer transition. Far-field noise measurements have been taken at a single observer location and with a microphone array in the Virginia Tech Stability Wind Tunnel for a wall-mounted finite airfoil with aspect ratios of L / C = 1 - 3, at a range of Reynolds numbers (ReC = 7.9 ×105 - 1.6 ×106, based on chord) and geometric angles of attack (α = 0 - 6 °). At these Reynolds numbers, the wall-mounted finite airfoil produces a broadband noise contribution with a number of discrete equispaced tones at non-zero angles of attack. Spectral data are also presented for the noise produced due to three-dimensional vortex flow near the airfoil tip and wall junction to show the contributions of these flow features to airfoil noise generation. Tonal noise production is linked to the presence of a transitional flow state to the trailing edge and an accompanying region of mildly separated flow on the pressure surface. The separated flow region and tonal noise source location shift along the airfoil trailing edge towards the free-end region with increasing geometric angle of attack due to the influence of the tip flow field over the airfoil span. Tonal envelopes defining the operating conditions for tonal noise production from a wall-mounted finite airfoil are derived and show that the domain of tonal noise production differs significantly from that of a two-dimensional airfoil. Tonal noise production shifts to lower Reynolds numbers and higher geometric angles of attack as airfoil aspect ratio is reduced.

  17. Characteristics and Performance of Students in an Online Section of Business Statistics

    ERIC Educational Resources Information Center

    Dutton, John; Dutton, Marilyn

    2005-01-01

    We compare students in online and lecture sections of a business statistics class taught simultaneously by the same instructor using the same content, assignments, and exams in the fall of 2001. Student data are based on class grades, registration records, and two surveys. The surveys asked for information on preparedness, reasons for section…

  18. [Characteristic features of systemic hemodynamics during cesarean section under general anesthesia with ketamine].

    PubMed

    Moiseev, V N

    1983-02-01

    On the basis of a comparative investigation of the central hemodynamics by the method of integrative rheography of the body in two groups of women during the operation of cesarean section under general anesthesia with ether or ketamin the author makes a conclusion that ketamin is a good drug for anesthesia in urgent surgical situations.

  19. Method of characteristics design of a supersonic wind tunnel nozzle with square cross-section, supplement 3

    NASA Technical Reports Server (NTRS)

    Schneider, Steven P.

    1993-01-01

    Nozzle design codes developed earlier were modified and used in order to design a supersonic wind tunnel nozzle with square cross-sections. As part of the design process, a post-processing code that generates square nozzles from the output of the Sivells code was written. The method is based on the results of an axisymmetric method of characteristics code and an axisymmetric boundary layer code. These design codes are documented.

  20. Running characteristics of the magnetically levitated train in a curved track section

    SciTech Connect

    Ohashi, S.; Ohsaki, H.; Masada, E.

    1997-09-01

    A three-dimensional running simulation program of the train motion which includes both electromagnetic and mechanical systems has been developed. For verification of running stability, the motion of the Maglev train with four bogies and three cars running into the curved section was calculated. Because the spring strength of the secondary suspension between bogies and cabins was much smaller than that of the electrodynamic suspension, interaction among bogies was small. The lateral displacement reached 0.045m, which is about 50% of the lateral air gap. Therefore train can run without touching the wall of the guideway. As this lateral air gap caused roll torque, total roll angle of the bogie became 12.5% larger than the cant angle of the curved section.

  1. Individual characteristics and student’s engagement in scientific research: a cross-sectional study

    PubMed Central

    2012-01-01

    Background In light of the increasing recognition of the importance of physician scientists, and given the association between undergraduate research experiences with future scientific activity, it is important to identify and understand variables related to undergraduate students’ decision to engage in scientific research activities. The present study assessed the influence of individual characteristics, including personality traits and socio-demographic characteristics, on voluntary engagement in scientific research of undergraduate medical students. Methods For this study, all undergraduate students and alumni of the School of Health Sciences in Minho, Portugal were invited to participate in a survey about voluntary engagement in scientific research activities. Data were available on socio-demographic, personality and university admission variables, as part of an ongoing longitudinal study. A regression model was used to compare (1) engaged with (2) not engaged students. A classification and regression tree model was used to compare students engaged in (3) elective curricular research (4) and extra-curricular research. Results A total of 466 students (88%) answered the survey. A complete set of data was available for 435 students (83%). Higher scores in admission grade point average and the personality dimensions of “openness to experience” and “conscientiousness” increased chances of engagement. Higher “extraversion” scores had the opposite effect. Male undergraduate students were two times more likely than females to engage in curricular elective scientific research and were also more likely to engage in extra-curricular research activities. Conclusions This study demonstrated that student’ grade point average and individual characteristics, like gender, openness and consciousness have a unique and statistically significant contribution to students’ involvement in undergraduate scientific research activities. PMID:23066758

  2. Airfoil for a gas turbine engine

    SciTech Connect

    Liang, George

    2011-05-24

    An airfoil is provided for a turbine of a gas turbine engine. The airfoil comprises: an outer structure comprising a first wall including a leading edge, a trailing edge, a pressure side, and a suction side; an inner structure comprising a second wall spaced from the first wall and at least one intermediate wall; and structure extending between the first and second walls so as to define first and second gaps between the first and second walls. The second wall and the at least one intermediate wall define at least one pressure side supply cavity and at least one suction side supply cavity. The second wall may include at least one first opening near the leading edge of the first wall. The first opening may extend from the at least one pressure side supply cavity to the first gap. The second wall may further comprise at least one second opening near the trailing edge of the outer structure. The second opening may extend from the at least one suction side supply cavity to the second gap. The first wall may comprise at least one first exit opening extending from the first gap through the pressure side of the first wall and at least one second exit opening extending from the second gap through the suction side of the second wall.

  3. Control of Flow Separation Using Adaptive Airfoils

    NASA Technical Reports Server (NTRS)

    Chandrasekhara, M. S.; Wilder, M. C.; Carr, L. W.; Davis, Sanford S. (Technical Monitor)

    1996-01-01

    A novel way of controlling flow separation is reported. The approach involves using an adaptive airfoil geometry that changes its leading edge shape to adjust to the instantaneous flow at high angles of attack such that the flow over it remains attached. In particular, a baseline NACA 0012 airfoil, whose leading edge curvature could be changed dynamically by 400% was tested under quasi-steady compressible flow conditions. A mechanical drive system was used to produce a rounded leading edge to reduce the strong local flow acceleration around its nose and thus reduce the strong adverse pressure gradient that follows such a rapid acceleration. Tests in steady flow showed that at M = 0.3, the flow separated at about 14 deg. angle of attack for the NACA 0012 profile but could be kept attached up to an angle of about 18 deg by changing the nose curvature. No significant hysteresis effects were observed; the flow could be made to reattach from its separated state at high angles by changing the leading edge curvature.

