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Sample records for airfoil suction surface

  1. 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.

  2. Design considerations of advanced supercritical low drag suction airfoils

    NASA Technical Reports Server (NTRS)

    Pfenninger, W.; Reed, H. L.; Dagenhart, J. R.

    1980-01-01

    Supercritical low drag suction laminar flow airfoils were laid out for shock-free flow at design freestream Mach = 0.76, design lift coefficient = 0.58, and t/c = 0.13. The design goals were the minimization of suction laminarization problems and the assurance of shock-free flow at freestream Mach not greater than design freestream Mach (for design lift coefficient) as well as at lift coefficient not greater than design lift coefficient (for design freestream Mach); this involved limiting the height-to-length ratio of the supersonic zone at design to 0.35. High design freestream Mach numbers result with extensive supersonic flow (over 80% of the chord) on the upper surface, with a steep Stratford-type rear pressure rise with suction, as well as by carrying lift essentially in front- and rear-loaded regions of the airfoil with high static pressures on the carved out front and rear lower surface.

  3. 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.

  4. 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.

  5. Investigation of viscous/inviscid interaction in transonic flow over airfoils with suction

    NASA Technical Reports Server (NTRS)

    Vemuru, C. S.; Tiwari, S. N.

    1988-01-01

    The viscous/inviscid interaction over transonic airfoils with and without suction is studied. The streamline angle at the edge of the boundary layer is used to couple the viscous and inviscid flows. The potential flow equations are solved for the inviscid flow field. In the shock region, the Euler equations are solved using the method of integral relations. For this, the potential flow solution is used as the initial and boundary conditions. An integral method is used to solve the laminar boundary-layer equations. Since both methods are integral methods, a continuous interaction is allowed between the outer inviscid flow region and the inner viscous flow region. To avoid the Goldstein singularity near the separation point the laminar boundary-layer equations are derived in an inverse form to obtain solution for the flows with small separations. The displacement thickness distribution is specified instead of the usual pressure distribution to solve the boundry-layer equations. The Euler equations are solved for the inviscid flow using the finite volume technique and the coupling is achieved by a surface transpiration model. A method is developed to apply a minimum amount of suction that is required to have an attached flow on the airfoil. The turbulent boundary layer equations are derived using the bi-logarithmic wall law for mass transfer. The results are found to be in good agreement with available experimental data and with the results of other computational methods.

  6. 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.

  7. 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

    roughness and freestream turbulence, compared with data from the cambered vane airfoil. Stanton numbers, skin friction coefficients, aerodynamic losses, and Reynolds analogy behavior are numerically predicted for a turbine vane using the FLUENT with a k-epsilon RNG model to show the effects of Mach number, mainstream turbulence level, and surface roughness. Comparisons with wake aerodynamic loss experimental data are made. Numerically predicted skin friction coefficients and Stanton numbers are also used to deduce Reynolds analogy behavior on the vane suction and pressure sides.

  8. The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Suction coefficient analysis

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

    A swept supercritical wing incorporating laminar flow control at transonic flow conditions was designed and tested. The definition of an experimental suction coefficient and a derivation of the compressible and incompressible formulas for the computation of the coefficient from measurable quantities is presented. The suction flow coefficient in the highest velocity nozzles is shown to be overpredicted by as much as 12 percent through the use of an incompressible formula. However, the overprediction on the computed value of suction drag when some of the suction nozzles were operating in the compressible flow regime is evaluated and found to be at most 6 percent at design conditions.

  9. Perforated Sheets as the Porous Material for a Suction-flap Application

    NASA Technical Reports Server (NTRS)

    Dannenberg, Robert E; Weiburg, James A; Gambucci, Bruno J

    1957-01-01

    Two-dimensional tests were made of an NACA 0006 airfoil with area suction applied to a porous region on a 0.3-chord trailing-edge flap deflected 50 degrees. The lift with suction approached the value computed from thin-airfoil theory. The lift gains and suction quantity requirements were unaffected by the perforation patterns of the surface over a wide range of hole sizes and spacings.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. The NASA Langley laminar-flow-control experiment on a swept, supercritical airfoil: Evaluation of initial perforated configuration

    NASA Technical Reports Server (NTRS)

    Harris, Charles D.; Brooks, Cuyler W., Jr.; Clukey, Patricia G.; Stack, John P.

    1992-01-01

    The initial evaluation of a large-chord, swept, supercritical airfoil incorporating an active laminar-flow-control (LFC) suction system with a perforated upper surface is documented in a chronological manner, and the deficiencies in the suction capability of the perforated panels as designed are described. The experiment was conducted in the Langley 8-Foot Transonic Pressure Tunnel. Also included is an evaluation of the influence of the proximity of the tunnel liner to the upper surface of the airfoil pressure distribution.

  17. Airfoil-shaped micro-mixers for reducing fouling on membrane surfaces

    SciTech Connect

    Ho, Clifford K; Altman, Susan J; Clem, Paul G; Hibbs, Michael; Cook, Adam W

    2012-10-23

    An array of airfoil-shaped micro-mixers that enhances fluid mixing within permeable membrane channels, such as used in reverse-osmosis filtration units, while minimizing additional pressure drop. The enhanced mixing reduces fouling of the membrane surfaces. The airfoil-shaped micro-mixer can also be coated with or comprised of biofouling-resistant (biocidal/germicidal) ingredients.

  18. 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.

  19. 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.

  20. Integrated axial and tangential serpentine cooling circuit in a turbine airfoil

    SciTech Connect

    Lee, Ching-Pang; Jiang, Nan; Marra, John J; Rudolph, Ronald J; Dalton, John P

    2015-05-05

    A continuous serpentine cooling circuit forming a progression of radial passages (44, 45, 46, 47A, 48A) between pressure and suction side walls (52, 54) in a MID region of a turbine airfoil (24). The circuit progresses first axially, then tangentially, ending in a last radial passage (48A) adjacent to the suction side (54) and not adjacent to the pressure side (52). The passages of the axial progression (44, 45, 46) may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (47A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (47A, 48A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.

  1. 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.

  2. 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.

  3. Tests of Large Airfoils in the Propeller Research Tunnel, Including Two with Corrugated Surfaces

    NASA Technical Reports Server (NTRS)

    Wood, Donald H

    1930-01-01

    This report gives the results of the tests of seven 2 by 12 foot airfoils (Clark Y, smooth and corrugated, Gottingen 398, N.A.C.A. M-6, and N.A.C.A. 84). The tests were made in the propeller research tunnel of the National Advisory Committee for Aeronautics at Reynolds numbers up to 2,000,000. The Clark Y airfoil was tested with three degrees of surface smoothness. Corrugating the surface causes a flattening of the lift curve at the burble point and an increase in drag at small flying angles.

  4. A study of flow past an airfoil with a jet issuing from its lower surface

    NASA Technical Reports Server (NTRS)

    Krothapalli, A.; Leopold, D.

    1984-01-01

    The aerodynamics of a NACA 0018 airfoil with a rectangular jet of finite aspect ratio exiting from its lower surface at 90 deg to the chord were investigated. The jet was located at 50% of the wing chord. Measurements include static pressures on the airfoil surface, total pressures in the near wake, and local velocity vectors in different planes of the wake. The effects of jet cross flow interaction on the aerodynamics of the airfoil are studied. It is indicated that at all values of momentum coefficients, the jet cross flow interaction produces a strong contra-rotating vortex structure in the near wake. The flow behind the jet forms a closed recirculation region which extends up to a chord length down stream of the trailing edge which results in the flow field to become highly three dimensional. The various aerodynamic force coefficients vary significantly along the span of the wing. The results are compared with a jet flap configuration.

  5. 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.

  6. An experimental investigation on the surface water transport process over an airfoil by using a digital image projection technique

    NASA Astrophysics Data System (ADS)

    Zhang, Kai; Wei, Tian; Hu, Hui

    2015-09-01

    In the present study, an experimental investigation was conducted to characterize the transient behavior of the surface water film and rivulet flows driven by boundary layer airflows over a NACA0012 airfoil in order to elucidate underlying physics of the important micro-physical processes pertinent to aircraft icing phenomena. A digital image projection (DIP) technique was developed to quantitatively measure the film thickness distribution of the surface water film/rivulet flows over the airfoil at different test conditions. The time-resolved DIP measurements reveal that micro-sized water droplets carried by the oncoming airflow impinged onto the airfoil surface, mainly in the region near the airfoil leading edge. After impingement, the water droplets formed thin water film that runs back over the airfoil surface, driven by the boundary layer airflow. As the water film advanced downstream, the contact line was found to bugle locally and developed into isolated water rivulets further downstream. The front lobes of the rivulets quickly advanced along the airfoil and then shed from the airfoil trailing edge, resulting in isolated water transport channels over the airfoil surface. The water channels were responsible for transporting the water mass impinging at the airfoil leading edge. Additionally, the transition location of the surface water transport process from film flows to rivulet flows was found to occur further upstream with increasing velocity of the oncoming airflow. The thickness of the water film/rivulet flows was found to increase monotonically with the increasing distance away from the airfoil leading edge. The runback velocity of the water rivulets was found to increase rapidly with the increasing airflow velocity, while the rivulet width and the gap between the neighboring rivulets decreased as the airflow velocity increased.

  7. Stick tight: suction adhesion on irregular surfaces in the northern clingfish

    PubMed Central

    Wainwright, Dylan K.; Kleinteich, Thomas; Kleinteich, Anja; Gorb, Stanislav N.; Summers, Adam P.

    2013-01-01

    The northern clingfish, Gobiesox maeandricus, is able to adhere to slippery, fouled and irregular surfaces in the marine intertidal environment. We have found that the fish can adhere equally well to surfaces with a broad range of surface roughness, from the finest sandpaper (Ra = 15 µm) to textures suitable for removing finish from flooring (Ra = 269 µm). The fishes outperform man-made suction cups, which only adhere to the smoothest surfaces. The adhesive forces of clingfish correspond to pressures 0.2–0.5 atm below ambient and are 80–230 times the body weight of the fish. The tenacity appears related to hierarchically structured microvilli around the edges of the adhesive disc that are similar in size and aspect ratio to the setae found on the feet of geckoes, spiders and insects. This points to a possible biomimetic solution to the problem of reversibly adhering to irregular, submerged surfaces. PMID:23637393

  8. An application of active surface heating for augmenting lift and reducing drag of an airfoil

    NASA Technical Reports Server (NTRS)

    Maestrello, Lucio; Badavi, Forooz F.; Noonan, Kevin W.

    1988-01-01

    Application of active control to separated flow on the RC(6)-08 airfoil at high angle of attack by localized surface heating is numerically simulated by integrating the compressible 2-D nonlinear Navier-Stokes equation solver. Active control is simulated by local modification of the temperature boundary condition over a narrow strip of the upper surface of the airfoil. Both mean and perturbed profiles are favorably altered when excited with the same natural frequency of the shear layer by moderate surface heating for both laminar and turbulent separation. The shear layer is found to be very sensitive to localized surface heating in the vicinity of the separation point. The excitation field at the surface sufficiently altered both the local as well as the global circulation to cause a significant increase in lift and reduction in drag.

  9. 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

  10. Experimental investigation of moving surfaces for boundary layer and circulation control of airfoils and wings

    NASA Astrophysics Data System (ADS)

    Vets, Robert

    An experimental study was conducted to assess the application of a moving surface to affect boundary layers and circulation around airfoils for the purpose of altering and enhancing aerodynamic performance of finite wings at moderate Reynolds numbers. The moving surface was established by a wide, lightweight, nylon belt that enveloped a wing's symmetric airfoil profile articulated via a friction drive cylinder such that the direction of the upper surface was in the direction of the free stream. A water tunnel visualization study accompanied wind tunnel testing at the University of Washington, Kirsten Wind Tunnel of finite wings. An experimental study was conducted to assess the application of a moving surface to affect boundary layers and circulation around airfoils for the purpose of altering and enhancing aerodynamic performance of finite wings at moderate Reynolds numbers. The moving surface was established by a wide, lightweight, nylon belt that enveloped a wing's symmetric airfoil profile articulated via a friction drive cylinder such that the direction of the upper surface was in the direction of the free stream. A water tunnel visualization study accompanied wind tunnel testing at the University of Washington, Kirsten Wind Tunnel of finite wings. The defining non-dimensional parameter for the system is the ratio of the surface velocity to the free stream velocity, us/Uo. Results show a general increase in lift with increasing us/Uo. The endurance parameter served as an additional metric for the system's performance. Examining the results of the endurance parameter shows general increase in endurance and lift with the moving surface activated. Peak performance in terms of increased endurance along with increased lift occurs at or slightly above us/Uo = 1. Water tunnel visualization showed a marked difference in the downwash for velocity ratios greater than 1, supporting the measured data. Reynolds numbers for this investigation were 1.9E5 and 4.3E5, relevant

  11. 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.

  12. Analysis of the separated boundary layer flow on the surface and in the wake of blunt trailing edge airfoils

    NASA Technical Reports Server (NTRS)

    Goradia, S. H.; Mehta, J. M.; Shrewsbury, G. S.

    1977-01-01

    The viscous flow phenomena associated with sharp and blunt trailing edge airfoils were investigated. Experimental measurements were obtained for a 17 percent thick, high performance GAW-1 airfoil. Experimental measurements consist of velocity and static pressure profiles which were obtained by the use of forward and reverse total pressure probes and disc type static pressure probes over the surface and in the wake of sharp and blunt trailing edge airfoils. Measurements of the upper surface boundary layer were obtained in both the attached and separated flow regions. In addition, static pressure data were acquired, and skin friction on the airfoil upper surface was measured with a specially constructed device. Comparison of the viscous flow data with data previously obtained elsewhere indicates reasonable agreement in the attached flow region. In the separated flow region, considerable differences exist between these two sets of measurements.

  13. Deriving the suction stress of unsaturated soils from water retention curve, based on wetted surface area in pores

    NASA Astrophysics Data System (ADS)

    Greco, Roberto; Gargano, Rudy

    2016-04-01

    The evaluation of suction stress in unsaturated soils has important implications in several practical applications. Suction stress affects soil aggregate stability and soil erosion. Furthermore, the equilibrium of shallow unsaturated soil deposits along steep slopes is often possible only thanks to the contribution of suction to soil effective stress. Experimental evidence, as well as theoretical arguments, shows that suction stress is a nonlinear function of matric suction. The relationship expressing the dependence of suction stress on soil matric suction is usually indicated as Soil Stress Characteristic Curve (SSCC). In this study, a novel equation for the evaluation of the suction stress of an unsaturated soil is proposed, assuming that the exchange of stress between soil water and solid particles occurs only through the part of the surface of the solid particles which is in direct contact with water. The proposed equation, based only upon geometric considerations related to soil pore-size distribution, allows to easily derive the SSCC from the water retention curve (SWRC), with the assignment of two additional parameters. The first parameter, representing the projection of the external surface area of the soil over a generic plane surface, can be reasonably estimated from the residual water content of the soil. The second parameter, indicated as H0, is the water potential, below which adsorption significantly contributes to water retention. For the experimental verification of the proposed approach such a parameter is considered as a fitting parameter. The proposed equation is applied to the interpretation of suction stress experimental data, taken from the literature, spanning over a wide range of soil textures. The obtained results show that in all cases the proposed relationships closely reproduces the experimental data, performing better than other currently used expressions. The obtained results also show that the adopted values of the parameter H0

  14. Instability and Transition of Flow at, and Near, an Attachment-Line: Including Control by Surface Suction

    NASA Technical Reports Server (NTRS)

    Smith, A.; Poll, D. I. A.

    1998-01-01

    Experiments have been performed on an untapered, swept cylinder model in the Cranfield College of Aeronautics 8 ft x 6 ft low-speed wind tunnel to investigate the effect of surface transpiration on the process of relaminarization in the attachment-line boundary layer. Suction coefficients for complete suppression of turbulence were determined as a function of Reynolds number and spanwise distance. The effect of attachment-line suction on the spanwise propagation of gross disturbances emanating from the fuselage-wing junction region was also studied. Finally, the effect of blowing on a laminar attachment-line boundary layer was also considered and excellent agreement was achieved with previous studies.

  15. 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

  16. 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.

  17. 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.

  18. 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)

  19. 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

  20. Wavy flow cooling concept for turbine airfoils

    DOEpatents

    Liang, George

    2010-08-31

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

  1. 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.

  2. Improvements in surface singularity analysis and design methods. [applicable to airfoils

    NASA Technical Reports Server (NTRS)

    Bristow, D. R.

    1979-01-01

    The coupling of the combined source vortex distribution of Green's potential flow function with contemporary numerical techniques is shown to provide accurate, efficient, and stable solutions to subsonic inviscid analysis and design problems for multi-element airfoils. The analysis problem is solved by direct calculation of the surface singularity distribution required to satisfy the flow tangency boundary condition. The design or inverse problem is solved by an iteration process. In this process, the geometry and the associated pressure distribution are iterated until the pressure distribution most nearly corresponding to the prescribed design distribution is obtained. Typically, five iteration cycles are required for convergence. A description of the analysis and design method is presented, along with supporting examples.

  3. A two dimensional study of rotor/airfoil interaction in hover

    NASA Technical Reports Server (NTRS)

    Lee, Chyang S.

    1988-01-01

    A two dimensional model for the chordwise flow near the wing tip of the tilt rotor in hover is presented. The airfoil is represented by vortex panels and the rotor is modeled by doublet panels. The rotor slipstream and the airfoil wake are simulated by free point vortices. Calculations on a 20 percent thick elliptical airfoil under a uniform rotor inflow are performed. Variations on rotor size, spacing between the rotor and the airfoil, ground effect, and the influence upper surface blowing in download reduction are analyzed. Rotor size has only a minor influence on download when it is small. Increase of the rotor/airfoil spacing causes a gradual decrease on download. Proximity to the ground effectively reduces the download and makes the wake unsteady. The surface blowing changes the whole flow structure and significantly reduces the download within the assumption of a potential solution. Improvement on the present model is recommended to estimate the wall jets induced suction on the airfoil lower surface.

  4. 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

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. Disturbance functions of the Goertler instability on an airfoil

    NASA Technical Reports Server (NTRS)

    Dagenhart, J. R.; Mangalam, S. M.

    1986-01-01

    Goertler vortices arise in boundary layers along concave surfaces due to centrifugal effects. This paper presents some results of an experiment conducted to study the development of these vortices on an airfoil with a pressure gradient in the concave region where an attached laminar boundary layer was insured with suction through a perforated panel. A sublimating chemical technique was used to visualize Goertler vortices and the velocity field was measured by laser velocimetry. Experimental disturbance functions are compared with those predicted by the linear stability theory. The trend of vortex amplification in the concave zone and damping in the following convex region is shown to essentially follow the theoretical predictions.

  10. The NASA Langley Laminar-Flow-Control Experiment on a Swept Supercritical Airfoil: Basic Results for Slotted Configuration

    NASA Technical Reports Server (NTRS)

    Harris, Charles D.; Brooks, Cuyler W., Jr.; Clukey, Patricia G.; Stack, John P.

    1989-01-01

    The effects of Mach number and Reynolds number on the experimental surface pressure distributions and transition patterns for a large chord, swept supercritical airfoil incorporating an active Laminar Flow Control suction system with spanwise slots are presented. The experiment was conducted in the Langley 8 foot Transonic Pressure Tunnel. Also included is a discussion of the influence of model/tunnel liner interactions on the airfoil pressure distribution. Mach number was varied from 0.40 to 0.82 at two chord Reynolds numbers, 10 and 20 x 1,000,000, and Reynolds number was varied from 10 to 20 x 1,000,000 at the design Mach number.

  11. 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.

  12. Intermittent Behavior of the Separated Boundary Layer along the Suction Surface of a Low Pressure Turbine Blade under Periodic Unsteady Flow Conditions

    NASA Technical Reports Server (NTRS)

    Oeztuerk, B; Schobeiri, M. T.; Ashpis, David E.

    2005-01-01

    The paper experimentally and theoretically studies the effects of periodic unsteady wake flow and aerodynamic characteristics on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experiments were carried out at Reynolds number of 110,000 (based on suction surface length and exit velocity). For one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, intermittency behaviors were experimentally and theoretically investigated. The current investigation attempts to extend the intermittency unsteady boundary layer transition model developed in previously to the LPT cases, where separation occurs on the suction surface at a low Reynolds number. The results of the unsteady boundary layer measurements and the intermittency analysis were presented in the ensemble-averaged and contour plot forms. The analysis of the boundary layer experimental data with the flow separation, confirms the universal character of the relative intermittency function which is described by a Gausssian function.

  13. 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.

  14. 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.

  15. 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.

  16. Recent advances in laser triangulation-based measurement of airfoil surfaces

    NASA Astrophysics Data System (ADS)

    Hageniers, Omer L.

    1995-01-01

    The measurement of aircraft jet engine turbine and compressor blades requires a high degree of accuracy. This paper will address the development and performance attributes of a noncontact electro-optical gaging system specifically designed to meet the airfoil dimensional measurement requirements inherent in turbine and compressor blade manufacture and repair. The system described consists of the following key components: a high accuracy, dual channel, laser based optical sensor, a four degree of freedom mechanical manipulator system and a computer based operator interface. Measurement modes of the system include point by point data gathering at rates up to 3 points per second and an 'on-the-fly' mode where points can be gathered at data rates up to 20 points per second at surface scanning speeds of up to 1 inch per second. Overall system accuracy is +/- 0.0005 inches in a configuration that is useable in the blade manufacturing area. The systems ability to input design data from CAD data bases and output measurement data in a CAD compatible data format is discussed.

  17. The Ultimate Flow Controlled Wind Turbine Blade Airfoil

    NASA Astrophysics Data System (ADS)

    Seifert, Avraham; Dolgopyat, Danny; Friedland, Ori; Shig, Lior

    2015-11-01

    Active flow control is being studied as an enabling technology to enhance and maintain high efficiency of wind turbine blades also with contaminated surface and unsteady winds as well as at off-design operating conditions. The study is focused on a 25% thick airfoil (DU91-W2-250) suitable for the mid blade radius location. Initially a clean airfoil was fabricated and tested, as well as compared to XFoil predictions. From these experiments, the evolution of the separation location was identified. Five locations for installing active flow control actuators are available on this airfoil. It uses both Piezo fluidic (``Synthetic jets'') and the Suction and Oscillatory Blowing (SaOB) actuators. Then we evaluate both actuation concepts overall energy efficiency and efficacy in controlling boundary layer separation. Since efficient actuation is to be found at low amplitudes when placed close to separation location, distributed actuation is used. Following the completion of the baseline studies the study has focused on the airfoil instrumentation and extensive wind tunnel testing over a Reynolds number range of 0.2 to 1.5 Million. Sample results will be presented and outline for continued study will be discussed.

  18. Shock Wave/Stable Vortex Interaction over A NACA 0012 Airfoil: A Numerical Study

    NASA Astrophysics Data System (ADS)

    Alammar, Khalid

    2002-11-01

    While many studies have been conducted on shock wave/vortex interaction in general, not much attention has been given to shock wave/stable vortex interaction over airfoils or wings, and the affect of vortices on transonic airfoil performance. This work is intended to numerically investigate shock wave/stable vortex interaction over airfoils, and to quantify vortex affect on airfoil performance at transonic speeds. To accomplish the objective, a steady, transonic turbulent flow around a 0.5-m NACA 0012 airfoil at alpha = 1 degree was simulated. The simulation was carried out using one, three, and no vortices. The stable vortices were placed on the suction side using cavities (dimples of 15-mm diameter). The simulation was conducted using the commercial code "Fluent". The second-order, coupled solver was invoked. Spalart-Almaras model was used in the formulation. The ideal-gas model and Sutherland's law were used for density and viscosity calculations, respectively. The computation was carried out at Mach 0.8 and Reynolds number of 9.1x106. Due to geometric complexity of the dimples, an unstructured mesh was used. The commercial code "Gambit" was utilized to construct the mesh. Three mesh blocks were generated to accommodate the boundary layer, the wake region, and the remainder of the computational space. A 3-mm, 20-layer boundary layer was constructed, and the first row was 0.01-mm high. The mesh consisted of 156,000 cells (tetrahedral for the domain and wedges for the boundary layer). Grid independence was checked by doubling the number of cells around the airfoil and in the wake region. No significant changes in the results were observed. The far field was 20 chords away from the surface. The simulation revealed the stable vortical flow structure inside the dimples. Small separation and reattachment was Predicted in all cases. It was found that the shock wave on the suction side of the airfoil was pushed up-stream by the stable vortices. Three vortices induced

  19. Instability and Transition of Flow at, and Near, an Attachment-line - Including Control by Surface Suction

    NASA Technical Reports Server (NTRS)

    Smith, A.

    1996-01-01

    Advances in aviation during and following the Second World War led to an enormous improvement in the performance of aircraft. The push for enhanced efficiency brought cruise speeds into the transonic range, where the associated drag rise due to the appearance of shock-waves became a limiting factor. Wing sweep was adopted to delay the onset of this drag rise, but with this development came several new and unforeseen problems. Preliminary theoretical work assumed that the boundary layer transition characteristics of a swept wing would be subject to the independence principle, so the chordwise transition position could be predicted from two-dimensional work Gas turbine development has now reached a point where additional increases in efficiency are both difficult and expensive to achieve. Consequently, aircraft manufacturers are looking elsewhere for ways to reduce Direct Operating Costs (DOC's) or increase military performance. The attention of industry is currently focusing on Hybrid Laminar Flow Control (HLFC) as a possible method of reducing DOC's for civil aircraft. Following this study and discussions with NASA Langley and Boeing a different series of questions have been addressed in the present work. There are five areas of interest: Relaminarisation of the attachment-line boundary layer when the value of R exceeds 600. The effects of large suction levels on transition in the attachment-line boundary layer (ie critical oversuction). The transition characteristics of a relaminarised attachment-line flow which encounters a non-porous surface. The effect of attachment-line suction on the spanwise propagation of gross disturbances emanating from the wing-fuselage junction. The attachment-line transition caused by surface blowing.

  20. A Feasibility Study to Control Airfoil Shape Using THUNDER

    NASA Technical Reports Server (NTRS)

    Pinkerton, Jennifer L.; Moses, Robert W.

    1997-01-01

    The objective of this study was to assess the capabilities of a new out-of-plane displacement piezoelectric actuator called thin-layer composite-unimorph ferroelectric driver and sensor (THUNDER) to alter the upper surface geometry of a subscale airfoil to enhance performance under aerodynamic loading. Sixty test conditions, consisting of combinations of five angles of attack, four dc applied voltages, and three tunnel velocities, were studied in a tabletop wind tunnel. Results indicated that larger magnitudes of applied voltage produced larger wafer displacements. Wind-off displacements were also consistently larger than wind-on. Higher velocities produced larger displacements than lower velocities because of increased upper surface suction. Increased suction also resulted in larger displacements at higher angles of attack. Creep and hysteresis of the wafer, which were identified at each test condition, contributed to larger negative displacements for all negative applied voltages and larger positive displacements for the smaller positive applied voltage (+102 V). An elastic membrane used to hold the wafer to the upper surface hindered displacements at the larger positive applied voltage (+170 V). Both creep and hysteresis appeared bounded based on the analysis of several displacement cycles. These results show that THUNDER can be used to alter the camber of a small airfoil under aerodynamic loads.

  1. 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.

  2. 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.

  3. Control of unsteady separated flow associated with the dynamic stall of airfoils

    NASA Technical Reports Server (NTRS)

    Wilder, M. C.

    1995-01-01

    An effort to understand and control the unsteady separated flow associated with the dynamic stall of airfoils was funded for three years through the NASA cooperative agreement program. As part of this effort a substantial data base was compiled detailing the effects various parameters have on the development of the dynamic stall flow field. Parameters studied include Mach number, pitch rate, and pitch history, as well as Reynolds number (through two different model chord lengths) and the condition of the boundary layer at the leading edge of the airfoil (through application of surface roughness). It was found for free stream Mach numbers as low as 0.4 that a region of supersonic flow forms on the leading edge of the suction surface of the airfoil at moderate angles of attack. The shocks which form in this supersonic region induce boundary-layer separation and advance the dynamic stall process. Under such conditions a supercritical airfoil profile is called for to produce a flow field having a weaker leading-edge pressure gradient and no leading-edge shocks. An airfoil having an adaptive-geometry, or dynamically deformable leading edge (DDLE), is under development as a unique active flow-control device. The DDLE, formed of carbon-fiber composite and fiberglass, can be flexed between a NACA 0012 profile and a supercritical profile in a controllable fashion while the airfoil is executing an angle-of-attack pitch-up maneuver. The dynamic stall data were recorded using point diffraction interferometry (PDI), a noninvasive measurement technique. A new high-speed cinematography system was developed for recording interferometric images. The system is capable of phase-locking with the pitching airfoil motion for real-time documentation of the development of the dynamic stall flow field. Computer-aided image analysis algorithms were developed for fast and accurate reduction of the images, improving interpretation of the results.

  4. Experimental measurement of the aerodynamic charateristics of two-dimensional airfoils for an unmanned aerial vehicle

    NASA Astrophysics Data System (ADS)

    Velazquez, Luis; Nožička, Jiří; Vavřín, Jan

    2012-04-01

    This paper is part of the development of an airfoil for an unmanned aerial vehicle (UAV) with internal propulsion system; the investigation involves the analysis of the aerodynamic performance for the gliding condition of two-dimensional airfoil models which have been tested. This development is based on the modification of a selected airfoil from the NACA four digits family. The modification of this base airfoil was made in order to create a blowing outlet with the shape of a step on the suction surface since the UAV will have an internal propulsion system. This analysis involved obtaining the lift, drag and pitching moment coefficients experimentally for the situation where there is not flow through the blowing outlet, called the no blowing condition by means of wind tunnel tests. The methodology to obtain the forces experimentally was through an aerodynamic wire balance. Obtained results were compared with numerical results by means of computational fluid dynamics (CFD) from references and found in very good agreement. Finally, a selection of the airfoil with the best aerodynamic performance is done and proposed for further analysis including the blowing condition.

  5. Virtual Shaping of a Two-dimensional NACA 0015 Airfoil Using Synthetic Jet Actuator

    NASA Technical Reports Server (NTRS)

    Chen, Fang-Jenq; Beeler, George B.

    2002-01-01

    The Aircraft Morphing Program at NASA Langley envisions an aircraft without conventional control surfaces. Instead of moving control surfaces, the vehicle control systems may be implemented with a combination of propulsive forces, micro surface effectors, and fluidic devices dynamically operated by an intelligent flight control system to provide aircraft maneuverability over each mission segment. As a part of this program, a two-dimensional NACA 0015 airfoil model was designed to test mild maneuvering capability of synthetic jets in a subsonic wind tunnel. The objective of the experiments is to assess the applicability of using unsteady suction and blowing to alter the aerodynamic shape of an airfoil with a purpose to enhance lift and/or to reduce drag. Synthetic jet actuation at different chordwise locations, different forcing frequencies and amplitudes, under different freestream velocities are investigated. The effect of virtual shape change is indicated by a localized increase of surface pressure in the neighborhood of synthetic jet actuation. That causes a negative lift to the airfoil with an upper surface actuation. When actuation is applied near the airfoil leading edge, it appears that the stagnation line is shifted inducing an effect similar to that caused by a small angle of attack to produce an overall lift change.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. Control of Pitching Airfoil Aerodynamics by Vorticity Flux Modification using Active Bleed

    NASA Astrophysics Data System (ADS)

    Kearney, John; Glezer, Ari

    2014-11-01

    Distributed active bleed driven by pressure differences across a pitching airfoil is used to regulate the vorticity flux over the airfoil's surface and thereby to control aerodynamic loads in wind tunnel experiments. The range of pitch angles is varied beyond the static stall margin of the 2-D VR-7 airfoil at reduced pitching rates up to k = 0.42. Bleed is regulated dynamically using piezoelectric louvers between the model's pressure side near the trailing edge and the suction surface near the leading edge. The time-dependent evolution of vorticity concentrations over the airfoil and in the wake during the pitch cycle is investigated using high-speed PIV and the aerodynamic forces and moments are measured using integrated load cells. The timing of the dynamic stall vorticity flux into the near wake and its effect on the flow field are analyzed in the presence and absence of bleed using proper orthogonal decomposition (POD). It is shown that bleed actuation alters the production, accumulation, and advection of vorticity concentrations near the surface with significant effects on the evolution, and, in particular, the timing of dynamic stall vortices. These changes are manifested by alteration of the lift hysteresis and improvement of pitch stability during the cycle, while maintaining cycle-averaged lift to within 5% of the base flow level with significant implications for improvement of the stability of flexible wings and rotor blades. This work is supported by the Rotorcraft Center (VLRCOE) at Georgia Tech.

  11. 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.

  12. A Quantitative Investigation of Surface Roughness Effects on Airfoil Boundary Layer Transition Using Infrared Thermography

    NASA Astrophysics Data System (ADS)

    Beeby, Todd Daniel

    An investigation of the impact of subcritical leading edge distributed roughness elements on airfoil boundary layer transition location has been undertaken using infrared thermography. In particular, a quantitative approach to boundary layer transition location detection using a differential energy balance method was implemented using a heating pad to produce constant heat flux. This was performed on a S809 airfoil model at Re c = 0.75 and 1.0 x 106, using roughness elements of height k/c = 3.75, 4.25 and 5.00 x 10 --4, pattern densities of 2 to 10 %, and roughness locations of 1 to 6 % chord. Turbulator tape of height k/c = 6.67 x 10--4 was also examined. Results indicate significant impact on transition for all roughness cases, and a more pronounced influence of roughness density as compared to roughness element height. The phenomenon of early laminar bubble collapse was also found to occur for some roughness configurations. The quantitative method used was found to be an effective means for automated transition location determination.

  13. Design philosophy of long range LFC transports with advanced supercritical LFC airfoils. [laminar flow control

    NASA Technical Reports Server (NTRS)

    Pfenninger, Werner; Vemuru, Chandra S.

    1988-01-01

    The achievement of 70 percent laminar flow using modest boundary layer suction on the wings, empennage, nacelles, and struts of long-range LFC transports, combined with larger wing spans and lower span loadings, could make possible an unrefuelled range halfway around the world up to near sonic cruise speeds with large payloads. It is shown that supercritical LFC airfoils with undercut front and rear lower surfaces, an upper surface static pressure coefficient distribution with an extensive low supersonic flat rooftop, a far upstream supersonic pressure minimum, and a steep subsonic rear pressure rise with suction or a slotted cruise flap could alleviate sweep-induced crossflow and attachment-line boundary-layer instability. Wing-mounted superfans can reduce fuel consumption and engine tone noise.

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

    NASA Astrophysics Data System (ADS)

    Phan, Nhan Huu

    input. The CFD results show that airfoil circulation control is achieved by the varying the CFF intake flow rate and the momentum of the CFF exhaust jet (e.g. through airfoil AoA or fan rotational speed). The presence of the CFF has the effect of moving the stagnation point on the airfoil pressure surface from the CFF airfoil LE region near the CFF to as far back as the airfoil trailing edge. At high AoA operation, LE flow separation on the airfoil suction surface is delayed by flow entrainment of the high-energy jet leaving the CFF. Detailed analysis of the flow field through the crossflow fan and its housing were carried out to understand its fluid-dynamics behavior, and it is found that the airfoil geometry acts as inlet guide vanes to the crossflow fan as the angle-of-attack is varied, thus introducing pre-swirl or co-swirl into the first stage of the crossflow fan. An experimental study of the LEEF concept confirmed that the concept works and it is robust. Finally, as application examples, the LEEF technology is applied to a Remote Control model and to a generic tiltrotor aircraft similar in characteristics to DARPA's Aerial Reconfigurable Embedded System. These aircraft configurations were analyzed using 2D and 3D CFD.

  15. 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.

  16. 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.

  17. 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.

  18. 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

    sequentially, which is the condition for a ladder jump to occur. Lower amplification factor A for the T-S waves, which can result in a radiation of lower noise levels for the broadband hump peak at fs. This phenomenon will proportionally reduce the noise level difference for fn and fn+1, thus making an identification of a ladder jump event more difficult. Finally, we believe that the tone noise generated in this experimental study is of genuine tones of an isolated airfoil. This can be supported by the fact that, when considering either a straight trailing edge or a serrated trailing edge, the overall airfoil geometry at the same angle of attack is still retained and the wind tunnel setup and locations of the laboratory equipment, which could potentially become an anchor point for the acoustic feedback loop, are exactly the same. However, the straight S0 and serrated S2" trailing edges have been shown to produce systematically different spectral characteristics, especially in the forms of tonal rungs, which can be predicted accurately by the original and slightly modified acoustic feedback Model A respectively. In summary, the trailing edge serration is a useful device for the suppression of airfoil instability self-noise. For greater control effectiveness, however, the laminar separation bubble must be situated within the serration region of the trailing edge. This connection imposes restrictions on the angle of attack and velocity over which trailing edge serrations are effective. The feedback loop structure about the wake noise source and the suction surface of the airfoil in Model B is ignored in the present case. This assumption should be reasonably valid given that, in our previous study [3], we cannot identify any significant role of the boundary layer flow at the suction surface in contributing the instability tonal noise radiation across a wide range of Reynolds numbers.

  19. Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Intermittency Behavior Along the Suction Surface of a Low Pressure Turbine Blade

    NASA Technical Reports Server (NTRS)

    Schobeiri, M. T.; Ozturk, B.; Ashpis, David E.

    2007-01-01

    The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

  20. Surface Modeling and Grid Generation for Iced Airfoils (SmaggIce)

    NASA Technical Reports Server (NTRS)

    Hammond, Brandy M.