  4. Characteristics of the Langley 8-foot Transonic Tunnel with Slotted Test Section

    NASA Technical Reports Server (NTRS)

    Wright, Ray H; Ritchie, Virgil S; Pearson, Albin O

    1958-01-01

    A large wind tunnel, approximately 8 feet in diameter, has been converted to transonic operation by means of slots in the boundary extending in the direction of flow. The usefulness of such a slotted wind tunnel, already known with respect to the reduction of the subsonic blockage interference and the production of continuously variable supersonic flows, has been augmented by devising a slot shape with which a supersonic test region with excellent flow quality could be produced. Experimental locations of detached shock waves ahead of axially symmetric bodies at low supersonic speeds in the slotted test section agreed satisfactorily with predictions obtained by use of existing approximate methods.

  5. Unsteady Newton-Busemann flow theory. I - Airfoils

    NASA Technical Reports Server (NTRS)

    Hui, W. H.; Tobak, M.

    1981-01-01

    Newtonian flow theory for unsteady flow at very high Mach numbers is completed by the addition of a centrifugal force correction to the impact pressures. The correction term is the unsteady counterpart of Busemann's centrifugal force correction to impact pressures in steady flow. For airfoils of arbitary shape, exact formulas for the unsteady pressure and stiffness and damping-in-pitch derivatives are obtained in closed form, which require only numerical quadratures of terms involving the airfoil shape. They are applicable to airfoils of arbitrary thickness having sharp or blunt leading edges. For wedges and thin airfoils these formulas are greatly simplified, and it is proved that the pitching motions of thin airfoils of convex shape and of wedges of arbitrary thickness are always dynamically stable according to Newton-Busemann theory. Leading-edge bluntness is shown to have a favorable effect on the dynamic stability; on the other hand, airfoils of concave shape tend toward dynamic instability over a range of axis positions if the surface curvature exceeds a certain limit. As a byproduct, it is also shown that a pressure formula recently given by Barron and Mandl for unsteady Newtonian flow over a pitching power-law shaped airfoil is erroneous and that their conclusion regarding the effect of pivot position on the dynamic stability is misleading.

  6. On the Theory of the Unsteady Motion of an Airfoil

    NASA Technical Reports Server (NTRS)

    Sedov, L. I.

    1947-01-01

    The paper presents a systematical analysis of the problem of the determination of the unsteady motion about an airfoil moving in an infinite fluid that contains a system of vortices and the determination of the hydrodynamical forces acting on the airfoil. The hydrodynamical problem is reduced to the determination of the function f (xi) which transforms conformally the external region of the airfoil into the interior of a circle. The proposed methods of determining the irrotational motion of a fluid that is produced by any motion of the airfoil are especially simple and effective if the function f (xi) is rational. As an example the flow is determined for the case of an arbitrary motion of an airfoil of the Joukowsky type. The formulas obtained for the determination of the hydrodynamical forces by means of contour integration are similar to those given by S. Chaplygin. These formulas are used to determine the force acting on the airfoil in the cases where the unsteady motion is potential throughout and the circulation about the airfoil is constant and also when the fluid contains a system of vortices. A full discussion is given of the concept of virtual masses together with practical formulas for computing the virtual mass coefficients.

  7. Boundary layer separation and reattachment detection on airfoils by thermal flow sensors.

    PubMed

    Sturm, Hannes; Dumstorff, Gerrit; Busche, Peter; Westermann, Dieter; Lang, Walter

    2012-10-24

    A sensor concept for detection of boundary layer separation (flow separation, stall) and reattachment on airfoils is introduced in this paper. Boundary layer separation and reattachment are phenomena of fluid mechanics showing characteristics of extinction and even inversion of the flow velocity on an overflowed surface. The flow sensor used in this work is able to measure the flow velocity in terms of direction and quantity at the sensor's position and expected to determine those specific flow conditions. Therefore, an array of thermal flow sensors has been integrated (flush-mounted) on an airfoil and placed in a wind tunnel for measurement. Sensor signals have been recorded at different wind speeds and angles of attack for different positions on the airfoil. The sensors used here are based on the change of temperature distribution on a membrane (calorimetric principle). Thermopiles are used as temperature sensors in this approach offering a baseline free sensor signal, which is favorable for measurements at zero flow. Measurement results show clear separation points (zero flow) and even negative flow values (back flow) for all sensor positions. In addition to standard silicon-based flow sensors, a polymer-based flexible approach has been tested showing similar results.

  8. Blade Design Trade-Offs Using Low-Lift Airfoils for Stall-Regulated HAWTs

    SciTech Connect

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

    1999-04-08

    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 lift 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 reducing the maximum lift coefficient below the upper limit considered in this study increases the cost of energy independently of the wind regime. As a consequence, higher maximum lift coefficient airfoils for the tip-region of the blade become more desirable as machine size increases, as long as they provide gentle stall characteristics. The conclusions are applicable to large wind turbines that use passive or active stall to regulate peak power.

  9. Boundary Layer Separation and Reattachment Detection on Airfoils by Thermal Flow Sensors

    PubMed Central

    Sturm, Hannes; Dumstorff, Gerrit; Busche, Peter; Westermann, Dieter; Lang, Walter

    2012-01-01

    A sensor concept for detection of boundary layer separation (flow separation, stall) and reattachment on airfoils is introduced in this paper. Boundary layer separation and reattachment are phenomena of fluid mechanics showing characteristics of extinction and even inversion of the flow velocity on an overflowed surface. The flow sensor used in this work is able to measure the flow velocity in terms of direction and quantity at the sensor's position and expected to determine those specific flow conditions. Therefore, an array of thermal flow sensors has been integrated (flush-mounted) on an airfoil and placed in a wind tunnel for measurement. Sensor signals have been recorded at different wind speeds and angles of attack for different positions on the airfoil. The sensors used here are based on the change of temperature distribution on a membrane (calorimetric principle). Thermopiles are used as temperature sensors in this approach offering a baseline free sensor signal, which is favorable for measurements at zero flow. Measurement results show clear separation points (zero flow) and even negative flow values (back flow) for all sensor positions. In addition to standard silicon-based flow sensors, a polymer-based flexible approach has been tested showing similar results. PMID:23202160

  10. Controlled aeroelastic response and airfoil shaping using adaptive materials and integrated systems

    NASA Astrophysics Data System (ADS)

    Pinkerton, Jennifer L.; McGowan, Anna-Maria R.; Moses, Robert W.; Scott, Robert C.; Heeg, Jennifer

    1996-05-01

    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 charcteristics (ATTACH) project. The PARTI program demonstrated active flutter control and significant reductions 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. The 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-unimorph piezoelectric driver and sensor (THUNDER) wafers to control airfoil aerodynamic characteristics was investigated. Plans for future applications are also discussed.

  11. Computer-aided roll pass design in rolling of airfoil shapes

    NASA Technical Reports Server (NTRS)

    Akgerman, N.; Lahoti, G. D.; Altan, T.

    1980-01-01

    This paper describes two computer-aided design (CAD) programs developed for modeling the shape rolling process for airfoil sections. The first program, SHPROL, uses a modular upper-bound method of analysis and predicts the lateral spread, elongation, and roll torque. The second program, ROLPAS, predicts the stresses, roll separating force, the roll torque and the details of metal flow by simulating the rolling process, using the slab method of analysis. ROLPAS is an interactive program; it offers graphic display capabilities and allows the user to interact with the computer via a keyboard, CRT, and a light pen. The accuracy of the computerized models was evaluated by (a) rolling a selected airfoil shape at room temperature from 1018 steel and isothermally at high temperature from Ti-6Al-4V, and (b) comparing the experimental results with computer predictions. The comparisons indicated that the CAD systems, described here, are useful for practical engineering purposes and can be utilized in roll pass design and analysis for airfoil and similar shapes.