    2004-01-01

    Many of the troubles associated with problem solving are alleviated when there is a model that can be used to represent the problem. Through the Advanced Graphics and Visualization (G-VIS) Laboratory and other facilities located within the Research Analysis Center, the Computer Services Division (CSD) is able to develop and maintain programs and software that allow for the modeling of various situations. For example, the Icing Research Branch is devoted to investigating the effect of ice that forms on the wings and other airfoils of airplanes while in flight. While running tests that physically generate ice and wind on airfoils within the laboratories and wind tunnels on site are done, it would be beneficial if most of the preliminary work could be done outside of the lab. Therefore, individuals from within CSD have collaborated with Icing Research in order to create SmaggIce. This software allows users to create ice patterns on clean airfoils or open files containing a variety of icing situations, manipulate and measure these forms, generate, divide, and merge grids around these elements for more explicit analysis, and specify and rediscretize subcurves. With the projected completion date of Summer 2005, the majority of the focus of the Smagglce team is user-functionality and error handling. My primary responsibility is to test the Graphical User Interface (GUI) in SmaggIce in order to ensure the usability and verify the expected results of the events (buttons, menus, etc.) within the program. However, there is no standardized, systematic way in which to test all the possible combinations or permutations of events, not to mention unsolicited events such as errors. Moreover, scripting tests, if not done properly and with a view towards inevitable revision, can result in more apparent errors within the software and in effect become useless whenever the developers of the program make a slight change in the way a specific process is executed. My task therefore

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. Cold-air aerodynamic study in a two-dimensional cascade of a turbine stator blade with suction-surface film cooling

    NASA Technical Reports Server (NTRS)

    Brown, D. B.; Helon, R. M.

    1973-01-01

    The effect on aerodynamic performance of a single row of spanwise-spaced coolant holes located at four positions chordwise along the suction surface was investigated. In addition, multiple-row data were obtained. The data are presented in terms of primary efficiency, coolant-to-primary-flow percentage, and coolant pressure ratio. Primary efficiencies of multirow blade configurations compare satisfactorily with that calculated from the single-row efficiency increments. At any given coolant flow percentage, the efficiency was about the same for all row locations. For a given coolant pressure, the efficiency varied by as much as 1 percent, depending on the row location.

  7. 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.

  8. Flatback airfoil wind tunnel experiment.

    SciTech Connect

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

    2008-04-01

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

  9. Numerical Studies on a Rotor with Distributed Suction for Noise Reduction

    NASA Astrophysics Data System (ADS)

    Lutz, Thorsten; Arnold, Benjamin; Wolf, Alexander; Krämer, Ewald

    2014-06-01

    Minimizing the flow-induced noise is an important issue in the design of modern onshore wind turbines. There is a number of proven passive means to reduce the aeroacoustic noise, such as the implementation of serrations, porous trailing edges or the aeroacoustic airfoil design. The noise emission can be further reduced by active flow control techniques. In the present study the impact of distributed boundary layer suction on the noise emission of an airfoil and a complete rotor is investigated. Aerodynamic and aeroacoustic wind tunnel tests were performed for the NACA 64-418 airfoil and supplemented by numerical calculations. The aeroacoustic analyses have been conducted by means of the institute's Rnoise prediction scheme. The 2D studies have shown that noise reductions of 5 dB can be achieved by suction at moderate mass flow rates. To study the impact of three-dimensional effects numerical investigations have been conducted on the example of the generic NREL 5MW rotor with suction applied in the outer part of the blade. The predictions for the complete rotor provided smaller benefits compared to those for the isolated airfoil, mainly because the examined suction configurations were not optimized with respect to the extent of the suction patch and suction distribution.

  10. 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.

  11. Simulated-airline-service flight tests of laminar-flow control with perforated-surface suction system

    NASA Technical Reports Server (NTRS)

    Maddalon, Dal V.; Braslow, Albert L.

    1990-01-01

    The effectiveness and practicality of candidate leading edge systems for suction laminar flow control transport airplanes were investigated in a flight test program utilizing a modified JetStar airplane. The leading edge region imposes the most severe conditions on systems required for any type of laminar flow control. Tests of the leading edge systems, therefore, provided definitive results as to the feasibility of active laminar flow control on airplanes. The test airplane was operated under commercial transport operating procedures from various commercial airports and at various seasons of the year.

  12. 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.

  13. 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.

  14. 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.

  15. Use of a liquid-crystal, heater-element composite for quantitative, high-resolution heat transfer coefficients on a turbine airfoil, including turbulence and surface roughness effects

    NASA Astrophysics Data System (ADS)

    Hippensteele, Steven A.; Russell, Louis M.; Torres, Felix J.

    1987-05-01

    Local heat transfer coefficients were measured along the midchord of a three-times-size turbine vane airfoil in a static cascade operated at roon temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a Mylar sheet with a layer of cholestric liquid crystals, which change color with temperature, and a heater made of a polyester sheet coated with vapor-deposited gold, which produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. Tests were conducted at two free-stream turbulence intensities: 0.6 percent, which is typical of wind tunnels; and 10 percent, which is typical of real engine conditions. In addition to a smooth airfoil, the effects of local leading-edge sand roughness were also examined for a value greater than the critical roughness. The local heat transfer coefficients are presented for both free-stream turbulence intensities for inlet Reynolds numbers from 1.20 to 5.55 x 10 to the 5th power. Comparisons are also made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code.

  16. Use of a liquid-crystal, heater-element composite for quantitative, high-resolution heat transfer coefficients on a turbine airfoil, including turbulence and surface roughness effects

    NASA Technical Reports Server (NTRS)

    Hippensteele, Steven A.; Russell, Louis M.; Torres, Felix J.

    1987-01-01

    Local heat transfer coefficients were measured along the midchord of a three-times-size turbine vane airfoil in a static cascade operated at roon temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a Mylar sheet with a layer of cholestric liquid crystals, which change color with temperature, and a heater made of a polyester sheet coated with vapor-deposited gold, which produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. Tests were conducted at two free-stream turbulence intensities: 0.6 percent, which is typical of wind tunnels; and 10 percent, which is typical of real engine conditions. In addition to a smooth airfoil, the effects of local leading-edge sand roughness were also examined for a value greater than the critical roughness. The local heat transfer coefficients are presented for both free-stream turbulence intensities for inlet Reynolds numbers from 1.20 to 5.55 x 10 to the 5th power. Comparisons are also made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code.

  17. Hydroponic Feed With Suction

    NASA Technical Reports Server (NTRS)

    Cox, William M.; Brown, Christopher S.; Dreschel, Thomas W.

    1994-01-01

    Placing nutrient solution under suction increases growth. Foam plug seals growing stem of plant, making it possible to maintain suction in nutrient liquid around roots. Jar wrapped in black tape to keep out light. Potential use in terrestrial applications in arid climates or in labor-intensive agricultural situations.

  18. Modeling an increase in the lift and aerodynamic efficiency of a thick Göttingen airfoil with optimum arrangement

    NASA Astrophysics Data System (ADS)

    Isaev, S. A.; Sudakov, A. G.; Usachov, A. E.; Kharchenko, V. B.

    2015-06-01

    The Reynolds equations closed using the Menter shear-stress-transfer model modified with allowance for the curvature of flow line have been numerically solved jointly with the energy equation. The obtained solution has been used to calculate subsonic flow (at M = 0.05 and 5° angle of attack) past a thick (24% chord) Göttingen airfoil with variable arrangement of a small-sized (about 10% chord) circular vortex cell with fixed distributed suction Cq = 0.007 from the surface of a central body. It is established that the optimum arrangement of the vortex cell provides a twofold decrease in the bow drag coefficient Cx, a threefold increase in the lift coefficient Cy, and an about fivefold increase in the aerodynamic efficiency at Re = 105 in comparison to the smooth airfoil.

  19. Flow Control on Low-Pressure Turbine Airfoils Using Vortex Generator Jets

    NASA Technical Reports Server (NTRS)

    Volino, Ralph J.; Ibrahim, Mounir B.; Kartuzova, Olga

    2010-01-01

    Motivation - Higher loading on Low-Pressure Turbine (LPT) airfoils: Reduce airfoil count, weight, cost. Increase efficiency, and Limited by suction side separation. Growing understanding of transition, separation, wake effects: Improved models. Take advantage of wakes. Higher lift airfoils in use. Further loading increases may require flow control: Passive: trips, dimples, etc. Active: plasma actuators, vortex generator jets (VGJs). Can increased loading offset higher losses on high lift airfoils. Objectives: Advance knowledge of boundary layer separation and transition under LPT conditions. Demonstrate, improve understanding of separation control with pulsed VGJs. Produce detailed experimental data base. Test and develop computational models.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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

    nearly remain unchanged. In dynamic stall, leading edge defect imposes a greater influence on the aerodynamic characteristics of airfoil than steady conditions. By increasing in defect length, it is found that the separated area becomes more intense and moves forward along the suction surface. Conclusions Leading edge defect has significant influence on the aerodynamic and flow characteristics of the airfoil, which will reach a stable status with enough large defect size. The leading edge separation bubble, circulation in the defect cavity and intense tailing edge vortex are the main features of flow around defective airfoils. PMID:27658310

  6. 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.

  7. 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.

  8. Four-wall turbine airfoil with thermal strain control for reduced cycle fatigue

    SciTech Connect

    Cambell, Christian X

    2013-09-17

    A turbine airfoil (20B) with a thermal expansion control mechanism that increases the airfoil camber (60, 61) under operational heating. The airfoil has four-wall geometry, including pressure side outer and inner walls (26, 28B), and suction side outer and inner walls (32, 34B). It has near-wall cooling channels (31F, 31A, 33F, 33A) between the outer and inner walls. A cooling fluid flow pattern (50C, 50W, 50H) in the airfoil causes the pressure side inner wall (28B) to increase in curvature under operational heating. The pressure side inner wall (28B) is thicker than walls (26, 34B) that oppose it in camber deformation, so it dominates them in collaboration with the suction side outer wall (32), and the airfoil camber increases. This reduces and relocates a maximum stress area (47) from the suction side outer wall (32) to the suction side inner wall (34B, 72) and the pressure side outer wall (26).

  9. 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.

  10. 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.

  11. Combined Influence of Hall Current and Soret Effect on Chemically Reacting Magnetomicropolar Fluid Flow from Radiative Rotating Vertical Surface with Variable Suction in Slip-Flow Regime

    PubMed Central

    Jain, Preeti

    2014-01-01

    An analysis study is presented to study the effects of Hall current and Soret effect on unsteady hydromagnetic natural convection of a micropolar fluid in a rotating frame of reference with slip-flow regime. A uniform magnetic field acts perpendicularly to the porous surface which absorbs the micropolar fluid with variable suction velocity. The effects of heat absorption, chemical reaction, and thermal radiation are discussed and for this Rosseland approximation is used to describe the radiative heat flux in energy equation. The entire system rotates with uniform angular velocity Ω about an axis normal to the plate. The nonlinear coupled partial differential equations are solved by perturbation techniques. In order to get physical insight, the numerical results of translational velocity, microrotation, fluid temperature, and species concentration for different physical parameters entering into the analysis are discussed and explained graphically. Also, the results of the skin-friction coefficient, the couple stress coefficient, Nusselt number, and Sherwood number are discussed with the help of figures for various values of flow pertinent flow parameters. PMID:27350957

  12. 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.

  13. 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).

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Intraoral transmental suction lipectomy.

    PubMed

    Mommaerts, M Y; Abeloos, J V S; De Clerq, C A S; Neyt, L F

    2002-08-01

    Intraoral transmental suction lipectomy (TMSL) is performed by entering the suction canula through the chin osteotomy/ostectomy gap into the sub- and/or supraplatysmal fat tissue layers. The aim of the study was to know patients' and operators' satisfaction with the procedure, and to know the kind and frequency of the complications. Twenty patients were consecutively treated and reviewed after a minimum of 5 years. All were satisfied with the overall results. It proved difficult to differentiate between the results of the liposuction and those of the genioplasty and/or orthognathic profile correction. From a surgeon's point of view, 11 showed excellent, nine good and one moderate results. Complications included one local subcutaneous infection, four transient neurosensory disturbances at the lower lip and two marginal branch weaknesses. All complications were resolved by the time of the long-term follow-up appointment. TMSL offers the psychological advantage of being performed without skin incision. Cosmetic results and complications are similar to those obtained with the transcutaneous liposuction techniques. PMID:12361067

  1. About the effects of an oscillating miniflap upon the wake on an airfoil, all immersed in turbulent flow

    NASA Astrophysics Data System (ADS)

    S, Delnero J.; J, Marañón Di Leo; Colman; J; M, Camocardi; Sainz M, García; F, Muñoz

    2011-12-01

    The present research analyzes the asymmetry in the rolling up shear layers behind the blunt trailing edge of an airfoil 4412 with a miniflap acting as active flow control device and its wake organization. Experimental investigations relating the asymmetry of the vortex flow in the near wake region, able to distort the flow increasing the downwash of an airfoil, have been performed. All of these in a free upstream turbulent flow (1.8% intensity). We examine the near wake region characteristics of a wing model with a 4412 airfoil without and with a rotating miniflap located on the lower surface, near the trailing edge. The flow in the near wake, for 3 x-positions (along chord line) and 20 vertical points in each x-position, was explored, for three different rotating frequencies, in order to identify signs of asymmetry of the initial counter rotating vortex structures. Experimental evidence is presented showing that for typical lifting conditions the shear layer rollup process within the near wake is different for the upper and lower vortices: the shear layer separating from the pressure side of the airfoil begins its rollup immediately behind the trailing edge, creating a stronger vortex while the shear layer from the suction side begins its rollup more downstream creating a weaker vortex. The experimental data were processed by classical statistics methods. Aspects of a mechanism connecting the different evolution and pattern of these initial vortex structures with lift changes and wake alleviating processes, due to these miniflaps, will be studied in future works.

  2. Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Re-attachment along the Suction Surface of a Low Pressure Turbine Blade

    NASA Technical Reports Server (NTRS)

    Ozturk, B.; Schobeiri, M. T.; Ashpis, David E.

    2005-01-01

    The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. Dynamical effects of suction/heating on turbulent boundary layers

    NASA Astrophysics Data System (ADS)

    Blackwelder, Ron

    1992-01-01

    The main emphasis of this quarter has been to test the effects of suction in a controlled environment with the emulated wall eddy structure. A study of the curved working wall of the test section in the Goertler Wind Tunnel showed that there were sufficient stresses within the plexiglas that cutting elongated slits for suction would probably cause the surface to develop step-type roughnesses. Thus several individual holes were initially drilled along the streamline direction in a spanwise region between two vortices. Air was withdrawn through this series of holes to provide a semi-continuous region of suction. Differing rates of suction through these holes were used to explore the effects upon the eddy structure. These preliminary results were obtained using visualization; i.e., smoke as introduced via a smoke wire into the boundary layer. Images were captured using a video camera and analyzed to determine the best suction rates. The preliminary results showed that suction has a large effect upon individual streaks of low speed fluid. Without the suction, the low speed region lying in the upwelling zone between two streamwise vortices was broken down by a secondary instability. This instability typically caused the low speed fluid marked with the smoke to oscillate from side to side in a manifestation of an inflectional instability in the spanwise direction as found and reported earlier in this research. With increasing distance downstream, the oscillation amplitude grew very rapidly until it broke down into complete turbulence.

  8. 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.

  9. 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.

  10. Design optimization of transonic airfoils

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. Use of a liquid-crystal and heater-element composite for quantitative, high-resolution heat-transfer coefficients on a turbine airfoil including turbulence and surface-roughness effects

    NASA Astrophysics Data System (ADS)

    Hippensteele, S. A.; Russell, L. M.; Torres, F. J.

    Local heat transfer coefficients were measured along the midchord of a three-times-size turbine vane airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a Mylar sheet with a layer of cholestric liquid crystals, which change color with temperature, and a heater made of a polyester sheet coated with vapor-deposited gold, which produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. Tests were conducted at two free-stream turbulence intensities: 0.6 percent, which is typical of wind tunnels; and 10 percent, which is typical of real engine conditions. In addition to a smooth airfoil, the effects of local leading-edge sand roughness were also examined for a value greater than the critical roughness. The local heat transfer coefficients are presented for both free-stream turbulence intensities for inlet Reynolds numbers from 1.20 to 5.55 x 10 to the 5th power. Comparisons are also made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code.

  15. Use of a liquid-crystal and heater-element composite for quantitative, high-resolution heat-transfer coefficients on a turbine airfoil including turbulence and surface-roughness effects

    NASA Technical Reports Server (NTRS)

    Hippensteele, S. A.; Russell, L. M.; Torres, F. J.

    1987-01-01

    Local heat transfer coefficients were measured along the midchord of a three-times-size turbine vane airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a Mylar sheet with a layer of cholestric liquid crystals, which change color with temperature, and a heater made of a polyester sheet coated with vapor-deposited gold, which produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. Tests were conducted at two free-stream turbulence intensities: 0.6 percent, which is typical of wind tunnels; and 10 percent, which is typical of real engine conditions. In addition to a smooth airfoil, the effects of local leading-edge sand roughness were also examined for a value greater than the critical roughness. The local heat transfer coefficients are presented for both free-stream turbulence intensities for inlet Reynolds numbers from 1.20 to 5.55 x 10 to the 5th power. Comparisons are also made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code.

  16. 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.

  17. 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.

  18. 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.

  19. Suction muffler for refrigeration compressor

    DOEpatents

    Nelson, Richard T.; Middleton, Marc G.

    1983-01-01

    A hermetic refrigeration compressor includes a suction muffler formed from two pieces of plastic material mounted on the cylinder housing. One piece is cylindrical in shape with an end wall having an aperture for receiving a suction tube connected to the cylinder head. The other piece fits over and covers the other end of the cylindrical piece, and includes a flaring entrance horn which extends toward the return line on the sidewall of the compressor shell.

  20. Suction muffler for refrigeration compressor

    DOEpatents

    Nelson, R.T.; Middleton, M.G.

    1983-01-25

    A hermetic refrigeration compressor includes a suction muffler formed from two pieces of plastic material mounted on the cylinder housing. One piece is cylindrical in shape with an end wall having an aperture for receiving a suction tube connected to the cylinder head. The other piece fits over and covers the other end of the cylindrical piece, and includes a flaring entrance horn which extends toward the return line on the sidewall of the compressor shell. 5 figs.