  12. Accurate load prediction by BEM with airfoil data from 3D RANS simulations

    NASA Astrophysics Data System (ADS)

    Schneider, Marc S.; Nitzsche, Jens; Hennings, Holger

    2016-09-01

    In this paper, two methods for the extraction of airfoil coefficients from 3D CFD simulations of a wind turbine rotor are investigated, and these coefficients are used to improve the load prediction of a BEM code. The coefficients are extracted from a number of steady RANS simulations, using either averaging of velocities in annular sections, or an inverse BEM approach for determination of the induction factors in the rotor plane. It is shown that these 3D rotor polars are able to capture the rotational augmentation at the inner part of the blade as well as the load reduction by 3D effects close to the blade tip. They are used as input to a simple BEM code and the results of this BEM with 3D rotor polars are compared to the predictions of BEM with 2D airfoil coefficients plus common empirical corrections for stall delay and tip loss. While BEM with 2D airfoil coefficients produces a very different radial distribution of loads than the RANS simulation, the BEM with 3D rotor polars manages to reproduce the loads from RANS very accurately for a variety of load cases, as long as the blade pitch angle is not too different from the cases from which the polars were extracted.

  13. A harmonic analysis method for unsteady transonic flow and its application to the flutter of airfoils

    NASA Technical Reports Server (NTRS)

    Ehlers, F. E.; Weatherill, W. H.

    1982-01-01

    A finite difference method for solving the unsteady transonic flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady differential equation for small disturbances. The differential equation for the unsteady velocity potential is linear with spatially varying coefficients and with the time variable eliminated by assuming harmonic motion. A study is presented of the shock motion associated with an oscillating airfoil and its representation by the harmonic procedure. The effects of the shock motion and the resulting pressure pulse are shown to be included in the harmonic pressure distributions and the corresponding generalized forces. Analytical and experimental pressure distributions for the NACA 64A010 airfoil are compared for Mach numbers of 0.75, 0.80 and 0.842. A typical section, two-degree-of-freedom flutter analysis of a NACA 64A010 airfoil is performed. The results show a sharp transonic bucket in one case and abrupt changes in instability modes.

  14. An experimental study of static and oscillating rotor blade sections in reverse flow

    NASA Astrophysics Data System (ADS)

    Lind, Andrew Hume

    The rotorcraft community has a growing interest in the development of high-speed helicopters to replace outdated fleets. One barrier to the design of such helicopters is the lack of understanding of the aerodynamic behavior of retreating rotor blades in the reverse flow region. This work considers two fundamental models of this complex unsteady flow regime: static and oscillating (i.e., pitching) airfoils in reverse flow. Wind tunnel tests have been performed at the University of Maryland (UMD) and the United States Naval Academy (USNA). Four rotor blade sections are considered: two featuring a sharp geometric trailing edge (NACA 0012 and NACA 0024) and two featuring a blunt geometric trailing edge (ellipse and cambered ellipse). Static airfoil experiments were performed at angles of attack through 180 deg and Reynolds numbers up to one million, representative of the conditions found in the reverse flow region of a full-scale high-speed helicopter. Time-resolved velocity field measurements were used to identify three unsteady flow regimes: slender body vortex shedding, turbulent wake, and deep stall vortex shedding. Unsteady airloads were measured in these three regimes using unsteady pressure transducers. The magnitude of the unsteady airloads is high in the turbulent wake regime when the separated shear layer is close to the airfoil surface and in deep stall due to periodic vortex-induced flow. Oscillating airfoil experiments were performed on a NACA 0012 and cambered ellipse to investigate reverse flow dynamic stall characteristics by modeling cyclic pitching kinematics. The parameter space spanned three Reynolds numbers (165,000; 330,000; and 500,000), five reduced frequencies between 0.100 and 0.511, three mean pitch angles (5,10, and 15 deg), and two pitch amplitudes (5 deg and 10 deg). The sharp aerodynamic leading edge of the NACA 0012 airfoil forces flow separation resulting in deep dynamic stall. The number of associated vortex structures depends strongly

  15. Characteristics of injuries caused by paragliding accidents: A cross-sectional study

    PubMed Central

    Canbek, Umut; İmerci, Ahmet; Akgün, Ulaş; Yeşil, Murat; Aydin, Ali; Balci, Yasemin

    2015-01-01

    BACKGROUND: This study was undertaken to analyze the characteristics and risk factors relating to fatalities and injuries caused by paragliding. METHODS: The judicial examination reports and hospital documents of 82 patients traumatized in 64 accidents during 242 355 paragliding jumps between August 2004 and September 2011 were analyzed. RESULTS: In these accidents, 18 of the 82 patients lost their lives. In the patients with a confirmed cause of accident, most of them were involved with multiple fractures and internal organ injuries (n=8, 44.4%). CONCLUSION: We investigated the incidence of paragliding injuries, the types of the injuries, and the severity of affected anatomical regions. The findings are significant for the prevention of paragliding injuries and future research. PMID:26401185

  16. Characteristics and implications of prolonged fusicoccin-induced growth of Avena coleoptile sections

    NASA Technical Reports Server (NTRS)

    Cleland, R. E.

    1994-01-01

    A study has been made of the prolonged growth of Avena coleoptile sections in response to fusicoccin (FC), a phytotoxin that promotes apoplastic acidification. The final amount of FC-induced growth is a function of the FC concentration. Removal of the epidermis speeds up the initial rate of elongation and shortens the duration of the response, without affecting the total amount of extension. A suboptimal FC concentration (7 x 10(-8) M) which induces the same rate of proton excretion as does optimal indoleacetic acid (IAA) (1 x 10(-5) M), causes elongation which is 60-75% of that induced by IAA in 4 h or 50-65% in 7 h. This suggests that acid-induced extension could make a major contribution to auxin-induced growth for at least 7 h.

  17. Discontinuous Galerkin methodology for Large-Eddy Simulations of wind turbine airfoils

    NASA Astrophysics Data System (ADS)

    Frére, A.; Sørensen, N. N.; Hillewaert, K.; Winckelmans, G.