  1. Airfoil family design for large offshore wind turbine blades

    NASA Astrophysics Data System (ADS)

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

    2014-06-01

    Wind turbine blades size has scaled-up during last years due to wind turbine platform increase especially for offshore applications. The EOLIA project 2007-2010 (Spanish Goverment funded project) was focused on the design of large offshore wind turbines for deep waters. The project was managed by ACCIONA Energia and the wind turbine technology was designed by ACCIONA Windpower. The project included the design of a wind turbine airfoil family especially conceived for large offshore wind turbine blades, in the order of 5MW machine. Large offshore wind turbines suffer high extreme loads due to their size, in addition the lack of noise restrictions allow higher tip speeds. Consequently, the airfoils presented in this work are designed for high Reynolds numbers with the main goal of reducing blade loads and mantainig power production. The new airfoil family was designed in collaboration with CENER (Spanish National Renewable Energy Centre). The airfoil family was designed using a evolutionary algorithm based optimization tool with different objectives, both aerodynamic and structural, coupled with an airfoil geometry generation tool. Force coefficients of the designed airfoil were obtained using the panel code XFOIL in which the boundary layer/inviscid flow coupling is ineracted via surface transpiration model. The desing methodology includes a novel technique to define the objective functions based on normalizing the functions using weight parameters created from data of airfoils used as reference. Four airfoils have been designed, here three of them will be presented, with relative thickness of 18%, 21%, 25%, which have been verified with the in-house CFD code, Wind Multi Block WMB, and later validated with wind tunnel experiments. Some of the objectives for the designed airfoils concern the aerodynamic behavior (high efficiency and lift, high tangential coefficient, insensitivity to rough conditions, etc.), others concern the geometry (good for structural design

  2. Theoretical effect of modifications to the upper surface of two NACA airfoils using smooth polynomial additional thickness distributions which emphasize leading edge profile and which vary quadratically at the trailing edge. [using flow equations and a CDC 7600 computer

    NASA Technical Reports Server (NTRS)

    Merz, A. W.; Hague, D. S.

    1975-01-01

    An investigation was conducted on a CDC 7600 digital computer to determine the effects of additional thickness distributions to the upper surface of the NACA 64-206 and 64 sub 1 - 212 airfoils. The additional thickness distribution had the form of a continuous mathematical function which disappears at both the leading edge and the trailing edge. The function behaves as a polynomial of order epsilon sub 1 at the leading edge, and a polynomial of order epsilon sub 2 at the trailing edge. Epsilon sub 2 is a constant and epsilon sub 1 is varied over a range of practical interest. The magnitude of the additional thickness, y, is a second input parameter, and the effect of varying epsilon sub 1 and y on the aerodynamic performance of the airfoil was investigated. Results were obtained at a Mach number of 0.2 with an angle-of-attack of 6 degrees on the basic airfoils, and all calculations employ the full potential flow equations for two dimensional flow. The relaxation method of Jameson was employed for solution of the potential flow equations.

  3. 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.

  4. 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.

  5. 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.

  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. 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.

  8. 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.

  9. Octopus-like suction cups: from natural to artificial solutions.

    PubMed

    Tramacere, F; Follador, M; Pugno, N M; Mazzolai, B

    2015-05-13

    Octopus suckers are able to attach to all nonporous surfaces and generate a very strong attachment force. The well-known attachment features of this animal result from the softness of the sucker tissues and the surface morphology of the portion of the sucker that is in contact with objects or substrates. Unlike artificial suction cups, octopus suckers are characterized by a series of radial grooves that increase the area subjected to pressure reduction during attachment. In this study, we constructed artificial suction cups with different surface geometries and tested their attachment performances using a pull-off setup. First, smooth suction cups were obtained for casting; then, sucker surfaces were engraved with a laser cutter. As expected, for all the tested cases, the engraving treatment enhanced the attachment performance of the elastomeric suction cups compared with that of the smooth versions. Moreover, the results indicated that the surface geometry with the best attachment performance was the geometry most similar to octopus sucker morphology. The results obtained in this work can be utilized to design artificial suction cups with higher wet attachment performance.

  10. Octopus-like suction cups: from natural to artificial solutions.

    PubMed

    Tramacere, F; Follador, M; Pugno, N M; Mazzolai, B

    2015-06-01

    Octopus suckers are able to attach to all nonporous surfaces and generate a very strong attachment force. The well-known attachment features of this animal result from the softness of the sucker tissues and the surface morphology of the portion of the sucker that is in contact with objects or substrates. Unlike artificial suction cups, octopus suckers are characterized by a series of radial grooves that increase the area subjected to pressure reduction during attachment. In this study, we constructed artificial suction cups with different surface geometries and tested their attachment performances using a pull-off setup. First, smooth suction cups were obtained for casting; then, sucker surfaces were engraved with a laser cutter. As expected, for all the tested cases, the engraving treatment enhanced the attachment performance of the elastomeric suction cups compared with that of the smooth versions. Moreover, the results indicated that the surface geometry with the best attachment performance was the geometry most similar to octopus sucker morphology. The results obtained in this work can be utilized to design artificial suction cups with higher wet attachment performance. PMID:25970079

  11. TAIR- TRANSONIC AIRFOIL ANALYSIS COMPUTER CODE

    NASA Technical Reports Server (NTRS)

    Dougherty, F. C.

    1994-01-01

    . Output from TAIR may include aerodynamic coefficients, the airfoil surface solution, convergence histories, and printer plots of Mach number and density contour maps. The TAIR program is written in FORTRAN IV for batch execution and has been implemented on a CDC 7600 computer with a central memory requirement of approximately 155K (octal) of 60 bit words. The TAIR program was developed in 1981.

  12. 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

  13. 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.

  14. 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.

  15. 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

  16. Self-pumping suction/propulsion for laminar flow bodies

    NASA Astrophysics Data System (ADS)

    Rogers, K. H.; King, D. A.

    1984-06-01

    An analysis is presented to investigate the feasibility of a self-pumping suction system for a very low drag suction laminar flow control (SLFC) underwater test body. The nose and afterbody of a torpedo-like body are contoured such that a prominent low-pressure region in the aft part of the body can serve as a suction pump to suck the boundary layer fluid through the circumferential surface-slots and thus laminarize the entire body length forward of the aft low-pressure peak. The results indicate that it is feasible to laminarize a test body in this fashion at a design speed, such as 40 knots; but that the laminarization of a particular configuration is limited to a band of speeds at and near the design speed. If an SLFC test body with a wide range of speed capability is desired, then a controllable-speed suction pump and controllable suction distribution along the body are indicated. The analysis includes a suction system design calculation example and should be a useful reference for future development of undersea SLFC vehicles.

  17. The Influence of Sweep on the Aerodynamic Loading of an Oscillating NACA0012 Airfoil. Volume 2: Data Report

    NASA Technical Reports Server (NTRS)

    St.hilaire, A. O.; Carta, F. O.

    1979-01-01

    The effect of sweep on the dynamic response of the NACA 0012 airfoil was investigated. Unsteady chordwise distributed pressure data were obtained from a tunnel spanning wing equipped with 21 single surface transducers (13 on the suction side and 8 on the pressure side of the airfoil). The pressure data were obtained at pitching amplitudes of 8 and 10 degrees over a tunnel Mach number range of 0.10 to 0.46 and a pitching frequency range of 2.5 to 10.6 cycles per second. The wing was oscillated in the unswept and swept positions about the quarter-chord pivot axis relative to mean incidence angle settings of 0, 9, 12, and 15 degrees. A compilation of all the response data obtained during the test program is presented. These data are in the form of normal force, chord force, lift force, pressure drag, and moment hysteresis loops derived from chordwise integrations of the unsteady pressure distributions. The hysteresis loops are organized in two main sections. In the first section, the loop data are arranged to show the effect of sweep (lambda = 0 and 30 deg) for all available combinations of mean incidence angle, pitching amplitude, reduced frequency, and chordwise Mach number. The second section shows the effect of chordwise Mach number (MC = 0.30 and MC = 0.40) on the swept wing response for all available combinations of mean incidence angle, pitching amplitude, and reduced frequency.

  18. 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.

  19. Suction patterns in preterm infants.

    PubMed

    Martell, M; Martínez, G; González, M; Díaz Rosselló, J L

    1993-01-01

    The suction pattern for breast and bottle feeding in two groups of preterm infants is described. The time elapsed between birth and the moment of suction was longer in preterm neonates born at lower gestational ages for both groups studied, breast and bottle fed (figure 1). The evolution of suckling in breastfeeding was analyzed in a composite study (longitudinal and transverse) in a group of 16 neonates starting from 32 weeks of gestation. The velocity of milk extraction during suckling varied with gestational age. It was uniform at lower gestational ages, then it became faster in the first minutes and at the 36th week, it was very similar to that of mature neonates (figure 2 and table I). The evaluation of bottle feeding was performed in a transverse study in 46 preterm neonates which had been exclusively bottle fed during 1 or 2 weeks. All of them had previously been fed using an orogastric tube. Nourishing time was shorter than in breastfeeding; the average duration was 3.7 minutes (table II). The greatest volume was ingested in the first minute, 40% (range between 44 and 25%) (figure 3). The frequency of suction did not change the duration of feeding, but it was found that the efficiency of suction (number of suctions to ingest 1 cc) was significantly lower in the first minute (Anova, p < 0.05) (figure 4).

  20. Measurements in a Transitional Boundary Layer Under Low-Pressure Turbine Airfoil Conditions

    NASA Technical Reports Server (NTRS)

    Simon, Terrence W.; Qiu, Songgang; Yuan, Kebiao; Ashpis, David (Technical Monitor); Simon, Fred (Technical Monitor)

    2000-01-01

    This report presents the results of an experimental study of transition from laminar to turbulent flow in boundary layers or in shear layers over separation zones on a convex-curved surface which simulates the suction surface of a low-pressure turbine airfoil. Flows with various free-stream turbulence intensity (FSTI) values (0.5%, 2.5% and 10%), and various Reynolds numbers (50,000, 100,000 200,000 and 300,000) are investigated. Reynold numbers in the present study are based on suction surface length and passage exit mean velocity. Flow separation followed by transition within the separated flow region is observed for the lower-Re cases at each of the FSTI levels. At the highest Reynolds numbers and at elevated FSn, transition of the attached boundary layer begins before separation, and the separation zone is small. Transition proceeds in the shear layer over the separation bubble. For both the transitional boundary layer and the transitional shear layer, mean velocity, turbulence intensity and intermittency (the fraction of the time the flow is turbulent) distributions are presented. The present data are compared to published distribution models for bypass transition, intermittency distribution through transition, transition start position, and transition length. A model developed for transition of separated flows is shown to adequately predict the location of the beginning of transition, for these cases, and a model developed for transitional boundary layer flows seems to adequately predict the path of intermittency through transition when the transition start and end are known. These results are useful for the design of low-pressure turbine stages which are known to operate under conditions replicated by these tests.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Method for forming a liquid cooled airfoil for a gas turbine

    DOEpatents

    Grondahl, Clayton M.; Willmott, Leo C.; Muth, Myron C.

    1981-01-01

    A method for forming a liquid cooled airfoil for a gas turbine is disclosed. A plurality of holes are formed at spaced locations in an oversized airfoil blank. A pre-formed composite liquid coolant tube is bonded into each of the holes. The composite tube includes an inner member formed of an anti-corrosive material and an outer member formed of a material exhibiting a high degree of thermal conductivity. After the coolant tubes have been bonded to the airfoil blank, the airfoil blank is machined to a desired shape, such that a portion of the outer member of each of the composite tubes is contiguous with the outer surface of the machined airfoil blank. Finally, an external skin is bonded to the exposed outer surface of both the machined airfoil blank and the composite tubes.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. A critical evaluation of the predictions of the NASA-Lockheed multielement airfoil computer program

    NASA Technical Reports Server (NTRS)

    Brune, G. W.; Manke, J. W.

    1978-01-01

    Theoretical predictions of several versions of the multielement airfoil computer program are evaluated. The computed results are compared with experimental high lift data of general aviation airfoils with a single trailing edge flap, and of airfoils with a leading edge flap and double slotted trailing edge flaps. Theoretical and experimental data include lift, pitching moment, profile drag and surface pressure distributions, boundary layer integral parameters, skin friction coefficients, and velocity profiles.

  13. 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.

  14. 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.

  15. 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.

  16. An experimental and numerical investigation on the formation of stall-cells on airfoils

    NASA Astrophysics Data System (ADS)

    Manolesos, M.; Papadakis, G.; Voutsinas, S.

    2014-12-01

    Stall Cells (SCs) are large scale three-dimensional structures of separated flow that have been observed on the suction side of airfoils designed for or used on wind turbine blades. SCs are unstable in nature but can be stabilised by means of a localized disturbance; here in the form of a zigzag tape covering 10% of the wing span. Based on extensive tuft flow visualisations, the resulting flow was found macroscopically similar to the undisturbed flow. Next a combined investigation was carried out including pressure recordings, Stereo-PIV measurements and CFD simulations. The investigation parameters were the aspect ratio, the angle of attack and the Re number. Tuft and pressure data were found in good agreement. The 3D CFD simulations reproduced the structure of the SCs in qualitative agreement with the experimental data but had a delay of ~3deg in capturing the first appearance of a SC. The error in Cl max prediction was 7% compared to 19% for the 2D cases. Tests show that SCs grow with Re number and angle of attack. Also analysis of the time averaged computational results indicated the presence of three types of vortices: (a) the trailing edge line vortex (TELV) in the wake, (b) the separation line vortex (SLV) over the wing and (c) the SC vortices. The TELV and SLV run parallel to the trailing edge and are of opposite sign, while the SC vortices start normal to the wing suction surface, then bend towards the SC centre and later extend downstream, with their vorticity parallel to the free stream.

  17. Suction is kid's play: extremely fast suction in newborn seahorses.

    PubMed

    Van Wassenbergh, Sam; Roos, Gert; Genbrugge, Annelies; Leysen, Heleen; Aerts, Peter; Adriaens, Dominique; Herrel, Anthony

    2009-04-23

    Ongoing anatomical development typically results in a gradual maturation of the feeding movements from larval to adult fishes. Adult seahorses are known to capture prey by rotating their long-snouted head extremely quickly towards prey, followed by powerful suction. This type of suction is powered by elastic recoil and requires very precise coordination of the movements of the associated feeding structures, making it an all-or-none phenomenon. Here, we show that newborn Hippocampus reidi are able to successfully feed using an extremely rapid and powerful snout rotation combined with a high-volume suction, surpassing that observed in adult seahorses. An inverse dynamic analysis shows that an elastic recoil mechanism is also used to power head rotation in newborn H. reidi. This illustrates how extreme levels of performance in highly complex musculoskeletal systems can be present at birth given a delayed birth and rapid development of functionally important structures. The fact that the head skeleton of newborn seahorses is still largely cartilaginous may not be problematic because the hydrodynamic stress on the rotating snout appeared considerably lower than in adult syngnathids.

  18. 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.

  19. 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.

  20. A rectangular capillary suction apparatus

    SciTech Connect

    Lee, D.J. . Dept. of Chemical Engineering); Hsu, Y.H. . Dept. of Chemical Engineering)

    1994-06-01

    Fluid flow and cake formation in a rectangular capillary suction apparatus (RCSA) are investigated experimentally and theoretically. Water, methanol, ethanol, and ethylene glycol are used to study the effects of liquid properties, and CaCO[sub 3], kaolin, and bentonite slurries are employed for studying the effects of cake formation on capillary suction-time (CST). A theory based on a diffusion-like approach is developed. The liquid saturation under the inner cell will approach a constant value when the wet front distance is large. A method based on this experimental finding for estimating the cake specific resistance is proposed. The agreement between experiments and calculations is close. The RCSA is superior to the cylindrical CSA when treating liquids with small diffusivities or slurries with high solid concentration and/or with high averaged specific resistance.

  1. 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.

  2. 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.

  3. 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.

  4. 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

  5. Dual-durometer soft suction foot robot for concrete inspection

    NASA Astrophysics Data System (ADS)

    Huston, Dryver; Burns, Dylan; Gardner-Morse, John; Montane, Paul; Angola, Enrique

    2014-03-01

    Climbing on concrete, masonry and brick with automated machines is difficult due to the uneven surfaces that prevent getting a good grip. This paper describes developments in using dual-durometer pneumatic suction feet for gripping onto concrete surfaces as part of a multi-legged robotic climbing system for inspecting concrete structures with vertical walls. The dual durometer technique presents a compliant suction tip to the concrete thereby producing a good seal against an irregular surface, and stiff component to deliver the structural rigidity needed for walking and climbing. Individually actuated pneumatic Venturi vacuum generators provide the suction from positive pneumatic pressure in a manner that is robust against leaks that cause the systemic vacuum collapse that can plague other vacuum configurations. The feet are attached to a six-legged robot that with a nominal floor walking capability and gait. Climbing a wall requires modification to leg actuation and gait, along with suction feet. System design, integration, concrete wall climbing performance and sensor deployment in the form of a lightweight ground penetrating radar system are presented.

  6. Numerical exploration of the origin of aerodynamic enhancements in [low-Reynolds number] corrugated airfoils

    NASA Astrophysics Data System (ADS)

    Barnes, Caleb J.; Visbal, Miguel R.

    2013-11-01

    This paper explores the flow structure of a corrugated airfoil using a high-fidelity implicit large eddy simulation approach. The first three-dimensional simulations for a corrugated wing section are presented considering a range of Reynolds numbers of Rec = 5 × 103 to 5.8 × 104 bridging the gap left by previous numerical and experimental studies. Several important effects are shown to result from the corrugations in the leading-edge region. First, interaction between the detached shear layer and the first corrugation peak promotes recirculation upstream and enhances transition to turbulence due to flow instabilities. Thus, early transitional flow develops on the corrugated wing which helps to delay stall even at Reynolds numbers as low as Rec = 1 × 104. Transition is shown to occur as early as Rec = 7.5 × 103 and quickly advances toward the leading-edge as Reynolds number is increased. Modification of the first corrugation peak height produces significantly reduced separation and improved aerodynamic forces demonstrating the sensitivity of flow behavior to leading-edge geometry. Second, the unusual orientation of the corrugated surface and strong suction resulting from rapidly turning fluid over the separated region upstream of the first corrugation produces a new effect which serves to reduce drag. This effect was amplified through the enhanced interaction produced by a modified geometry. Corrugations were found to be most advantageous in the leading-edge region and could be optimized to properly take advantage of the flow field under different operating conditions.

  7. Effect of Axial Velocity Density Ratio on the Performance of a Controlled Diffusion Airfoil Compressor Cascade

    NASA Astrophysics Data System (ADS)

    Senthil Kumaran, R.; Kamble, Sachin; Swamy, K. M. M.; Nagpurwala, Q. H.; Bhat, Ananthesha

    2015-12-01

    Axial Velocity Density Ratio (AVDR) is an important parameter to check the two-dimensionality of cascade flows. It can have significant influence on the cascade performance and the secondary flow structure. In the present study, the effect of AVDR has been investigated on a highly loaded Controlled Diffusion airfoil compressor cascade. Detailed 3D Computational Fluid Dynamics (CFD) studies were carried out with the cascade at five different AVDRs. Key aerodynamic performance parameters and flow structure through the cascade were analyzed in detail. CFD results of one AVDR were validated with the experimental cascade test data and were seen to be in good agreement. Loss characteristics of the cascade varied significantly with change in AVDR. Increase in AVDR postponed the point of separation on the suction surface, produced thinner boundary layers and caused substantial drop in the pressure loss coefficient. Strong end wall vortices were noticed at AVDR of 1.177. At higher AVDRs, the flow was well guided even close to the end wall and the secondary flows diminished. The loading initially improved with increase in AVDR. Beyond a certain limit, further increase in AVDR offered no improvements to the loading but rather resulted in drop in diffusion and deviation.

  8. Active Control of Separation on a Low Reynolds Number Airfoil Using Synthetic Jet Actuation

    NASA Astrophysics Data System (ADS)

    Feero, Mark

    Wind tunnel experiments were used to study the effect of excitation amplitude and frequency on flow separation using synthetic jet actuation. A synthetic jet actuator was located near the leading edge of a NACA0025 airfoil at a chord-based Reynolds number of 100,000 and angle-of-attack of 10°. Under these flow conditions, the boundary layer separated from the suction surface and failed to reattach. Low-frequency excitation was used to target flow instabilities, while high-frequency excitation was performed at time scales an order of magnitude smaller. Low-frequency excitation at the separated shear layer frequency was found to be the most effective technique for flow reattachment and drag reduction. The results suggested that flow reattachment depended on exceeding a threshold momentum coefficient that varied with excitation frequency. Furthermore, a local minimum in drag independent of excitation frequency was achieved when the momentum coefficient corresponded to an average jet velocity that matched the freestream velocity.

  9. The leading-edge stall of airfoils with various nose shapes

    NASA Astrophysics Data System (ADS)

    Kraljic, Matthew; Rusak, Zvi; Wang, Shixiao

    2015-11-01

    We study the inception of leading-edge stall on stationary, smooth thin airfoils with various nose shapes of the form xa (where 0 < a < 1 / 2) at low to moderately high chord Reynolds number flows. A reduced-order, multi-scale model problem is developed and solved using numerical simulations. The asymptotic theory demonstrates that a subsonic flow about a thin airfoil can be described in terms of an outer region, around most of the airfoil's chord, and an inner region, around the nose, that asymptotically match each other. The flow in the outer region is dominated by the classical thin airfoil theory. Scaled (magnified) coordinates and a modified (smaller) Reynolds number ReM are used to correctly account for the nonlinear behavior and extreme velocity changes in the inner region, where both the near-stagnation and high suction areas occur. The inner region problem is solved numerically to determine the inception of leading-edge stall on the nose. It is found that stall is delayed to higher angles of attack with the decrease of nose parameter a. Specifically, new airfoil shapes are proposed with increased stall angle at subsonic speeds and higher critical Mach numbers at transonic speeds.

  10. 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.

  11. 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.

  12. Optimum Suction Distribution for Transition Control

    NASA Technical Reports Server (NTRS)

    Balakumar, P.; Hall, P.

    1996-01-01

    The optimum suction distribution which gives the longest laminar region for a given total suction is computed. The goal here is to provide the designer with a method to find the best suction distribution subject to some overall constraint applied to the suction. We formulate the problem using the Lagrangian multiplier method with constraints. The resulting non-linear system of equations is solved using the Newton-Raphson technique. The computations are performed for a Blasius boundary layer on a flat-plate and crossflow cases. For the Blasius boundary layer, the optimum suction distribution peaks upstream of the maximum growth rate region and remains flat in the middle before it decreases to zero at the end of the transition point. For the stationary and travelling crossflow instability, the optimum suction peaks upstream of the maximum growth rate region and decreases gradually to zero.

  13. 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

  14. 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.

  15. 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)

  16. 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.

  17. 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.

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

    NASA Technical Reports Server (NTRS)

    Nagamatsu, H. T.; Ficarra, R.; Orozco, R.

    1983-01-01

    The optimization of passive shock wave/boundary layer control for supercritical airfoil drag reduction was investigated in a 3 in. x 15.4 in. Transonic Blowdown Wind Tunnel. A 14% thick supercritical airfoil was tested with 0%, 1.42% and 2.8% porosities at Mach numbers of .70 to .83. The 1.42% case incorporated a linear increase in porosity with the flow direction while the 2.8% case was uniform porosity. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid-surface airfoils. While the results show that linear 1.42% porosity actually led to a slight increase in drag it was found that the uniform 2.8% porosity can lead to a drag reduction of 46% at M = .81.

  19. 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.

  20. Making Large Suction Panels For Laminar-Flow Control

    NASA Technical Reports Server (NTRS)

    Maddalon, Dal V.

    1991-01-01

    Perforated titanium panels used to identify and resolve issues related to manufacture. Recently, relatively large suction panels with aerodynamically satisfactory surface perforations and with surface contours and smoothness characteristics necessary for Laminar-Flow Control (LFC) designed, fabricated, and tested. Requirements of production lines for commercial transport airplanes carefully considered in development of panels. Sizes of panels representative of what is used on wing of commercial transport airplane. Tests of perforated panels in transonic wind tunnel demonstrated aerodynamic stability at flight mach numbers.

  1. The acoustics and unsteady wall pressure of a circulation control airfoil

    NASA Astrophysics Data System (ADS)

    Silver, Jonathan C.

    A Circulation Control (CC) airfoil uses a wall jet exiting onto a rounded trailing edge to generate lift via the Coanda effect. The aerodynamics of the CC airfoil have been studied extensively. The acoustics of the airfoil are, however, much less understood. The primary goal of the present work was to study the radiated sound and unsteady surface pressures of a CC airfoil. The focus of this work can be divided up into three main categories: characterizing the unsteady surface pressures, characterizing the radiated sound, and understanding the acoustics from surface pressures. The present work is the first to present the unsteady surface pressures from the trailing edge cylinder of a circulation control airfoil. The auto-spectral density of the unsteady surface pressures at various locations around the trailing edge are presented over a wide range of the jets momentum coefficient. Coherence of pressure and length scales were computed and presented. Single microphone measurements were made at a range of angles for a fixed observer distance in the far field. Spectra are presented for select angles to show the directivity of the airfoil's radiated sound. Predictions of the acoustics were made from unsteady surface pressures via Howe's curvature noise model and a modified Curle's analogy. A summary of the current understanding of the acoustics from a CC airfoil is given along with suggestions for future work.

  2. Rime ice accretion and its effect on airfoil performance. Ph.D. Thesis. Final Report

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.

    1982-01-01

    A methodology was developed to predict the growth of rime ice, and the resulting aerodynamic penalty on unprotected, subcritical, airfoil surfaces. The system of equations governing the trajectory of a water droplet in the airfoil flowfield is developed and a numerical solution is obtained to predict the mass flux of super cooled water droplets freezing on impact. A rime ice shape is predicted. The effect of time on the ice growth is modeled by a time-stepping procedure where the flowfield and droplet mass flux are updated periodically through the ice accretion process. Two similarity parameters, the trajectory similarity parameter and accumulation parameter, are found to govern the accretion of rime ice. In addition, an analytical solution is presented for Langmuir's classical modified inertia parameter. The aerodynamic evaluation of the effect of the ice accretion on airfoil performance is determined using an existing airfoil analysis code with empirical corrections. The change in maximum lift coefficient is found from an analysis of the new iced airfoil shape. The drag correction needed due to the severe surface roughness is formulated from existing iced airfoil and rough airfoil data. A small scale wind tunnel test was conducted to determine the change in airfoil performance due to a simulated rime ice shape.

  3. 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.

  4. Turbine Vane External Heat Transfer. Volume 1: Analytical and Experimental Evaluation of Surface Heat Transfer Distributions with Leading Edge Showerhead Film Cooling

    NASA Technical Reports Server (NTRS)

    Turner, E. R.; Wilson, M. D.; Hylton, L. D.; Kaufman, R. M.

    1985-01-01

    Progress in predictive design capabilities for external heat transfer to turbine vanes was summarized. A two dimensional linear cascade (previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils) was used to examine the effect of leading edge shower head film cooling on downstream heat transfer. The data were used to develop and evaluate analytical models. Modifications to the two dimensional boundary layer model are described. The results were used to formulate and test an effective viscosity model capable of predicting heat transfer phenomena downstream of the leading edge film cooling array on both the suction and pressure surfaces, with and without mass injection.

  5. Practical aspects of oronasopharyngeal suction in children.

    PubMed

    Knox, Tony

    2011-09-01

    Undergoing oronasopharyngeal suction is an unpleasant experience, but the intervention may prevent the deterioration of children who cannot clear their secretions. A successful procedure requires a practitioner with appropriate knowledge, dexterity and communication skills. Thorough training should be provided and a careful risk-benefit assessment is important before suction is performed.

  6. Is nasogastric suction necessary in acute pancreatitis?

    PubMed Central

    Naeije, R; Salingret, E; Clumeck, N; De Troyer, A; Devis, G

    1978-01-01

    Fifty-eight patients with mild to moderately severe acute pancreatitis were randomly allocated to treatment with or without nasogastric suction (27 and 31 patients respectively). Intravenous fluids and pethidine hydrochloride were also given. The two groups were comparable clinically at the start of the study. There were no differences between the two groups in the mean duration of the following features: abdominal pain or tenderness; absence of bowel movements; raised serum amylase concentration; time to resumption of oral feeding; and days in hospital. Prolonged hyperamylasaemia (serum amylase greater than 0.33 mU/l) occurred in one patient in the suction group and in three patients in the non-suction group. A mild recurrence of abdominal pain after resumption of oral feeding occurred in three patients in the suction group and in two patients in the non-suction group. Two patients in the suction group developed overt consumption coagulopathy and two others pulmonary complications. No patient in the non-suction group had complications. The findings suggest that most patients with mild to moderately severe acute pancreatitis do not benefit from nasogastric suction. The procedure should be elective rather than mandatory in treating this condition. PMID:698650

  7. Materials for Advanced Turbine Engines (MATE). Project 4: Erosion resistant compressor airfoil coating

    NASA Technical Reports Server (NTRS)

    Rashid, J. M.; Freling, M.; Friedrich, L. A.

    1987-01-01

    The ability of coatings to provide at least a 2X improvement in particulate erosion resistance for steel, nickel and titanium compressor airfoils was identified and demonstrated. Coating materials evaluated included plasma sprayed cobalt tungsten carbide, nickel carbide and diffusion applied chromium plus boron. Several processing parameters for plasma spray processing and diffusion coating were evaluated to identify coating systems having the most potential for providing airfoil erosion resistance. Based on laboratory results and analytical evaluations, selected coating systems were applied to gas turbine blades and evaluated for surface finish, burner rig erosion resistance and effect on high cycle fatigue strength. Based on these tests, the following coatings were recommended for engine testing: Gator-Gard plasma spray 88WC-12Co on titanium alloy airfoils, plasma spray 83WC-17Co on steel and nickel alloy airfoils, and Cr+B on nickel alloy airfoils.

  8. 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.

  9. Materials for advanced turbine engines (MATE). Project 4: erosion resistant compressor airfoil coating

    SciTech Connect

    Rashid, J.M.; Freling, M.; Friedrich, L.A.

    1987-05-01

    The ability of coatings to provide at least a 2X improvement in particulate erosion resistance for steel, nickel and titanium compressor airfoils was identified and demonstrated. Coating materials evaluated included plasma sprayed cobalt tungsten carbide, nickel carbide and diffusion applied chromium plus boron. Several processing parameters for plasma spray processing and diffusion coating were evaluated to identify coating systems having the most potential for providing airfoil erosion resistance. Based on laboratory results and analytical evaluations, selected coating systems were applied to gas turbine blades and evaluated for surface finish, burner rig erosion resistance and effect on high cycle fatigue strength. Based on these tests, the following coatings were recommended for engine testing: Gator-Gard plasma spray 88WC-12Co on titanium alloy airfoils, plasma spray 83WC-17Co on steel and nickel alloy airfoils, and Cr+B on nickel alloy airfoils.

  10. 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.

  11. An analytic study of nonsteady two-phase laminar boundary layer around an airfoil

    NASA Technical Reports Server (NTRS)

    Hsu, Yu-Kao

    1989-01-01

    Recently, NASA, FAA, and other organizations have focused their attention upon the possible effects of rain on airfoil performance. Rhode carried out early experiments and concluded that the rain impacting the aircraft increased the drag. Bergrum made numerical calculation for the rain effects on airfoils. Luers and Haines did an analytic investigation and found that heavy rain induces severe aerodynamic penalties including both a momentum penalty due to the impact of the rain and a drag and lift penalty due to rain roughening of the airfoil and fuselage. More recently, Hansman and Barsotti performed experiments and declared that performance degradation of an airfoil in heavy rain is due to the effective roughening of the surface by the water layer. Hansman and Craig did further experimental research at low Reynolds number. E. Dunham made a critical review for the potential influence of rain on airfoil performance. Dunham et al. carried out experiments for the transport type airfoil and concluded that there is a reduction of maximum lift capability with increase in drag. There is a scarcity of published literature in analytic research of two-phase boundary layer around an airfoil. Analytic research is being improved. The following assumptions are made: the fluid flow is non-steady, viscous, and incompressible; the airfoil is represented by a two-dimensional flat plate; and there is only a laminar boundary layer throughout the flow region. The boundary layer approximation is solved and discussed.

  12. 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.

  13. 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.

  14. On the attenuating effect of permeability on the low frequency sound of an airfoil

    NASA Astrophysics Data System (ADS)

    Weidenfeld, M.; Manela, A.

    2016-08-01

    The effect of structure permeability on the far-field radiation of a thin airfoil is studied. Assuming low-Mach and high-Reynolds number flow, the near- and far-field descriptions are investigated at flapping-flight and unsteady flow conditions. Analysis is carried out using thin-airfoil theory and compact-body-based calculations for the hydrodynamic and acoustic fields, respectively. Airfoil porosity is modeled via Darcy's law, governed by prescribed distribution of surface intrinsic permeability. Discrete vortex model is applied to describe airfoil wake evolution. To assess the impact of penetrability, results are compared to counterpart predictions for the sound of an impermeable airfoil. Considering the finite-chord airfoil as "acoustically transparent", the leading-order contribution of surface porosity is obtained in terms of an acoustic dipole. It is shown that, at all flow conditions considered, porosity causes attenuation in outcome sound level. This is accompanied by a time-delay in the pressure signal, reflecting the mediating effect of permeability on the interaction of fluid flow with airfoil edge points. To the extent that thin-airfoil theory holds (requiring small normal-to-airfoil flow velocities), the results indicate on a decrease of ~ 10 percent and more in the total energy radiated by a permeable versus an impermeable airfoil. This amounts to a reduction in system sound pressure level of 3 dB and above at pitching flight conditions, where the sound-reducing effect of the seepage dipole pressure becomes dominant. The applicability of Darcy's law to model the effect of material porosity is discussed in light of existing literature.

  15. 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.

  16. 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.

  17. 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.

  18. Experimental results for the Eppler 387 airfoil at low Reynolds numbers in the Langley low-turbulence pressure tunnel

    NASA Technical Reports Server (NTRS)

    Mcghee, Robert J.; Walker, Betty S.; Millard, Betty F.

    1988-01-01

    Experimental results were obtained for an Eppler 387 airfoil in the Langley Low Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.03 to 0.13 and a chord Reynolds number range for 60,000 to 460,000. Lift and pitching moment data were obtained from airfoil surface pressure measurements and drag data for wake surveys. Oil flow visualization was used to determine laminar separation and turbulent reattachment locations. Comparisons of these results with data on the Eppler 387 airfoil from two other facilities as well as the Eppler airfoil code are included.

  19. 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

  20. 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.

  1. Synthetic Vortex Generator Jets Used to Control Separation on Low-Pressure Turbine Airfoils

    NASA Technical Reports Server (NTRS)

    Ashpis, David E.; Volino, Ralph J.

    2005-01-01

    Low-pressure turbine (LPT) airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. When the adverse pressure gradient on the suction side of these airfoils becomes strong enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss of lift and a subsequent degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions because of the thinner air and lower Reynolds numbers at altitude. Component efficiency can drop significantly between takeoff and cruise conditions. The decrease is about 2 percent in large commercial transport engines, and it could be as large as 7 percent in smaller engines operating at higher altitudes. Therefore, it is very beneficial to eliminate, or at least reduce, the separation bubble.

  2. 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.

  3. 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.

  4. 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.

  5. A computed tomographic scan assessment of endotracheal suctioning-induced bronchoconstriction in ventilated sheep.

    PubMed

    Lu, Q; Capderou, A; Cluzel, P; Mourgeon, E; Abdennour, L; Law-Koune, J D; Straus, C; Grenier, P; Zelter, M; Rouby, J J

    2000-11-01

    This study was directed at assessing changes in bronchial cross-sectional surface areas (BCSA) and in respiratory resistance induced by endotracheal suctioning in nine anesthetized sheep. Cardiorespiratory parameters (Swan-Ganz catheter), respiratory resistance (inspiratory occlusion technique), BCSA, and lung aeration (computed tomography) were studied at baseline, during endotracheal suctioning, and after 20 consecutive hyperinflations. Measurements performed initially at an inspired oxygen fraction (FI(O(2))) of 0.3 were repeated at an FI(O(2)) of 1.0. At an FI(O(2)) of 0.3, endotracheal suctioning resulted in atelectasis, a reduction in BCSA of 29 +/- 23% (mean +/- SD), a decrease in arterial oxygen saturation from 95 +/- 3% to 87 +/- 12% (p = 0.02), an increase in venous admixture from 19 +/- 10% to 31 +/- 19% (p = 0. 006), and an increase in lung tissue resistance (DR(rs)) (p = 0. 0003). At an FI(O(2)) of 1.0, despite an extension of atelectasis and an increase in pulmonary shunt from 19 +/- 5% to 36 +/- 2% (p < 0.0001), arterial O(2) desaturation was prevented and BCSA decreased by only 7 +/- 32%. A recruitment maneuver after endotracheal suctioning entirely reversed the suctioning-induced increase in DR(rs) and atelectasis. In three lidocaine-pretreated sheep, the endotracheal suctioning-induced reduction of BCSA was entirely prevented. These data suggest that the endotracheal suctioning-induced decrease in BCSA is related to atelectasis and bronchoconstriction. Both effects can be reversed by hyperoxygenation maneuver before suctioning in combination with recruitment maneuver after suctioning.

  6. 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.

  7. Effects of a ground vortex on the aerodynamics of an airfoil

    NASA Technical Reports Server (NTRS)

    Krothapalli, A.; Leopold, D.

    1988-01-01

    An experimental investigation was carried out to study the aerodynamics of an airfoil with a rectangular jet exiting from its lower surface at fifty percent of the chord. The airfoil was tested with and without the influence of a ground plane. Surface static pressures were measured on the airfoil at jet to free stream velocity ratios ranging from 0 to 9. From these pressures, the variation of C sub L with velocity ratio was easily determined. The measurements indicated significant positive and negative pressure regions on the lower surface of the airfoil ahead of and after the nozzle exit respectively. The presence of a ground plane enhanced these pressure regions at low velocity ratios, but at a particular ratio for each plane location, a recirculation zone or vortex formed ahead of the jet resulting in decreased pressures and a drop in C sub L.