    2016-09-01

    This paper aims at evaluating the potential of the Discontinuous Galerkin (DG) methodology for Large-Eddy Simulation (LES) of wind turbine airfoils. The DG method has shown high accuracy, excellent scalability and capacity to handle unstructured meshes. It is however not used in the wind energy sector yet. The present study aims at evaluating this methodology on an application which is relevant for that sector and focuses on blade section aerodynamics characterization. To be pertinent for large wind turbines, the simulations would need to be at low Mach numbers (M ≤ 0.3) where compressible approaches are often limited and at large Reynolds numbers (Re ≥ 106) where wall-resolved LES is still unaffordable. At these high Re, a wall-modeled LES (WMLES) approach is thus required. In order to first validate the LES methodology, before the WMLES approach, this study presents airfoil flow simulations at low and high Reynolds numbers and compares the results to state-of-the-art models used in industry, namely the panel method (XFOIL with boundary layer modeling) and Reynolds Averaged Navier-Stokes (RANS). At low Reynolds number (Re = 6 x 104), involving laminar boundary layer separation and transition in the detached shear layer, the Eppler 387 airfoil is studied at two angles of attack. The LES results agree slightly better with the experimental chordwise pressure distribution than both XFOIL and RANS results. At high Reynolds number (Re = 1.64 x 106), the NACA4412 airfoil is studied close to stall condition. In this case, although the wall model approach used for the WMLES is very basic and not supposed to handle separation nor adverse pressure gradients, all three methods provide equivalent accuracy on averaged quantities. The present work is hence considered as a strong step forward in the use of LES at high Reynolds numbers.

  18. Design and experimental results for the S805 airfoil

    SciTech Connect

    Somers, D.M.

    1997-01-01

    An airfoil for horizontal-axis wind-turbine applications, the S805, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.

  19. Design and experimental results for the S809 airfoil

    SciTech Connect

    Somers, D M

    1997-01-01

    A 21-percent-thick, laminar-flow airfoil, the S809, for horizontal-axis wind-turbine applications, has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of restrained maximum lift, insensitive to roughness, and low profile drag have been achieved. The airfoil also exhibits a docile stall. Comparisons of the theoretical and experimental results show good agreement. Comparisons with other airfoils illustrate the restrained maximum lift coefficient as well as the lower profile-drag coefficients, thus confirming the achievement of the primary objectives.

  20. Experiences with optimizing airfoil shapes for maximum lift over drag

    NASA Technical Reports Server (NTRS)

    Doria, Michael L.

    1991-01-01

    The goal was to find airfoil shapes which maximize the ratio of lift over drag for given flow conditions. For a fixed Mach number, Reynolds number, and angle of attack, the lift and drag depend only on the airfoil shape. This then becomes a problem in optimization: find the shape which leads to a maximum value of lift over drag. The optimization was carried out using a self contained computer code for finding the minimum of a function subject to constraints. To find the lift and drag for each airfoil shape, a flow solution has to be obtained. This was done using a two dimensional Navier-Stokes code.

  1. Broadband Noise Predictions for an Airfoil in a Turbulent Stream

    NASA Technical Reports Server (NTRS)

    Casper, J.; Farassat, F.; Mish, P. F.; Devenport, W. J.

    2003-01-01

    Loading noise is predicted from unsteady surface pressure measurements on a NACA 0015 airfoil immersed in grid-generated turbulence. The time-dependent pressure is obtained from an array of synchronized transducers on the airfoil surface. Far field noise is predicted by using the time-dependent surface pressure as input to Formulation 1A of Farassat, a solution of the Ffowcs Williams - Hawkings equation. Acoustic predictions are performed with and without the effects of airfoil surface curvature. Scaling rules are developed to compare the present far field predictions with acoustic measurements that are available in the literature.

  2. MATE program: Erosion resistant compressor airfoil coating, volume 2

    NASA Technical Reports Server (NTRS)

    Freling, Melvin

    1987-01-01

    The performance of candidate erosion resistant airfoil coatings installed in ground tested experimental JT8D and JT9D engines and subjected to cyclic endurance at idle, takeoff and intermediate power conditions has been evaluated. Engine tests were terminated prior to the scheduled 1000 cycles of endurance test due to high cycle fatigue fracture of the Gator-Gard plasma sprayed 88WC-12Co coating on titanium alloy airfoils. Coated steel (AMS5616) and nickel base alloy (Incoloy 901) performed well in both engine tests. Post test airfoil analyses consisted of binocular, scanning electron microscope and metallographic examinations.

  3. Predictions of airfoil aerodynamic performance degradation due to icing

    NASA Technical Reports Server (NTRS)

    Shaw, Robert J.; Potapezuk, Mark G.; Bidwell, Colin S.

    1988-01-01

    An overview of NASA's ongoing efforts to develop an airfoil icing analysis capability is developed. An indication is given to the approaches being followed to calculate the water droplet trajectories past the airfoil, the buildup of ice on the airfoil, and the resultant changes in aerodynamic performance due to the leading edge ice accretion. Examples are given of current code capabilities/limitations through comparisons of predictions with experimental data gathered in various calibration/validation experiments. A brief discussion of future efforts to extend the analysis to handle three dimensional components is included.

  4. Wind tunnel testing of low-drag airfoils

    NASA Technical Reports Server (NTRS)

    Harvey, W. Donald; Mcghee, R. J.; Harris, C. D.

    1986-01-01

    Results are presented for the measured performance recently obtained on several airfoil concepts designed to achieve low drag by maintaining extensive regions of laminar flow without compromising high-lift performance. The wind tunnel results extend from subsonic to transonic speeds and include boundary-layer control through shaping and suction. The research was conducted in the NASA Langley 8-Ft Transonic Pressure Tunnel (TPT) and Low Turbulence Pressure Tunnel (LTPT) which have been developed for testing such low-drag airfoils. Emphasis is placed on identifying some of the major factors influencing the anticipated performance of low-drag airfoils.

  5. Influence of airfoil thickness on convected gust interaction noise

    NASA Technical Reports Server (NTRS)

    Kerschen, E. J.; Tsai, C. T.

    1989-01-01

    The case of a symmetric airfoil at zero angle of attack is considered in order to determine the influence of airfoil thickness on sound generated by interaction with convected gusts. The analysis is based on a linearization of the Euler equations about the subsonic mean flow past the airfoil. Primary sound generation is found to occur in a local region surrounding the leading edge, with the size of the local region scaling on the gust wavelength. For a parabolic leading edge, moderate leading edge thickness is shown to decrease the noise level in the low Mach number limit.

  6. Nature-Inspired Airfoils for Environmental Noise Reduction

    NASA Astrophysics Data System (ADS)

    Han, Suyeong; Kyung, Richard

    2013-11-01

    Recently, study on the insects' flapping flight became one of the challenging research subjects in the field of environmental engineering and aeronautics because of its potential applicability to intelligent micro-robots capable of autonomous flight and the next generation aerial-vehicles. In order to uncover its curious unsteady characteristics, many researchers have conducted experimental and computational studies on the unsteady aerodynamics of insects' flapping flight. In the present work, the unsteady flow physics around insect wings are conducted by utilizing numerical and computational simulation. The e-AIRS [6] (e-Science Aerospace Integrated Research System) gives a balanced service between computational and experimental aerodynamics, along with integrated research process of these two research activities. This paper presents the wing motions and their aerodynamics with a two dimensional approach to reduce environmental noise during the airflight. Also this paper shows an optimal phase angle, where the thrust is maximized at the position of minimized drag, which occurs when noise is minimized. Aside from the two-dimensional approach, stroke angles and phase angles of the airfoils are set as parameters, to determine which motion yields the best aerodynamic characteristics.