  8. Suction blister skin grafting--a modern application.

    PubMed

    Parbhoo, A V; Simpson, M T

    2014-03-01

    The suction blistering technique produces an ultra-thin skin graft with no morbidity at the donor site. Negative pressure using wall suction in outpatients is used to generate a graft that can be used for reconstruction, and it avoids the need for invasive procedures in patients with coexisting conditions. The harvested tissue has a low metabolic demand and survival is excellent. We used it in a patient when previous reconstructions after excision of skin cancer had failed. Graft survival was more than 95% by surface area and there was no donor site morbidity. We have found it particularly useful for grafting over Integra® dermal regeneration template (Integra LifeSciences Corporation, NJ, USA) to produce healing at difficult sites. Patients tolerate the procedure well and the donor site heals quickly. It is useful where recipient vascularity is poor or where coexisting conditions prevent complex procedures.

  9. High Reynolds Number Hybrid Laminar Flow Control (HLFC) Flight Experiment. Report 4; Suction System Design and Manufacture

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This document describes the design of the leading edge suction system for flight demonstration of hybrid laminar flow control on the Boeing 757 airplane. The exterior pressures on the wing surface and the required suction quantity and distribution were determined in previous work. A system consisting of porous skin, sub-surface spanwise passages ("flutes"), pressure regulating screens and valves, collection fittings, ducts and a turbocompressor was defined to provide the required suction flow. Provisions were also made for flexible control of suction distribution and quantity for HLFC research purposes. Analysis methods for determining pressure drops and flow for transpiration heating for thermal anti-icing are defined. The control scheme used to observe and modulate suction distribution in flight is described.

  10. [Suction-irrigator device for microsurgery: technical note].

    PubMed

    Gusmão, Sebastião; Silveira, Roberto Leal

    2003-06-01

    A modification of the conventional suction device for microsurgery is described. It consists of a built-in tube in another tube, being the first connected to the suction device and the second to the irrigation. This suction-irrigator device allows to accomplish the suction and irrigation simultaneously and in a precise way. PMID:12806518

  11. Calculation of vortex lift effect for cambered wings by the suction analogy

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

    An improved version of Woodward's chord plane aerodynamic panel method for subsonic and supersonic flow is developed for cambered wings exhibiting edge separated vortex flow, including those with leading edge vortex flaps. The exact relation between leading edge thrust and suction force in potential flow is derived. Instead of assuming the rotated suction force to be normal to wing surface at the leading edge, new orientation for the rotated suction force is determined through consideration of the momentum principle. The supersonic suction analogy method is improved by using an effective angle of attack defined through a semi-empirical method. Comparisons of predicted results with available data in subsonic and supersonic flow are presented.

  12. 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.

  13. Potential flow around two-dimensional airfoils using a singular integral method

    NASA Technical Reports Server (NTRS)

    Nguyen, Yves; Wilson, Dennis

    1987-01-01

    The problem of potential flow around two-dimensional airfoils is solved by using a new singular integral method. The potential flow equations for incompressible potential flow are written in a singular integral equation. The equation is solved at N collocation points on the airfoil surface. A unique feature of this method is that the airfoil geometry is specified as an independent variable in the exact integral equation. Compared to other numerical methods, the present calculation procedure is much simpler and gives remarkable accuracy for many body shapes. An advantage of the present method is that it allows the inverse design calculation and the results are extremely accurate.

  14. Fluid mechanics mechanisms in the stall process of airfoils for helicopters

    NASA Technical Reports Server (NTRS)

    Young, W. H., Jr.

    1981-01-01

    Phenomena that control the flow during the stall portion of a dynamic stall cycle are analyzed, and their effect on blade motion is outlined. Four mechanisms by which dynamic stall may be initiated are identified: (1) bursting of the separation bubble, (2) flow reversal in the turbulent boundary layer on the airfoil upper surface, (3) shock wave-boundary layer interaction behind the airfoil crest, and (4) acoustic wave propagation below the airfoil. The fluid mechanics that contribute to the identified flow phenomena are summarized, and the usefulness of a model that incorporates the required fluid mechanics mechanisms is discussed.

  15. 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.

  16. 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.

  17. [Ultrasonic aspirator with controllable suction system--variable action suction adapter and clinical experience with it].

    PubMed

    Nagasawa, S; Shimano, H; Kuroiwa, T

    2000-12-01

    While the ultrasonic aspirator (UA) has been widely used as one of the indispensable tools in the field of neurosurgery, a potential risk when using the present UA is injury to the neurovascular structures due to ultrasonic pulverization and constant forceful suction power. We have devised a small variable action suction adapter that can be used in a similar manner to conventional surgical suction tubes. The UA control unit and the handpiece used in this study were the Sonopet UST-2000 and HA-01, respectively (M & M Corporation Tokyo, Japan). The handpiece is slim, with the mid-portion diameter of 13 mm, and it weighs 100 grams. A variable action suction adapter was made from polycarbonate of 15 x 12 x 13 mm in size. The adapter was connected to the suction tube using a Y-shaped connector (Fig. 2 A), which was integrated into the handpiece. The suction power is regulated by variably closing the oval-shaped hole. The adapter can be variously placed on and rotated around the handpiece (Fig. 2 B and C) so that either the right or left hand handles it in a similar fashion to conventional suction tubes. We used this UA in surgery for 8 patients with large brain tumors (meningioma in 5 cases, metastatic brain tumor in 2 cases and glioma in one case). It reduced the risk of suction-related injury to the neurovascular structures and was handled in a similar manner to conventional suction tubes. This adapter ensures the complete control of suction power, which will reduce the risk of suction injury.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Hatton, Kenneth S.

    1999-01-01

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

  20. On the influence of airfoil deviations on the aerodynamic performance of wind turbine rotors

    NASA Astrophysics Data System (ADS)

    Winstroth, J.; Seume, J. R.

    2016-09-01

    The manufacture of large wind turbine rotor blades is a difficult task that still involves a certain degree of manual labor. Due to the complexity, airfoil deviations between the design airfoils and the manufactured blade are certain to arise. Presently, the understanding of the impact of manufacturing uncertainties on the aerodynamic performance is still incomplete. The present work analyzes the influence of a series of airfoil deviations likely to occur during manufacturing by means of Computational Fluid Dynamics and the aeroelastic code FAST. The average power production of the NREL 5MW wind turbine is used to evaluate the different airfoil deviations. Analyzed deviations include: Mold tilt towards the leading and trailing edge, thick bond lines, thick bond lines with cantilever correction, backward facing steps and airfoil waviness. The most severe influences are observed for mold tilt towards the leading and thick bond lines. By applying the cantilever correction, the influence of thick bond lines is almost compensated. Airfoil waviness is very dependent on amplitude height and the location along the surface of the airfoil. Increased influence is observed for backward facing steps, once they are high enough to trigger boundary layer transition close to the leading edge.

  1. On the effect of leading edge blowing on circulation control airfoil aerodynamics

    NASA Technical Reports Server (NTRS)

    Mclachlan, B. G.

    1987-01-01

    In the present context the term circulation control is used to denote a method of lift generation that utilizes tangential jet blowing over the upper surface of a rounded trailing edge airfoil to determine the location of the boundary layer separation points, thus setting an effective Kutta condition. At present little information exists on the flow structure generated by circulation control airfoils under leading edge blowing. Consequently, no theoretical methods exist to predict airfoil performance under such conditions. An experimental study of the flow field generated by a two dimensional circulation control airfoil under steady leading and trailing edge blowing was undertaken. The objective was to fundamentally understand the overall flow structure generated and its relation to airfoil performance. Flow visualization was performed to define the overall flow field structure. Measurements of the airfoil forces were also made to provide a correlation of the observed flow field structure to airfoil performance. Preliminary results are presented, specifically on the effect on the flow field structure of leading edge blowing, alone and in conjunction with trailing edge blowing.

  2. Effects of enviromentally imposed roughness on airfoil performance

    NASA Technical Reports Server (NTRS)

    Cebeci, Tuncer

    1987-01-01

    The experimental evidence for the effects of rain, insects, and ice on airfoil performance are examined. The extent to which the available information can be incorporated in a calculation method in terms of change of shape and surface roughness is discussed. The methods described are based on the interactive boundary layer procedure of Cebeci or on the thin layer Navier Stokes procedure developed at NASA. Cases presented show that extensive flow separation occurs on the rough surfaces.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Negative tracheal pressure during neonatal endotracheal suction.

    PubMed

    Kiraly, Nicholas J; Tingay, David G; Mills, John F; Morley, Colin J; Copnell, Beverley

    2008-07-01

    Endotracheal tube (ETT) suction is the most frequently performed invasive procedure in ventilated newborn infants and is associated with adverse effects related to negative tracheal pressure. We aimed to measure suction catheter gas flow and intratracheal pressure during ETT suction of a test lung and develop a mathematical model to predict tracheal pressure from catheter and ETT dimensions and applied pressure. Tracheal pressure and catheter flow were recorded during suction of ETT sizes 2.5-4.0 mm connected to a test lung with catheters 5-8 French Gauge and applied pressures of 80-200 mm Hg. The fraction of applied pressure transmitted to the trachea was calculated for each combination, and data fitted to three nonlinear models for analysis. Tracheal pressure was directly proportional to applied pressure (r = 0.82-0.99), and catheter flow fitted a turbulent flow model (R = 0.85-0.96). With each ETT, increasing catheter size resulted in greater catheter flow (p < 0.0001) and thus lower intratracheal pressure (p < 0.0001). The fraction of applied pressure transmitted to the trachea was accurately modeled using ETT and catheter dimensions (R = 0.98-0.99). Negative tracheal pressure during in vitro ETT suction is directly proportional to applied pressure. This relationship is determined by ETT and catheter dimensions.

  13. 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.

  14. Analysis of Tank 38H (HTF-38-15-119, 127) Surface, Subsurface and Tank 43H (HTF-43-15-116, 117 and 118) Surface, Feed Pump Suction and Jet Suction Subsurface Supernatant Samples in Support of Enrichment, Corrosion Control and Salt Batch Planning Programs

    SciTech Connect

    Oji, L.

    2015-12-17

    Compositional feed limits have been established to ensure that a nuclear criticality event for the 2H and 3H Evaporators is not possible. The Enrichment Control Program (ECP) requires feed sampling to determine the equivalent enriched uranium content prior to transfer of waste other than recycle transfers (requires sampling to determine the equivalent enriched uranium at two locations in Tanks 38H and 43H every 26 weeks) The Corrosion Control Program (CCP) establishes concentration and temperature limits for key constituents and periodic sampling and analysis to confirm that waste supernate is within these limits. This report provides the results of analyses on Tanks 38H and 43H surface and subsurface supernatant liquid samples in support of the ECP, the CCP, and the Salt Batch 10 Planning Program.

  15. 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.

  16. 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.

  17. 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

  18. 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.

  19. 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.

  20. 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.

  1. SmaggIce 2.0: Additional Capabilities for Interactive Grid Generation of Iced Airfoils

    NASA Technical Reports Server (NTRS)

    Kreeger, Richard E.; Baez, Marivell; Braun, Donald C.; Schilling, Herbert W.; Vickerman, Mary B.

    2008-01-01

    The Surface Modeling and Grid Generation for Iced Airfoils (SmaggIce) software toolkit has been extended to allow interactive grid generation for multi-element iced airfoils. The essential phases of an icing effects study include geometry preparation, block creation and grid generation. SmaggIce Version 2.0 now includes these main capabilities for both single and multi-element airfoils, plus an improved flow solver interface and a variety of additional tools to enhance the efficiency and accuracy of icing effects studies. An overview of these features is given, especially the new multi-element blocking strategy using the multiple wakes method. Examples are given which illustrate the capabilities of SmaggIce for conducting an icing effects study for both single and multi-element airfoils.

  2. 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.

  3. Grid generation by elliptic partial differential equations for a tri-element Augmentor-Wing airfoil

    NASA Technical Reports Server (NTRS)

    Sorenson, R. L.

    1982-01-01

    Two efforts to numerically simulate the flow about the Augmentor-Wing airfoil in the cruise configuration using the GRAPE elliptic partial differential equation grid generator algorithm are discussed. The Augmentor-Wing consists of a main airfoil with a slotted trailing edge for blowing and two smaller airfoils shrouding the blowing jet. The airfoil and the algorithm are described, and the application of GRAPE to an unsteady viscous flow simulation and a transonic full-potential approach is considered. The procedure involves dividing a complicated flow region into an arbitrary number of zones and ensuring continuity of grid lines, their slopes, and their point distributions across the zonal boundaries. The method for distributing the body-surface grid points is discussed.

  4. Compressible laminar streaks with wall suction

    NASA Astrophysics Data System (ADS)

    Ricco, Pierre; Shah, Daniel; Hicks, Peter D.

    2013-05-01

    The response of a compressible laminar boundary layer subject to free-stream vortical disturbances and steady mean-flow wall suction is studied. The theoretical frameworks of Leib et al. [J. Fluid Mech. 380, 169-203 (1999), 10.1017/S0022112098003504] and Ricco and Wu [J. Fluid Mech. 587, 97-138 (2007), 10.1017/S0022112007007070], based on the linearized unsteady boundary-region equations, are adopted to study the influence of suction on the kinematic and thermal streaks arising through the interaction between the free-stream vortical perturbations and the boundary layer. In the asymptotic limit of small spanwise wavelength compared with the boundary layer thickness, i.e., when the disturbance flow is conveniently described by the steady compressible boundary region equations, the effect of suction is mild on the velocity fluctuations and negligible on the temperature fluctuations. When the spanwise wavelength is comparable with the boundary layer thickness, small suction values intensify the supersonic streaks, while higher transpiration levels always stabilize the disturbances at all Mach numbers. At larger spanwise wavelengths, very small amplitudes of wall transpiration have a dramatic stabilizing effect on all boundary layer fluctuations, which can take the form of transiently growing thermal streaks, large amplitude streamwise oscillations, or oblique exponentially growing Tollmien-Schlichting waves, depending on the Mach number and the wavelengths. The range of wavenumbers for which the exponential growth occurs becomes narrower and the location of instability is significantly shifted downstream by mild suction, indicating that wall transpiration can be a suitable vehicle for delaying transition when the laminar breakdown is promoted by these unstable disturbances. The typical streamwise wavelength of these disturbances is instead not influenced by suction, and asymptotic triple deck theory predicts the strong changes in growth rate and the very mild

  5. Porous plug for reducing orifice induced pressure error in airfoils

    NASA Technical Reports Server (NTRS)

    Plentovich, Elizabeth B. (Inventor); Gloss, Blair B. (Inventor); Eves, John W. (Inventor); Stack, John P. (Inventor)

    1988-01-01

    A porous plug is provided for the reduction or elimination of positive error caused by the orifice during static pressure measurements of airfoils. The porous plug is press fitted into the orifice, thereby preventing the error caused either by fluid flow turning into the exposed orifice or by the fluid flow stagnating at the downstream edge of the orifice. In addition, the porous plug is made flush with the outer surface of the airfoil, by filing and polishing, to provide a smooth surface which alleviates the error caused by imperfections in the orifice. The porous plug is preferably made of sintered metal, which allows air to pass through the pores, so that the static pressure measurements can be made by remote transducers.

  6. Wind-tunnel Tests of the NACA 45-125 Airfoil: A Thick Airfoil for High-Speed Airplanes

    NASA Technical Reports Server (NTRS)

    Delano, James B.

    1940-01-01

    Investigations of the pressure distribution, the profile drag, and the location of transition for a 30-inch-chord 25-percent-thick N.A,C.A. 45-125 airfoil were made in the N.A.C.A 8-foot high-speed wind tunnel for the purpose of aiding in the development of a thick wing for high-speed airplanes. The tests were made at a lift coefficient of 0.1 for Reynolds Numbers from 1,750,000 to 8,690,000, corresponding to speeds from 80 to 440 miles per hour at 59 F. The effect on the profile drag of fixing the transition point was also investigated. The effect of compressibility on the rate of increase of pressure coefficients was found to be greater than that predicted by a simplified theoretical expression for thin wings. The results indicated that, for a lift coefficient of 0.1, the critical speed of the N.A.C,A. 45-125 airfoil was about 460 miles per hour at 59 F,. The value of the profile-drag coefficient at a Reynolds Number of 4,500,000 was 0.0058, or about half as large as the value for the N.A,C,A. 0025 airfoil. The increase in the profile-drag coefficient for a given movement of the transition point was about three times as large as the corresponding increase for the N.A.C,A. 0012 airfoil. Transition determinations indicated that, for Reynolds Numbers up to ?,000,000, laminar boundary 1ayers were maintained over approximately 40 percent of the upper and the lower surfaces of the airfoil.

  7. Remotely operated submersible underwater suction apparatus

    SciTech Connect

    Kristan, Louis L.

    1990-01-01

    A completely submersible, remotely operated underwater suction device for collection of irradiated materials in a nuclear pool is disclosed. The device includes a pump means for pumping water through the device, a filter means for capturing irradiated debris, remotely operated releasable connector means, a collection means and a means for remotely maneuvering the collection means. The components of the suction device may be changed and replaced underwater to take advantage of the excellent radiation shielding ability of water to thereby minimize exposure of personnel to radiation.

  8. Computation of subsonic flow around airfoil systems with multiple separation

    NASA Technical Reports Server (NTRS)

    Jacob, K.

    1982-01-01

    A numerical method for computing the subsonic flow around multi-element airfoil systems was developed, allowing for flow separation at one or more elements. Besides multiple rear separation also sort bubbles on the upper surface and cove bubbles can approximately be taken into account. Also, compressibility effects for pure subsonic flow are approximately accounted for. After presentation the method is applied to several examples and improved in some details. Finally, the present limitations and desirable extensions are discussed.

  9. Pump tank divider plate for sump suction sodium pumps

    DOEpatents

    George, John A.; Nixon, Donald R.

    1977-01-01

    A circular plate extends across the diameter of "sump suction" pump, with a close clearance between the edge of the plate and the wall of the pump tank. The plate is located above the pump impeller, inlet and outlet flow nozzles but below the sodium free surface and effectively divides the pump tank into two separate chambers. On change of pump speed, the close fitting flow restriction plate limits the rate of flow into or out of the upper chamber, thereby minimizing the rate of level change in the tank and permitting time for the pump cover gas pressure to be varied to maintain an essentially constant level.

  10. Effects of Suction on Swept-Wing Transition

    NASA Technical Reports Server (NTRS)

    Saric, William S.

    1998-01-01

    Stability experiments are conducted in the Arizona State University Unsteady Wind Tunnel on a 45 deg swept airfoil. The pressure gradient is designed to provide purely crossflow-dominated transition; that is, the boundary layer is subcritical to Tollmien-Schlichting disturbances. The airfoil surface is hand polished to a 0.25 microns rms finish. Under these conditions, stationary crossflow disturbances grow to nonuniform amplitude due to submicron surface irregularities near the leading edge. Uniform stationary crossflow waves are produced by controlling the initial conditions with spanwise arrays of micron-sized roughness elements near the attachment line. Hot-wire measurements provide detailed maps of the crossflow wave structure, and accurate spectral decompositions isolate individual-mode growth rates for the fundamental and harmonic disturbances. Roughness spacing, roughness height, and Reynolds number are varied to investigate the growth of all amplified wavelengths. The measurements show early nonlinear mode interaction causing amplitude saturation well before transition. Comparisons with nonlinear parabolized stability equations calculations show excellent agreement in both the disturbance amplitude and the mode-shape profiles.