  7. [Distribution Characteristics and Source Apportionment of n-Alkanes in Water from Yellow River in Henan Section].

    PubMed

    Feng, Jing-lan; Xi, Nan-nan; Zhang, Fei; Liu, Shu-hui; Sun, Jian-hui

    2016-03-15

    To investigate the distributions and possible sources of n-alkanes in water and suspended particulate matter from Yellow River in Henan section, 26 water and suspended particulate matter samples were collected in August 2010 and 22 n-alkanes (C₁₄-C₃₆) were quantitatively determined by gas chromatography-mass spectrometer (GC-MS). Potential sources of n-alkanes were analyzed using different characteristic parameters. The results indicated that total concentrations of 22 n-alkanes were 521-5,843 ng · L⁻¹ with a mean concentration of 1,409 ng · L⁻¹, while the total amounts of n-alkanes in the suspended particulate matter were 463-11,142 ng · L⁻¹ with a mean value of 1,951 ng · L⁻¹. The composition profiles of n-alkanes in water showed unimodal distribution with a peak at C₂₅ in water. However, the composition characteristics of n-alkanes in SPM were of bimodal type, but still with the advantage of high carbon hydrocarbons peak at C₂₅. Results of characteristic parameters including CPI, TAR, OEP and % WaxCn showed that n-alkanes in the studied area were derived mainly from combustion of fossil fuel, while terrestrial higher plant played a role in the existence of n-alkanes in water and suspended particulate matter from Yellow River in Henan section. PMID:27337879

  8. Subtyping Somatic Tinnitus: A Cross-Sectional UK Cohort Study of Demographic, Clinical and Audiological Characteristics

    PubMed Central

    Ward, Jamie; Vella, Claire; Hoare, Derek J.; Hall, Deborah A.

    2015-01-01

    Somatic tinnitus is the ability to modulate the psychoacoustic features of tinnitus by somatic manoeuvres. The condition is still not fully understood and further identification of this subtype is essential, particularly for the purpose of establishing protocols for both its diagnosis and treatment. This study aimed to investigate the characteristics of somatic tinnitus within a large UK cohort using a largely unselected sample. We believe this to be relatively unique in comparison to current literature on the topic. This was investigated by using a total of 608 participant assessments from a set of recognised tinnitus and audiology measures. Results from a set of chi-square tests of association found that amongst the individuals with somatic tinnitus, a higher proportion had pulsatile tinnitus (different from heartbeat), were under the age of 40, reported variation in the loudness of their tinnitus and reported temporomandibular joint (TMJ) disorder. The same pattern of results was confirmed using a multivariate analysis of the data based on logistic regression. These findings have strong implications towards the profiling of somatic tinnitus as a distinct subtype of general tinnitus. PMID:25996779

  9. Improvements in Heat Transfer for Anti-Icing of Gas-Heated Airfoils with Internal Fins and Partitions

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H.

    1950-01-01

    The effect of modifying the gas passage of hollow metal airfoils by the additIon of internal fins and partitions was experimentally investigated and comparisons were made among a basic unfinned airfoil section and two airfoil designs having metal fins attached at the leading edge of the internal gas passage. An analysis considering the effects of heat conduction in the airfoil metal was made to determine the internal modification effectiveness that may be obtained in gas-heated components, such as turbojet-inlet guide vanes, support struts, hollow propeller blades, arid. thin wings. Over a wide range of heated-gas flow and tunnel-air velocity, the increase In surface-heating rates with internal finning was marked (up to 3.5 times), with the greatest increase occurring at the leading edge where anti-icing heat requirements are most critical. Variations in the amount and the location of internal finning and. partitioning provided. control over the local rates of surface heat transfer and permitted efficient anti-icing utilization of the gas-stream heat content.

  10. Population characteristics, mechanisms of primary care and premature mortality in England: a cross-sectional study

    PubMed Central

    Baker, Richard; Honeyford, Kate; Levene, Louis S; Mainous, Arch G; Jones, David R; Bankart, M John; Stokes, Tim

    2016-01-01

    Objectives Health systems with strong primary care tend to have better population outcomes, but in many countries demand for care is growing. We sought to identify mechanisms of primary care that influence premature mortality. Design We developed a conceptual model of the mechanisms by which primary care influences premature mortality, and undertook a cross-sectional study in which population and primary care variables reflecting the model were used to explain variations in mortality of those aged under 75 years. The premature standardised mortality ratios (SMRs) for each practice, available from the Department of Health, had been calculated from numbers of deaths in the 5 years from 2006 to 2010. A regression model was undertaken with explanatory variables for the year 2009/2010, and repeated to check stability using data for 2008/2009 and 2010/2011. Setting All general practices in England were eligible for inclusion and, of the total of 8290, complete data were available for 7858. Results Population variables, particularly deprivation, were the most powerful predictors of premature mortality, but the mechanisms of primary care depicted in our model also affected mortality. The number of GPs/1000 population and detection of hypertension were negatively associated with mortality. In less deprived practices, continuity of care was also negatively associated with mortality. Conclusions Greater supply of primary care is associated with lower premature mortality even in a health system that has strong primary care (England). Health systems need to sustain the capacity of primary care to deliver effective care, and should assist primary care providers in identifying and meeting the needs of socioeconomically deprived groups. PMID:26868945

  11. Characteristic plateau in the L{sub 1}-subshell ionization cross section of Ag induced by proton collisions

    SciTech Connect

    Kavcic, M.; Smit, Z.

    2009-12-15

    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 Lbeta{sub 3} (L{sub 1}-M{sub 3}) and Lbeta{sub 6} (L{sub 3}-N{sub 1}) 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.

  12. Aerodynamic Characteristics at High Speeds of Related Full-Scale Propellers Having Different Blade-Section Cambers

    NASA Technical Reports Server (NTRS)

    Maynard, Julian D; Salters, Leland B , Jr

    1957-01-01

    Wind-tunnel tests of a full-scale two-blade NACA 10-(10)(08)-03 (high camber) propeller have been made for a range of blade angles from 20 degrees to 55 degrees at airspeeds up to 500 miles per hour. The results of these tests have been compared with results from previous tests of the NACA 10-(3) (08)-03 (low camber) and NACA 10-(5)(08)-03 (medium camber) propellers to evaluate the effects of blade-section camber on propeller aerodynamic characteristics.

  13. Comparison of pressure distributions on model and full-scale NACA 64-621 airfoils with ailerons for wind turbine application

    NASA Technical Reports Server (NTRS)

    Gregorek, G. M.; Kuniega, R. J.; Nyland, T. W.

    1988-01-01

    The aerodynamic similarity between a small (4-inch chord) wind tunnel model and a full-scale wind turbine blade (24-foot tip section with a 36-inch chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-inch High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-foot Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the 3-D effects of blade twist and downwash in the blade tip section. The results of the two different experimental methods correlated well for angles of attack from minus 4 to 36 degrees and aileron reflections from 0 to 90 degrees.

  14. Heat Transfer of Airfoils and Plates

    NASA Technical Reports Server (NTRS)

    Seibert, Otto

    1943-01-01

    The few available test data on the heat dissipation of wholly or partly heated airfoil models are compared with the corresponding data for the flat plate as obtained by an extension of Prandtl's momentum theory, with differentiation between laminar and turbulent boundary layer and transitional region between both, the extent and appearance of which depend upon certain critical factors. The satisfactory agreement obtained justifies far-reaching conclusions in respect to other profile forms and arrangements of heated surface areas. The temperature relationship of the material quantities in its effect on the heat dissipation is discussed as far as is possible at tk.e present state of research, and it is shown that the profile drag of heated wing surfaces can increase or decrease with the temperature increase depending upon the momentarily existent structure of the boundary layer.

  15. Turbine airfoil with a compliant outer wall

    DOEpatents

    Campbell, Christian X.; Morrison, Jay A.

    2012-04-03

    A turbine airfoil usable in a turbine engine with a cooling system and a compliant dual wall configuration configured to enable thermal expansion between inner and outer layers while eliminating stress formation in the outer layer is disclosed. The compliant dual wall configuration may be formed a dual wall formed from inner and outer layers separated by a support structure. The outer layer may be a compliant layer configured such that the outer layer may thermally expand and thereby reduce the stress within the outer layer. The outer layer may be formed from a nonplanar surface configured to thermally expand. In another embodiment, the outer layer may be planar and include a plurality of slots enabling unrestricted thermal expansion in a direction aligned with the outer layer.

  16. Cooled airfoil in a turbine engine

    SciTech Connect

    Vitt, Paul H; Kemp, David A; Lee, Ching-Pang; Marra, John J

    2015-04-21

    An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall.

  17. Design and experimental results for the S814 airfoil

    SciTech Connect

    Somers, D.M.

    1997-01-01

    A 24-percent-thick airfoil, the S814, for the root region of a horizontal-axis wind-turbine blade has been designed and analyzed theoretically and verified experimentally in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory, The Netherlands. The two primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on the pitching moment and the airfoil thickness have been satisfied. Comparisons of the theoretical and experimental results show good agreement with the exception of maximum lift which is overpredicted. Comparisons with other airfoils illustrate the higher maximum lift and the lower profile drag of the S814 airfoil, thus confirming the achievement of the objectives.

  18. Comparison of Analytical and CAA Solutions: Single Airfoil Gust Response

    NASA Technical Reports Server (NTRS)

    Scott, James R.

    2000-01-01

    The purpose of this problem is to test the ability of a computational fluid dynamics/computational aeroacoustics code to accurately predict the unsteady aerodynamic and aeroacoustic response of a single airfoil to a two-dimensional, periodic vortical gust.

  19. Aeroacoustics and aerodynamic performance of a rotor with flatback airfoils.

    SciTech Connect

    Paquette, Joshua A.; Barone, Matthew Franklin; Christiansen, Monica; Simley, Eric

    2010-06-01

    The aerodynamic performance and aeroacoustic noise sources of a rotor employing flatback airfoils have been studied in field test campaign and companion modeling effort. The field test measurements of a sub-scale rotor employing nine meter blades include both performance measurements and acoustic measurements. The acoustic measurements are obtained using a 45 microphone beamforming array, enabling identification of both noise source amplitude and position. Semi-empirical models of flatback airfoil blunt trailing edge noise are developed and calibrated using available aeroacoustic wind tunnel test data. The model results and measurements indicate that flatback airfoil noise is less than drive train noise for the current test turbine. It is also demonstrated that the commonly used Brooks, Pope, and Marcolini model for blunt trailing edge noise may be over-conservative in predicting flatback airfoil noise for wind turbine applications.

  20. Unsteady transonic flow control around an airfoil in a channel

    NASA Astrophysics Data System (ADS)

    Hamid, Md. Abdul; Hasan, A. B. M. Toufique; Ali, Mohammad; Mitsutake, Yuichi; Setoguchi, Toshiaki; Yu, Shen

    2016-04-01

    Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings (perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square (RMS) of pressure oscillation around the airfoil have been reduced with the control method.

  1. High-flaps for natural laminar flow airfoils

    NASA Technical Reports Server (NTRS)

    Morgan, Harry L.

    1986-01-01

    A review of the NACA and NASA low-drag airfoil research is presented with particular emphasis given to the development of mechanical high-lift flap systems and their application to general aviation aircraft. These flap systems include split, plain, single-slotted, and double-slotted trailing-edge flaps plus slat and Krueger leading-edge devices. The recently developed continuous variable-camber high-lift mechanism is also described. The state-of-the-art of theoretical methods for the design and analysis of multi-component airfoils in two-dimensional subsonic flow is discussed, and a detailed description of the Langley MCARF (Multi-Component Airfoil Analysis Program) computer code is presented. The results of a recent effort to design a single- and double-slotted flap system for the NASA high speed natural laminar flow (HSNLF) (1)-0213 airfoil using the MCARF code are presented to demonstrate the capabilities and limitations of the code.

  2. Experimental study of airfoil separation control using synthetic jets

    NASA Astrophysics Data System (ADS)

    Xia, Xi; Mohseni, Kamran

    2010-11-01

    The flow control over an airfoil is studied experimentally in a wind tunnel. Synthetic jets are placed on the top surface of the airfoil as flow actuators. The position and the angle of the jet orifice, together with the frequency and jet strength could be varied in order to adjust the separation or reattachment points on the surface. An Array of hot-film sensors are located on the surface in order to detect the location of the reattachment point. The airfoil is mounted on a 6 d.o.f force balance system to dynamically measure the drag and lift forces on the airfoil. Results from the hot-film sensor array measurement are correlated to the measured drag and lift forces.

  3. Grid Sensitivity and Aerodynamic Optimization of Generic Airfoils

    NASA Technical Reports Server (NTRS)

    Sadrehaghighi, Ideen; Smith, Robert E.; Tiwari, Surendra N.

    1995-01-01

    An algorithm is developed to obtain the grid sensitivity with respect to design parameters for aerodynamic optimization. The procedure is advocating a novel (geometrical) parameterization using spline functions such as NURBS (Non-Uniform Rational B- Splines) for defining the airfoil geometry. An interactive algebraic grid generation technique is employed to generate C-type grids around airfoils. The grid sensitivity of the domain with respect to geometric design parameters has been obtained by direct differentiation of the grid equations. A hybrid approach is proposed for more geometrically complex configurations such as a wing or fuselage. The aerodynamic sensitivity coefficients are obtained by direct differentiation of the compressible two-dimensional thin-layer Navier-Stokes equations. An optimization package has been introduced into the algorithm in order to optimize the airfoil surface. Results demonstrate a substantially improved design due to maximized lift/drag ratio of the airfoil.

  4. Active Control of Flow Separation Over an Airfoil

    NASA Technical Reports Server (NTRS)

    Ravindran, S. S.