  11. 21 CFR 884.1175 - Endometrial suction curette and accessories.

    Code of Federal Regulations, 2013 CFR

    2013-04-01

    ... suction curette is a device used to remove material from the uterus and from the mucosal lining of the uterus by scraping and vacuum suction. This device is used to obtain tissue for biopsy or for...

  12. 21 CFR 884.1175 - Endometrial suction curette and accessories.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ... suction curette is a device used to remove material from the uterus and from the mucosal lining of the uterus by scraping and vacuum suction. This device is used to obtain tissue for biopsy or for...

  13. 21 CFR 884.1175 - Endometrial suction curette and accessories.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ... suction curette is a device used to remove material from the uterus and from the mucosal lining of the uterus by scraping and vacuum suction. This device is used to obtain tissue for biopsy or for...

  14. 21 CFR 884.1175 - Endometrial suction curette and accessories.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ... suction curette is a device used to remove material from the uterus and from the mucosal lining of the uterus by scraping and vacuum suction. This device is used to obtain tissue for biopsy or for...

  15. 21 CFR 884.1175 - Endometrial suction curette and accessories.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... suction curette is a device used to remove material from the uterus and from the mucosal lining of the uterus by scraping and vacuum suction. This device is used to obtain tissue for biopsy or for...

  16. On the lift increments with the occurrence of airfoil tones at low Reynodls numbers

    NASA Astrophysics Data System (ADS)

    Ikeda, Tomoaki; Fujimoto, Daisuke; Inasawa, Ayumu; Asai, Masahito

    2015-11-01

    The aeroacoustic effects on the aerodynamics of an NACA 0006 airfoil are investigated experimentally at relatively low Reynolds numbers, Re = 30 , 000 - 70 , 000 . By employing two wind-testing airfoil models at different chord lengths, L = 40 and 100 [mm], the aerodynamic dependence on Mach number is examined at a given Reynolds number. In a particular range of Reynolds number, tonal peaks of trailing-edge noise are obtained from a shorter-chord airfoil, while no apparent tones are observed with longer chord length at a lower Mach number. Surprisingly, the occurrence of a tonal noise leads to a greater lift slope in the present wind-tunnel experiment, evaluated via a PIV approach. The lift curves obtained experimentally at higher Mach numbers agree well with two-dimensional numerical simulations, performed at M = 0 . 2 . At the Mach number, the numerical results clearly indicate the occurrence of an acoustic feedback loop with discrete tones, within a range of angle of attack. A few three dimensional numerical results are also presented. In the simulation at Re = 50 , 000 , the suppression of tonal noise corresponds to the development of a turbulent wedge in the suction-side boundary layer at the angle of attack 4 . 0 [deg.], which agrees with the experiment. This work was supported by Grant-in-Aid for Scientific Research from Japan Society for the Promotion of Science (Grant No. 25420139).

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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…

  2. 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.

  3. An experimental study on a suction flow control method to reduce the unsteadiness of the wind loads acting on a circular cylinder

    NASA Astrophysics Data System (ADS)

    Chen, Wen-Li; Li, Hui; Hu, Hui

    2014-04-01

    An experimental investigation was conducted to assess the effectiveness of a suction flow control method for vortex-induced vibration (VIV) suppression. The flow control method uses a limited number of isolated suction holes to manipulate the vortex shedding in the wake behind a circular cylinder in order to reduce the unsteadiness of the dynamic wind loads acting on the cylinder. The experimental study was performed at Re ≈ 3.0 × 104, i.e., in the typical Reynolds number range of VIV for the cables of cable-stayed bridges. In addition to measuring the surface pressure distributions to determine the resultant dynamic wind loads acting on the test model, a digital particle image velocimetry system was used to conduct detailed flow field measurements to reveal the changes in the shedding process of the unsteady wake vortex structures from the test model with and without the suction flow control. The effects of important controlling parameters (i.e., the azimuthal locations of the suction holes in respect to the oncoming airflow, the spanwise spacing between the suction holes, and the suction flow rate through the suction holes) on the wake flow characteristics, the surface pressure distributions, and the resultant dynamic wind loads were assessed quantitatively. While a higher suction flow rate and smaller spanwise spacing between the suction holes were beneficial to the effectiveness of the suction flow control, the azimuthal locations of the suction holes were found to be very critical for reducing the fluctuating amplitudes of the dynamic wind loads acting on the test model using the suction flow control method. With the suction holes located at the proper azimuthal locations on the test model (i.e., at the azimuthal angle of θ = 90° and 270° for the present study), the characteristics of the wake flow behind the test model were found to change significantly along the entire span of the test model, even though only a limited number of the isolated suction

  4. 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.

  5. Experimental study of flow due to an isolated suction hole and a partially plugged suction slot

    NASA Technical Reports Server (NTRS)

    Goglia, G. L.; Wilkinson, S. P.

    1980-01-01

    Details for construction of a model of a partially plugged, laminar flow control, suction slot and an isolated hole are presented. The experimental wind tunnel facility and instrumentation is described. Preliminary boundary layer velocity profiles (without suction model) are presented and shown to be in good agreement with the Blasius laminar profile. Recommendations for the completion of the study are made. An experimental program for study of transition on a rotating disk is described along with preliminary disturbance amplification rate data.

  6. 21 CFR 870.5050 - Patient care suction apparatus.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... 21 Food and Drugs 8 2010-04-01 2010-04-01 false Patient care suction apparatus. 870.5050 Section 870.5050 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES... suction apparatus. (a) Identification. A patient care suction apparatus is a device used with...

  7. 21 CFR 870.5050 - Patient care suction apparatus.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ... 21 Food and Drugs 8 2011-04-01 2011-04-01 false Patient care suction apparatus. 870.5050 Section 870.5050 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES... suction apparatus. (a) Identification. A patient care suction apparatus is a device used with...

  8. 21 CFR 874.5350 - Suction antichoke device.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ...) MEDICAL DEVICES EAR, NOSE, AND THROAT DEVICES Therapeutic Devices § 874.5350 Suction antichoke device. (a) Identification. A suction antichoke device is a device intended to be used in an emergency situation to remove... 21 Food and Drugs 8 2010-04-01 2010-04-01 false Suction antichoke device. 874.5350 Section...

  9. 21 CFR 874.5350 - Suction antichoke device.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ...) MEDICAL DEVICES EAR, NOSE, AND THROAT DEVICES Therapeutic Devices § 874.5350 Suction antichoke device. (a) Identification. A suction antichoke device is a device intended to be used in an emergency situation to remove... 21 Food and Drugs 8 2011-04-01 2011-04-01 false Suction antichoke device. 874.5350 Section...

  10. Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications

    SciTech Connect

    Law, S.P.; Gregorek, G.M.

    1987-07-01

    An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. x 22 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar sharp trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30% thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.

  11. Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications

    NASA Astrophysics Data System (ADS)

    Law, S. P.; Gregorek, G. M.

    1987-07-01

    An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. by 21 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 deg at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30 percent thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.

  12. Effect of cavity on shock oscillation in transonic flow over RAE2822 supercritical airfoil

    NASA Astrophysics Data System (ADS)

    Rahman, M. Rizwanur; Labib, Md. Itmam; Hasan, A. B. M. Toufique; Ali, M.; Mitsutake, Y.; Setoguchi, T.

    2016-07-01

    Transonic flow past a supercritical airfoil is strongly influenced by the interaction of shock wave with boundary layer. This interaction induces unsteady self-sustaining shock wave oscillation, flow instability, drag rise and buffet onset which limit the flight envelop. In the present study, a computational analysis has been carried out to investigate the flow past a supercritical RAE2822 airfoil in transonic speeds. To control the shock wave oscillation, a cavity is introduced on the airfoil surface where shock wave oscillates. Different geometric configurations have been investigated for finding optimum cavity geometry and dimension. Unsteady Reynolds averaged Navier-Stokes equations (RANS) are computed at Mach 0.729 with an angle of attack of 5°. Computed results are well validated with the available experimental data in case of baseline airfoil. However, in case of airfoil with control cavity; it has been observed that the introduction of cavity completely suppresses the unsteady shock wave oscillation. Further, significant drag reduction and successive improvement of aerodynamic performance have been observed in airfoil with shock control cavity.

  13. Wind tunnel evaluation of a truncated NACA 64-621 airfoil for wind turbine applications

    NASA Technical Reports Server (NTRS)

    Law, S. P.; Gregorek, G. M.

    1987-01-01

    An experimental program to measure the aerodynamic performance of a NACA 64-621 airfoil with a truncated trailing edge for wind turbine applications has been conducted in the Ohio State University Aeronautical and Astronautical Research Laboratory 6 in. by 21 in. pressurized wind tunnel. The blunted or trailing edge truncated (TET) airfoil has an advantage over similar trailing edge airfoils because it is able to streamline a larger spar structure, while also providing aerodynamic properties that are quite good. Surface pressures were measured and integrated to determine the lift, pressure drag, and moment coefficients over angles of attack ranging from -14 to +90 deg at Mach 0.2 and Reynolds numbers of 1,000,000 and 600,000. Results are compared to the NACA 0025, 0030, and 0035 thick airfoils with sharp trailing edges. Comparison shows that the 30 percent thick NACA 64-621-TET airfoil has higher maximum lift, higher lift curve slope, lower drag at higher lift coefficients, and higher chordwise force coefficient than similar thick airfoils with sharp trailing edges.

  14. 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.

  15. 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.

  16. 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.

  17. Transonic flow about a thick circular-arc airfoil

    NASA Technical Reports Server (NTRS)

    Mcdevitt, J. B.; Levy, L. L., Jr.; Deiwert, G. S.

    1975-01-01

    An experimental and theoretical study of transonic flow over a thick airfoil, prompted by a need for adequately documented experiments that could provide rigorous verification of viscous flow simulation computer codes, is reported. Special attention is given to the shock-induced separation phenomenon in the turbulent regime. Measurements presented include surface pressures, streamline and flow separation patterns, and shadowgraphs. For a limited range of free-stream Mach numbers the airfoil flow field is found to be unsteady. Dynamic pressure measurements and high-speed shadowgraph movies were taken to investigate this phenomenon. Comparisons of experimentally determined and numerically simulated steady flows using a new viscous-turbulent code are also included. The comparisons show the importance of including an accurate turbulence model. When the shock-boundary layer interaction is weak the turbulence model employed appears adequate, but when the interaction is strong, and extensive regions of separation are present, the model is inadequate and needs further development.

  18. 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.

  19. 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.

  20. Control of Vortex Shedding on an Airfoil using Mini Flaps at Low Reynolds Number

    NASA Astrophysics Data System (ADS)

    Oshiyama, Daisuke; Numata, Daiju; Asai, Keisuke

    2015-11-01

    In this study, the effects of mini flaps (MFs) on a NACA0012 airfoil were investigated experimentally at low Reynolds number. MFs are small flat plates attached to the trailing edge of an airfoil perpendicularly. All the tests were conducted at the Tohoku-University Basic Aerodynamic Research Tunnel at the chord Reynolds number of 25,000. Aerodynamic forces were measured using a 3-component balance and the surface flow was visualized by luminescent oil film technique. The results of force measurement show that attachment of MFs enhances lift and the enhanced lift increases with MF height. On the other hand, the results of oil flow visualization show that attachment of MFs enlarges the separated region on the airfoil rather than diminishes it. To understand the physical mechanism of MFs for lift enhancement, the flow around the airfoil was visualized by the smoke-wire method and the wake profile behind the airfoil was measured using a hot wire anemometer. It was found that vortices shed periodically from the tip of the MFs and interact with the separated shear layer from the upper surface. This unsteady vortex shedding forms a low-pressure region on the upper surface, generating higher lift. These results suggest that the height of MFs controls the frequency of vortex shedding behind the MF, forcing the separated shear layer on the upper surface flow in unsteady manner.

  1. Separated shear layer transition over an airfoil at a low Reynolds number

    NASA Astrophysics Data System (ADS)

    Boutilier, Michael S. H.; Yarusevych, Serhiy

    2012-08-01

    Shear layer development over a NACA 0018 airfoil at a chord Reynolds number of 100 000 was investigated using a combination of flow visualization, velocity field mapping, surface pressure fluctuation measurements, and stability analysis. The results provide a detailed description of shear layer transition on an airfoil at low Reynolds numbers. An extensive comparison of measured surface pressure and velocity fluctuations demonstrated that time-resolved surface pressure sensor arrays can be used to identify the presence of flow separation, estimate the extent of the separated flow region, and measure disturbance growth rate spectra in significantly less time than is required by conventional techniques. Surface pressure sensor measurements of disturbance growth rate, wave number, and convection speed are found to compare well with predictions of linear stability theory, supporting the claim that convection speeds measured in separation bubbles over low Reynolds number airfoils are associated with wave packets of growing disturbances propagating through the shear layer. Through a comparison of measured convection speeds in this investigation and prior low Reynolds number airfoil experiments, it is shown that disturbance convection speeds of between 30% and 50% of the edge velocity are typical for this type of flow, consistent with phase speed estimates from previous analytical studies on transitional separation bubbles. Modal RMS velocity profiles were measured and found to be reasonably predicted by stability theory. The results suggest that, even for the relatively thick NACA 0018 airfoil profile, disturbance development over the majority of the laminar separated shear layer is primarily governed by a linear inviscid mechanism.

  2. Removing Boundary Layer by Suction

    NASA Technical Reports Server (NTRS)

    Ackeret, J

    1927-01-01

    Through the utilization of the "Magnus effect" on the Flettner rotor ship, the attention of the public has been directed to the underlying physical principle. It has been found that the Prandtl boundary-layer theory furnishes a satisfactory explanation of the observed phenomena. The present article deals with the prevention of this separation or detachment of the flow by drawing the boundary layer into the inside of a body through a slot or slots in its surface.

  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. 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.

  5. Instabilities orginating from suction holes used for Laminar Flow Control (LFC)

    NASA Technical Reports Server (NTRS)

    Watmuff, Jonathan H.

    1994-01-01

    A small-scale wind tunnel previously used for turbulent boundary layer studies has been modified for experiments in laminar flow control. The facility incorporates suction through interchangeable porous test surfaces which are used to stabilize the boundary layer and delay transition to turbulent flow. The thin porous test surfaces are supported by a baffled plenum chamber box which also acts to gather the flow through the surface into tubes which are routed to a high pressure fan. An elliptic leading edge is attached to the assembly to establish a new layer on the test plate. A slot is used to remove the test section flow below the leading edge. The test section was lengthened and fitted with a new ceiling. Substantial modifications were also made to the 3D probe traverse. Detailed studies have been made using isolated holes to explore the underlying instability mechanisms. The suction is perturbed harmonically and data are averaged on the basis of the phase of the disturbance. Conditions corresponding to strong suction and without suction have been studied. In both cases, 3D contour surfaces in the vicinity of the hole show highly three-dimensional T-S waves that fan out away from the hole with streamwise distance. With suction, the perturbations on the centerline are much stronger and decay less rapidly, while the far field is similar to the case without suction. Downstream the contour surfaces of the bow-shaped TS waves develop spanwise irregularities which eventually form into clumps. The contours remain smooth when suction is not applied. Even without suction, the harmonic point source is challenging for CFD; e.g. DNS has been used for streamwise growth. With suction, grid resources are consumed by the hole and this makes DNS even more expensive. Limited DNS results so far indicate that the vortices which emanate from suction holes appear to be stable. The spanwise clumping observed in the experiment is evidence of a secondary instability that could be

  6. Some recent applications of the suction analogy to vortex-lift estimates

    NASA Technical Reports Server (NTRS)

    Lamar, J. E.

    1976-01-01

    An extension of the suction analogy for the estimation of vortex lift along the side edge of wings is reviewed along with the concept of an augmented vortex lift to account for the effect of the leading-edge vortex passing downstream over an aft part of the model. Applications of these extensions have resulted in an improved estimating capability for a wide range of isolated sharp-edge planforms and also for multiple lifting surfaces. Hence, the suction analogy concept can now have wider applicability at both subsonic and supersonic speeds, especially in the preliminary design cycle.

  7. 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.

  8. Bernoulli Suction Effect on Soap Bubble Blowing?

    NASA Astrophysics Data System (ADS)

    Davidson, John; Ryu, Sangjin

    2015-11-01

    As a model system for thin-film bubble with two gas-liquid interfaces, we experimentally investigated the pinch-off of soap bubble blowing. Using the lab-built bubble blower and high-speed videography, we have found that the scaling law exponent of soap bubble pinch-off is 2/3, which is similar to that of soap film bridge. Because air flowed through the decreasing neck of soap film tube, we studied possible Bernoulli suction effect on soap bubble pinch-off by evaluating the Reynolds number of airflow. Image processing was utilized to calculate approximate volume of growing soap film tube and the volume flow rate of the airflow, and the Reynolds number was estimated to be 800-3200. This result suggests that soap bubbling may involve the Bernoulli suction effect.

  9. 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.

  10. [Contusion-suction trauma after globe injuries].

    PubMed

    Kroll, P; Stoll, W; Kirchhoff, E

    1983-06-01

    An analysis of ball injuries treated during the last 3 years at Münster University Eye Hospital revealed a difference in the kind of traumata caused by air-filled balls and by solid, inelastic balls. The pathomechanism of a "contusion-suction trauma" is discussed; this would offer a satisfactory explanation not only for injuries of the anterior segment, but also for retinal changes at the outer periphery and the posterior pole.

  11. 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.

  12. Ultra-fast underwater suction traps.

    PubMed

    Vincent, Olivier; Weisskopf, Carmen; Poppinga, Simon; Masselter, Tom; Speck, Thomas; Joyeux, Marc; Quilliet, Catherine; Marmottant, Philippe

    2011-10-01

    Carnivorous aquatic Utricularia species catch small prey animals using millimetre-sized underwater suction traps, which have fascinated scientists since Darwin's early work on carnivorous plants. Suction takes place after mechanical triggering and is owing to a release of stored elastic energy in the trap body accompanied by a very fast opening and closing of a trapdoor, which otherwise closes the trap entrance watertight. The exceptional trapping speed--far above human visual perception--impeded profound investigations until now. Using high-speed video imaging and special microscopy techniques, we obtained fully time-resolved recordings of the door movement. We found that this unique trapping mechanism conducts suction in less than a millisecond and therefore ranks among the fastest plant movements known. Fluid acceleration reaches very high values, leaving little chance for prey animals to escape. We discovered that the door deformation is morphologically predetermined, and actually performs a buckling/unbuckling process, including a complete trapdoor curvature inversion. This process, which we predict using dynamical simulations and simple theoretical models, is highly reproducible: the traps are autonomously repetitive as they fire spontaneously after 5-20 h and reset actively to their ready-to-catch condition.