    1999-01-01

    Designing an aircraft without conventional control surfaces is of interest to aerospace community. In this direction, smart actuator devices such as synthetic jets have been proposed to provide aircraft maneuverability instead of control surfaces. In this article, a numerical study is performed to investigate the effects of unsteady suction and blowing on airfoils. The unsteady suction and blowing is introduced at the leading edge of the airfoil in the form of tangential jet. Numerical solutions are obtained using Reynolds-Averaged viscous compressible Navier-Stokes equations. Unsteady suction and blowing is investigated as a means of separation control to obtain lift on airfoils. The effect of blowing coefficients on lift and drag is investigated. The numerical simulations are compared with experiments from the Tel-Aviv University (TAU). These results indicate that unsteady suction and blowing can be used as a means of separation control to generate lift on airfoils.

  5. Natural laminar flow airfoil analysis and trade studies

    NASA Technical Reports Server (NTRS)

    1979-01-01

    An analysis of an airfoil for a large commercial transport cruising at Mach 0.8 and the use of advanced computer techniques to perform the analysis are described. Incorporation of the airfoil into a natural laminar flow transport configuration is addressed and a comparison of fuel requirements and operating costs between the natural laminar flow transport and an equivalent turbulent flow transport is addressed.

  6. An assessment of airfoil design by numerical optimization

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Murman, E. M.; Vanderplaats, G. N.

    1974-01-01

    A practical procedure for optimum design of aerodynamic shapes is demonstrated. The proposed procedure uses an optimization program based on the method of feasible directions coupled with an analysis program that uses a relaxation solution of the inviscid, transonic, small-disturbance equations. Results are presented for low-drag, nonlifting transonic airfoils. Extension of the method to lifting airfoils, other speed regimes, and to three dimensions if feasible.

  7. An inverse design method for 2D airfoil

    NASA Astrophysics Data System (ADS)

    Liang, Zhi-Yong; Cui, Peng; Zhang, Gen-Bao

    2010-03-01

    The computational method for aerodynamic design of aircraft is applied more universally than before, in which the design of an airfoil is a hot problem. The forward problem is discussed by most relative papers, but inverse method is more useful in practical designs. In this paper, the inverse design of 2D airfoil was investigated. A finite element method based on the variational principle was used for carrying out. Through the simulation, it was shown that the method was fit for the design.

  8. Incidence angle effects on convected gust airfoil noise

    NASA Astrophysics Data System (ADS)

    Kerschen, E. J.; Myers, M. R.

    1983-04-01

    An analysis is developed which predicts the influence of airfoil mean loading on noise generation due to convected gusts. The theory is based on a linearization of the exact inviscid equations about a nonuniform compressible mean flow and the solution is developed using singular perturbation techniques. The case of a flat plate airfoil, at incidence angle alpha, interacting with three-dimensional disturbances is analyzed. It is found that in the vicinity of the airfoil leading and trailing edges, local regions are present which scale on the disturbance wavelength, with the noise generation concentrated in these regions. Away from the airfoil edges, the mean flow variation is found to be slow compared to the disturbance wavelength and no significant noise generation occurs. The mean flow variation near the leading edge generates additional noise by distorting the convected gust. The cumulative effect of the airfoil mean loading in the trailing edge region produces a 0(1) phase shift between the disturbances on the upper and lower surfaces of the airfoil. A corresponding 0(1) decrease, compared to the alpha = 0 case, is found in the noise generated at the trailing edge.

  9. Computer programs for smoothing and scaling airfoil coordinates

    NASA Technical Reports Server (NTRS)

    Morgan, H. L., Jr.

    1983-01-01

    Detailed descriptions are given of the theoretical methods and associated computer codes of a program to smooth and a program to scale arbitrary airfoil coordinates. The smoothing program utilizes both least-squares polynomial and least-squares cubic spline techniques to smooth interatively 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. A technique for computing the camber and thickness distribution of the smoothed airfoil is also discussed. The scaling program can then be used to scale the thickness distribution generated by the smoothing program to a specific maximum thickness which is then combined with the camber distribution to obtain the final scaled airfoil contour. Computer listings of the smoothing and scaling programs are included.

  10. Symmetric airfoil geometry effects on leading edge noise.

    PubMed

    Gill, James; Zhang, X; Joseph, P

    2013-10-01

    Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number. PMID:24116405

  11. Multiple Solutions of Transonic Flow over NACA0012 Airfoil

    NASA Astrophysics Data System (ADS)

    Xiong, Juntao; Liu, Ya; Liu, Feng; Luo, Shijun; Zhao, Zijie; Ren, Xudong; Gao, Chao

    2012-11-01

    Multiple solutions of the small-disturbance potential equation and full potential equation were known for the NACA0012 airfoil in a certain range of transonic Mach numbers and at zero angle of attack. However the multiple solutions for this airfoil were not observed using Euler or Navier-Stokes equations under the above flow conditions. In the present work, both the Unsteady Reynolds-Averaged Navier-Stokes (URANS) computations and transonic wind tunnel experiments are performed under certain Reynolds numbers to further study the problem. The results of the two methods reveal that buffet appears in a narrow Mach number range where the potential flow methods predict multiple solutions. Boundary layer displacement thickness computed from URANS at the same flow condition is used to modify the geometry of the airfoil. Euler equations are then solved for the modified geometry. The results show that the addition of the boundary layer displacement thickness creates multiple solutions for the NACA0012 airfoil. Global linear stability analysis is also performed on the original and the modified airfoils. This shows a close relationship between the viscous unsteady shock buffet phenomenon of transonic airfoil flow and the existence of multiple solutions of the external inviscid flow. Postdoctoral Research Assistant.

  12. Symmetric airfoil geometry effects on leading edge noise.

    PubMed

    Gill, James; Zhang, X; Joseph, P

    2013-10-01

    Computational aeroacoustic methods are applied to the modeling of noise due to interactions between gusts and the leading edge of real symmetric airfoils. Single frequency harmonic gusts are interacted with various airfoil geometries at zero angle of attack. The effects of airfoil thickness and leading edge radius on noise are investigated systematically and independently for the first time, at higher frequencies than previously used in computational methods. Increases in both leading edge radius and thickness are found to reduce the predicted noise. This noise reduction effect becomes greater with increasing frequency and Mach number. The dominant noise reduction mechanism for airfoils with real geometry is found to be related to the leading edge stagnation region. It is shown that accurate leading edge noise predictions can be made when assuming an inviscid meanflow, but that it is not valid to assume a uniform meanflow. Analytic flat plate predictions are found to over-predict the noise due to a NACA 0002 airfoil by up to 3 dB at high frequencies. The accuracy of analytic flat plate solutions can be expected to decrease with increasing airfoil thickness, leading edge radius, gust frequency, and Mach number.

  13. An Experimental Investigation of Unsteady Surface Pressure on an Airfoil in Turbulence

    NASA Technical Reports Server (NTRS)

    Mish, Patrick F.; Devenport, William J.

    2003-01-01

    Measurements of fluctuating surface pressure were made on a NACA 0015 airfoil immersed in grid generated turbulence. The airfoil model has a 2 ft chord and spans the 6 ft Virginia Tech Stability Wind Tunnel test section. Two grids were used to investigate the effects of turbulence length scale on the surface pressure response. A large grid which produced turbulence with an integral scale 13% of the chord and a smaller grid which produced turbulence with an integral scale 1.3% of the chord. Measurements were performed at angles of attack, alpha from 0 to 20 . An array of microphones mounted subsurface was used to measure the unsteady surface pressure. The goal of this measurement was to characterize the effects of angle of attack on the inviscid response. Lift spectra calculated from pressure measurements at each angle of attack revealed two distinct interaction regions; for omega(sub r) = omega b / U(sub infinity) is less than 10 a reduction in unsteady lift of up to 7 decibels (dB) occurs while an increase occurs for omega(sub r) is greater than 10 as the angle of attack is increased. The reduction in unsteady lift at low omega(sub r) with increasing angle of attack is a result that has never before been shown either experimentally or theoretically. The source of the reduction in lift spectral level appears to be closely related to the distortion of inflow turbulence based on analysis of surface pressure spanwise correlation length scales. Furthermore, while the distortion of the inflow appears to be critical in this experiment, this effect does not seem to be significant in larger integral scale (relative to the chord) flows based on the previous experimental work of McKeough suggesting the airfoils size relative to the inflow integral scale is critical in defining how the airfoil will respond under variation of angle of attack. A prediction scheme is developed that correctly accounts for the effects of distortion when the inflow integral scale is small relative

  14. Global surface pressure measurements of static and dynamic stall on a wind turbine airfoil at low Reynolds number

    NASA Astrophysics Data System (ADS)

    Disotell, Kevin J.; Nikoueeyan, Pourya; Naughton, Jonathan W.; Gregory, James W.

    2016-05-01

    Recognizing the need for global surface measurement techniques to characterize the time-varying, three-dimensional loading encountered on rotating wind turbine blades, fast-responding pressure-sensitive paint (PSP) has been evaluated for resolving unsteady aerodynamic effects in incompressible flow. Results of a study aimed at demonstrating the laser-based, single-shot PSP technique on a low Reynolds number wind turbine airfoil in static and dynamic stall are reported. PSP was applied to the suction side of a Delft DU97-W-300 airfoil (maximum thickness-to-chord ratio of 30 %) at a chord Reynolds number of 225,000 in the University of Wyoming open-return wind tunnel. Static and dynamic stall behaviors are presented using instantaneous and phase-averaged global pressure maps. In particular, a three-dimensional pressure topology driven by a stall cell pattern is detected near the maximum lift condition on the steady airfoil. Trends in the PSP-measured pressure topology on the steady airfoil were confirmed using surface oil visualization. The dynamic stall case was characterized by a sinusoidal pitching motion with mean angle of 15.7°, amplitude of 11.2°, and reduced frequency of 0.106 based on semichord. PSP images were acquired at selected phase positions, capturing the breakdown of nominally two-dimensional flow near lift stall, development of post-stall suction near the trailing edge, and a highly three-dimensional topology as the flow reattaches. Structural patterns in the surface pressure topologies are considered from the analysis of the individual PSP snapshots, enabled by a laser-based excitation system that achieves sufficient signal-to-noise ratio in the single-shot images. The PSP results are found to be in general agreement with observations about the steady and unsteady stall characteristics expected for the airfoil.

  15. Application of Artificial Neural Networks to the Design of Turbomachinery Airfoils

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan; Madavan, Nateri

    1997-01-01

    Artificial neural networks are widely used in engineering applications, such as control, pattern recognition, plant modeling and condition monitoring to name just a few. In this seminar we will explore the possibility of applying neural networks to aerodynamic design, in particular, the design of turbomachinery airfoils. The principle idea behind this effort is to represent the design space using a neural network (within some parameter limits), and then to employ an optimization procedure to search this space for a solution that exhibits optimal performance characteristics. Results obtained for design problems in two spatial dimensions will be presented.

  16. A least squares finite element scheme for transonic flow around harmonically oscillating airfoils

    NASA Technical Reports Server (NTRS)

    Cox, C. L.; Fix, G. J.; Gunzburger, M. D.

    1983-01-01

    The present investigation shows that a finite element scheme with a weighted least squares variational principle is applicable to the problem of transonic flow around a harmonically oscillating airfoil. For the flat plate case, numerical results compare favorably with the exact solution. The obtained numerical results for the transonic problem, for which an exact solution is not known, have the characteristics of known experimental results. It is demonstrated that the performance of the employed numerical method is independent of equation type (elliptic or hyperbolic) and frequency. The weighted least squares principle allows the appropriate modeling of singularities, which such a modeling of singularities is not possible with normal least squares.

  17. Comparison of Several Methods of Cyclic De-Icing of a Gas-Heated Airfoil

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H.; Bowden, Dean T.

    1953-01-01

    Several methods of cyclic de-icing of a gas-heated airfoil were investigated to determine ice-removal characteristics and heating requirements. The cyclic de-icing system with a spanwise ice-free parting strip in the stagnation region and a constant-temperature gas-supply duct gave the quickest and most reliable ice removal. Heating requirements for the several methods of cyclic de-icing are compared, and the savings over continuous ice prevention are shown. Data are presented to show the relation of surface temperature, rate of surface heating, and heating time to the removal of ice.

  18. The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil - Drag equations

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    The Langley Research Center has designed a swept, supercritical airfoil incorporating Laminar Flow Control for testing at transonic speeds. Analytical expressions have been developed and an evaluation made of the experimental section drag, composed of suction drag and wake drag, using theoretical design information and experimental data. The analysis shows that, although the sweep-induced boundary-layer crossflow influence on the wake drag is too large to be ignored and there is not a practical method for evaluating these crossflow effects on the experimental wake data, the conventional unswept 2-D wake-drag computation used in the reduction of the experimental data is at worst 10 percent too high.

  19. Investigation of nonlinear inviscid and viscous flow effects in the analysis of dynamic stall. [air flow and chordwise pressure distribution on airfoil below stall condition

    NASA Technical Reports Server (NTRS)

    Crimi, P.

    1974-01-01

    A method for analyzing unsteady airfoil stall was refined by including nonlinear effects in the representation of the inviscid flow. Certain other aspects of the potential-flow model were reexamined and the effects of varying Reynolds number on stall characteristics were investigated. Refinement of the formulation improved the representation of the flow and chordwise pressure distribution below stall, but substantial quantitative differences between computed and measured results are still evident for sinusoidal pitching through stall. Agreement is substantially improved by assuming the growth rate of the dead-air region at the onset of leading-edge stall is of the order of the component of the free stream normal to the airfoil chordline. The method predicts the expected increase in the resistance to stalling with increasing Reynolds number. Results indicate that a given airfoil can undergo both trailing-edge and leading-edge stall under unsteady conditions.

  20. Pressure-distribution investigation of an N.A.C.A. 0009 airfoil with a 50-percent-chord plain flap and three tabs

    NASA Technical Reports Server (NTRS)

    Street, William G; Ames, Milton B

    1939-01-01

    Pressure-distribution tests of an N.A.C.A. 0009 airfoil with a 50-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made in the N.A.C.A. 4- by 6- foot vertical tunnel. The tests supplied aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. The experimental results and values computed by analytical methods are in good agreement for small flap and tab deflections. The results of the tests indicated that the effectiveness of all three tab sizes in reducing flap hinge moments decreased with increasing flap deflection.