  13. An Ultrasonic Suction Pump with No Physically Moving Parts

    NASA Astrophysics Data System (ADS)

    Yun, Cheol-Ho; Hasegawa, Takeshi; Nakamura, Kentaro; Ueha, Sadayuki

    2004-05-01

    A new ultrasonic suction pump is described in this paper. The pump uses the suction force of a rigid cylinder tube vibrating at an ultrasonic frequency and has no physically moving parts. The pump consists of a longitudinal bolt-clamped Langevin transducer (BLT) combined with a stepped horn working at a resonance frequency of 24 kHz. A glass tube with the length of the half-wavelength-resonance is glued at the tip of the horn. To enhance pump performance, we introduced a reflection plate and a thin rod installed to the end of the glass tube with a small gap. Maximum pressures of 7.2 kPa and 23.5 kPa were recorded using the reflection plate and the thin rod, respectively. In this study, we experimentally investigate the characteristics of the pump and the operating physics. The maximum pressure is a function of the vibration velocity of the end surface of the glass tube and of the gap.

  14. Behavior of Water Jet Accompanied with Air Suction

    NASA Astrophysics Data System (ADS)

    Kawakami, Hironobu; Ishido, Tsutomu; Ihara, Akio

    In order to atomize a liquid, the authors have investigated the behavior of air-water jets. In a series of experiments, we have discovered a strange phenomenon that the water jet accompanied with air suction from the free surface has made a periodic radial splash of water drop. The purpose of the present paper is to clear out the origin of this phenomenon and the behavior of water jet accompanied with air suction. The behavior of water jet has been photographed by a digital camera aided with a flashlight and high-speed video camera. Those experiments enable us to find the origin of a periodic radial splash due to a formation of single air bubble at the flow separation region inside the nozzle and due to explosive expansion of the bubble after injected in the free space. In order to analyze the radial splash of water, we have conducted the equation of spherical liquid membrane. The numerical results obtained have been compared with the experimental results and good agreement has been obtained in radial expansion velocity.

  15. Adaptive Suction and Blowing for Twin-Tail Buffet Control

    NASA Technical Reports Server (NTRS)

    Kandil, Osama A.; Yang, Zhi

    1999-01-01

    Adaptive active flow control for twin-tail buffet alleviation is investigated. The concept behind this technique is to place control ports on the tail outer and inner surfaces with flow suction or blowing applied through these ports in order to minimize the pressure difference across the tail. The suction or blowing volume flow rate from each port is proportional to the pressure difference across the tail at this location. A parametric study of the effects of the number and location of these ports on the buffet response is carried out. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. This complex multidisciplinary problem is solved sequentially using three sets of equations for the fluid flow, aeroelastic response and grid deformation, using a dynamic multi-block grid structure. The computational model is pitched at 30 deg angle of attack. The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. The model is investigated for the inboard position of the twin tails, which corresponds to a separation distance between the twin tails of 33% of the wing span. Comparison of the time history and power spectral density responses of the tails for various distributions of the control ports are presented and discussed.

  16. 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.

  17. Reduction of acoustic disturbances in the test section of supersonic wind tunnels by laminarizing their nozzle and test section wall boundary layers by means of suction

    NASA Technical Reports Server (NTRS)

    Pfenninger, W.; Syberg, J.

    1974-01-01

    The feasibility of quiet, suction laminarized, high Reynolds number (Re) supersonic wind tunnel nozzles was studied. According to nozzle wall boundary layer development and stability studies, relatively weak area suction can prevent amplified nozzle wall TS (Tollmien-Schlichting) boundary layer oscillations. Stronger suction is needed in and shortly upstream of the supersonic concave curvature nozzle area to avoid transition due to amplified TG (Taylor-Goertler) vortices. To control TG instability, moderately rapid and slow expansion nozzles require smaller total suction rates than rapid expansion nozzles, at the cost of larger nozzle length Re and increased TS disturbances. Test section mean flow irregularities can be minimized with suction through longitudinal or highly swept slots (swept behind local Mach cone) as well as finely perforated surfaces. Longitudinal slot suction is optimized when the suction-induced crossflow velocity increases linearly with surface distance from the slot attachment line toward the slot (through suitable slot geometry). Suction in supersonic blowdown tunnels may be operated by one or several individual vacuum spheres.

  18. 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.

  19. 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.

  20. 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.

  1. Application of Novel CO2 Laser-Suction Device.

    PubMed

    Straus, David; Moftakhar, Roham; Fink, Yoel; Patel, Deval; Byrne, Richard W

    2013-12-01

    Background Development of the flexible CO2 fiber has presented new opportunities for the use of precision laser cutting in cranial procedures. The efficacy of the CO2 scalpel is further enhanced by combining it with a fluid removal suction capability. Objectives We report our experience with a novel CO2 laser-suction device. Methods The novel laser-suction device was designed in conjunction with OmniGuide Inc. (Cambridge, Massachusetts, USA). We performed a case review of its use in firm tumors that were resistant to resection by bipolar, suction, and ultrasonic aspirator. Results The laser-suction device was applied in three tumors where resection with ultrasonic aspiration failed. Tumor resection using the laser-suction device was successful in all three cases. There were no complications related to the laser-suction device. There were no instances of intraoperative device malfunction. Discussion The CO2 laser combined with suction is a useful instrument for resection of firm tumors that prove to be resistant to ultrasonic aspiration. We also find it to be useful in settings where precise tissue incisions are desired with minimal manipulation. In our experience, the surgical efficiency of the CO2 laser is improved by the laser-suction device. This device allows the surgeon to utilize a suction device and laser in a single hand and enables concurrent use of bipolar electrocautery without repeated instrument changes.

  2. Transient technique for measuring heat transfer coefficients on stator airfoils in a jet engine environment

    NASA Astrophysics Data System (ADS)

    Gladden, H. J.; Proctor, M. P.

    A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.

  3. Transient technique for measuring heat transfer coefficients on stator airfoils in a jet engine environment

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.; Proctor, M. P.

    1985-01-01

    A transient technique was used to measure heat transfer coefficients on stator airfoils in a high-temperature annular cascade at real engine conditions. The transient response of thin film thermocouples on the airfoil surface to step changes in the gas stream temperature was used to determine these coefficients. In addition, gardon gages and paired thermocouples were also utilized to measure heat flux on the airfoil pressure surface at steady state conditions. The tests were conducted at exit gas stream Reynolds numbers of one-half to 1.9 million based on true chord. The results from the transient technique show good comparison with the steady-state results in both trend and magnitude. In addition, comparison is made with the STAN5 boundary layer code and shows good comparison with the trends. However, the magnitude of the experimental data is consistently higher than the analysis.

  4. Water-tunnel experiments on an oscillating airfoil at RE equals 21,000

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

    Flow visualization experiments were performed in a water tunnel on a modified NACA 0012 airfoil undergoing large amplitude harmonic oscillations in pitch. Hydrogen bubbles were used to: (1) create a conveniently striated and well preserved set of inviscid flow markers; and (2) to expose the succession of events occurring within the viscous domain during the onset of dynamic stall. Unsteady effects were shown to have an important influence on the progression of flow reversal along the airfoil surface prior to stall. A region of reversed flow underlying a free shear layer was found to momentarily exist over the entire upper surface without any appreciable disturbance of the viscous-inviscid boundary. A flow protuberance was observed to develop near the leading edge, while minor vortices evolve from an expanding instability of the free shear layer over the rear portion of the airfoil. The complete breakdown of this shear layer culminates in the successive formation of two dominant vortices.

  5. 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.

  6. 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.

  7. Evaluation of hypopharyngeal suction to eliminate aspiration: the Retro-Esophageal Suction (REScue) catheter.

    PubMed

    Belafsky, Peter C; Mehdizadeh, O B; Ledgerwood, L; Kuhn, M

    2015-02-01

    Profound oropharyngeal dysphagia (OPD) is common and costly. Treatment options are limited. The purpose of this investigation was to evaluate the utility of hypopharyngeal suction at the upper esophageal sphincter (UES) to eliminate aspiration. Five different catheters were passed retrograde up the esophagus and positioned at the UES in a cadaver model of profound OPD. Suction was affixed to each catheter. 10 cc of barium was administered into the pyriform sinus, and videofluoroscopy was utilized to evaluate the presence of aspiration. 6 trials were administered per catheter and for a no catheter control. The outcome measures were the incidence of aspiration, the NIH Swallow Safety Scale (NIH-SSS), and UES opening. Control trials with no suction resulted in an aspiration rate of 100 % (6/6 trials). Negative pressure through 16, 18, 24, and 30 Fr catheter resulted in an aspiration rate of 0 % (0/24 trials; p < 0.001), and suction through a 12-Fr catheter resulted in an aspiration rate of 33 % (2/6 trials; p > 0.05). The mean NIH-SSS improved from 7.0 (±0.0) in the control to 0 (±0.0) with hypopharyngeal suction (18 Fr nasogastric catheter; p < 0.001). Mean UES opening improved from 0.0 (±0.0) mm in the control condition to 8.6 (±0.2) mm with a hypopharyngeal catheter (16 Fr Foley catheter; p < 0.001). Negative pressure applied through retro-esophageal suction catheters (>12 Fr) at the level of the UES reduced aspiration by 100 % and significantly increased UES opening in a cadaveric model of profound oropharyngeal dysphagia.

  8. 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.

  9. 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

  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. 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.

  12. Wind-Tunnel Tests on Airfoil Boundary Layer Control Using a Backward-Opening Slot

    NASA Technical Reports Server (NTRS)

    Bamber, Millard J

    1932-01-01

    This report presents the results of an investigation to determine the effect of boundary layer control on the lift and drag of an airfoil. Boundary layer control was accomplished by means of a backward-opening slot in the upper surface of the hollow airfoil. Air was caused to flow through this slot by a pressure which was maintained inside the airfoil by a blower. Various slot locations, slot openings, and wing pressures were used. The tests were conducted in the 5-foot atmospheric wind tunnel of the Langley Memorial Aeronautical Laboratory. Under the test conditions, the maximum lift coefficient was increased about 96 per cent for one slot arrangement, and the minimum drag coefficient was decreased about 27 per cent for another, both being compared with the results obtained with the unslotted airfoil. It is believed from this investigation that the above effects may be increased by the use of larger slot openings, better slot locations, multiple slots, improved airfoil profiles, and trailing edge flaps.

  13. Efficient simulation of incompressible viscous flow over single and multi-element airfoils

    NASA Technical Reports Server (NTRS)

    Rogers, Stuart E.; Wiltberger, N. L.; Kwak, Dochan

    1992-01-01

    Incompressible viscous turbulent flows over single- and multiple-element airfoils are numerically simulated in an efficient manner by solving the incompressible Navier-Stokes equations. The solution algorithm uses the method of pseudocompressibility with an upwind-differencing scheme for the convective fluxes and an implicit line-relaxation scheme to study high-lift take-off and landing configurations and to compute lift and drag at various angles of attack up to stall. Two different turbulence models are tested in computing the flow over an NACA 4412 airfoil. The approach used for multiple-element airfoils involves the use of multiple zones of structured grids fitted to each element. Two different approaches are compared: a patched system of grids and an overlaid Chimera system of grids. Computational results are presented for two-element, three-element, and four-element airfoil configurations. Excellent agreement with experimental surface-pressure coefficients is seen. The code converges in less than 200 iterations, requiring on the order of one minute of CPU time on a CRAY YMP per element in the airfoil configuration.

  14. Efficient simulation of incompressible viscous flow over multi-element airfoils

    NASA Technical Reports Server (NTRS)

    Rogers, Stuart E.; Wiltberger, N. Lyn; Kwak, Dochan

    1993-01-01

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

  15. Efficient simulation of incompressible viscous flow over multi-element airfoils

    NASA Technical Reports Server (NTRS)

    Rogers, Stuart E.; Wiltberger, N. Lyn; Kwak, Dochan

    1992-01-01

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

  16. 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.

  17. 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.

  18. Flat slab deformation caused by interplate suction force

    NASA Astrophysics Data System (ADS)

    Ma, Yiran; Clayton, Robert W.

    2015-09-01

    We image the structure at the southern end of the Peruvian flat subduction zone, using receiver function and surface wave methods. The Nazca slab subducts to ~100 km depth and then remains flat for ~300 km distance before it resumes the dipping subduction. The flat slab closely follows the topography of the continental Moho above, indicating a strong suction force between the slab and the overriding plate. A high-velocity mantle wedge exists above the initial half of the flat slab, and the velocity resumes to normal values before the slab steepens again, indicating the resumption of dehydration and ecologitization. Two prominent midcrust structures are revealed in the 70 km thick crust under the Central Andes: molten rocks beneath the Western Cordillera and the underthrusting Brazilian Shield beneath the Eastern Cordillera.

  19. 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%.

  20. Film cooling air pocket in a closed loop cooled airfoil

    DOEpatents

    Yu, Yufeng Phillip; Itzel, Gary Michael; Osgood, Sarah Jane; Bagepalli, Radhakrishna; Webbon, Waylon Willard; Burdgick, Steven Sebastian

    2002-01-01

    Turbine stator vane segments have radially inner and outer walls with vanes extending between them. The inner and outer walls are compartmentalized and have impingement plates. Steam flowing into the outer wall plenum passes through the impingement plate for impingement cooling of the outer wall upper surface. The spent impingement steam flows into cavities of the vane having inserts for impingement cooling the walls of the vane. The steam passes into the inner wall and through the impingement plate for impingement cooling of the inner wall surface and for return through return cavities having inserts for impingement cooling of the vane surfaces. To provide for air film cooing of select portions of the airfoil outer surface, at least one air pocket is defined on a wall of at least one of the cavities. Each air pocket is substantially closed with respect to the cooling medium in the cavity and cooling air pumped to the air pocket flows through outlet apertures in the wall of the airfoil to cool the same.

  1. 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.

  2. 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.

  3. A consistent design procedure for supercritical airfoils in free air and a wind tunnel

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Malmuth, N. D.; Cole, J. D.

    1979-01-01

    A computational inverse procedure for transonic airfoils in which shapes are determined supporting prescribed pressure distributions is presented. The method uses the small disturbance equation and a consistent analysis-design differencing procedure at the airfoil surface. This avoids the intermediate analysis-design-analysis iterations. The effect of any openness at the trailing edge is taken onto account by adding an effective source term in the far field. The final results from a systematic expansion procedure which models the far field for solid, ideal slotted, and free jet tunnel walls are presented along with some design results for the associated boundary conditions and those for a free flight.

  4. Turbine airfoil with an internal cooling system having vortex forming turbulators

    SciTech Connect

    Lee, Ching-Pang

    2014-12-30

    A turbine airfoil usable in a turbine engine and having at least one cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels having a plurality of turbulators protruding from an inner surface and positioned generally nonorthogonal and nonparallel to a longitudinal axis of the airfoil cooling channel. The configuration of turbulators may create a higher internal convective cooling potential for the blade cooling passage, thereby generating a high rate of internal convective heat transfer and attendant improvement in overall cooling performance. This translates into a reduction in cooling fluid demand and better turbine performance.

  5. 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.

  6. 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.

  7. Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow

    NASA Technical Reports Server (NTRS)

    Callaghan, Edmund E.; Serafini, John S.

    1953-01-01

    Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.

  8. Computational Modeling For The Transitional Flow Over A Multi-Element Airfoil

    NASA Technical Reports Server (NTRS)

    Liou, William W.; Liu, Feng-Jun; Rumsey, Chris L. (Technical Monitor)

    2000-01-01

    The transitional flow over a multi-element airfoil in a landing configuration are computed using a two equation transition model. The transition model is predictive in the sense that the transition onset is a result of the calculation and no prior knowledge of the transition location is required. The computations were performed using the INS2D) Navier-Stokes code. Overset grids are used for the three-element airfoil. The airfoil operating conditions are varied for a range of angle of attack and for two different Reynolds numbers of 5 million and 9 million. The computed results are compared with experimental data for the surface pressure, skin friction, transition onset location, and velocity magnitude. In general, the comparison shows a good agreement with the experimental data.

  9. Theoretical Prediction of Pressure Distributions on Nonlifting Airfoils at High Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Spreiter, John R; Alksne, Alberta

    1955-01-01

    Theoretical pressure distributions on nonlifting circular-arc airfoils in two-dimensional flows with high subsonic free-stream velocity are found by determining approximate solutions, through an iteration process, of an integral equation for transonic flow proposed by Oswatitsch. The integral equation stems directly from the small-disturbance theory for transonic flow. This method of analysis possesses the advantage of remaining in the physical, rather than the hodograph, variable and can be applied in airfoils having curved surfaces. After discussion of the derivation of the integral equation and qualitative aspects of the solution, results of calculations carried out for circular-arc airfoils in flows with free-stream Mach numbers up to unity are described. These results indicate most of the principal phenomena observed in experimental studies.

  10. 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.

  11. CFD simulation of flow-induced vibration of an elastically supported airfoil

    NASA Astrophysics Data System (ADS)

    Šidlof, Petr

    2016-03-01

    Flow-induced vibration of lifting or control surfaces in aircraft may lead to catastrophic consequences. Under certain circumstances, the interaction between the airflow and the elastic structure may lead to instability with energy transferred from the airflow to the structure and with exponentially increasing amplitudes of the structure. In the current work, a CFD simulation of an elastically supported NACA0015 airfoil with two degrees of freedom (pitch and plunge) coupled with 2D incompressible airflow is presented. The geometry of the airfoil, mass, moment of inertia, location of the centroid, linear and torsional stiffness was matched to properties of a physical airfoil model used for wind-tunnel measurements. The simulations were run within the OpenFOAM computational package. The results of the CFD simulations were compared with the experimental data.

  12. Comparison of Evolutionary (Genetic) Algorithm and Adjoint Methods for Multi-Objective Viscous Airfoil Optimizations

    NASA Technical Reports Server (NTRS)

    Pulliam, T. H.; Nemec, M.; Holst, T.; Zingg, D. W.; Kwak, Dochan (Technical Monitor)

    2002-01-01

    A comparison between an Evolutionary Algorithm (EA) and an Adjoint-Gradient (AG) Method applied to a two-dimensional Navier-Stokes code for airfoil design is presented. Both approaches use a common function evaluation code, the steady-state explicit part of the code,ARC2D. The parameterization of the design space is a common B-spline approach for an airfoil surface, which together with a common griding approach, restricts the AG and EA to the same design space. Results are presented for a class of viscous transonic airfoils in which the optimization tradeoff between drag minimization as one objective and lift maximization as another, produces the multi-objective design space. Comparisons are made for efficiency, accuracy and design consistency.

  13. Airfoil wake and linear theory gust response including sub and superresonant flow conditions

    NASA Technical Reports Server (NTRS)

    Henderson, Gregory H.; Fleeter, Sanford

    1992-01-01

    The unsteady aerodynamic gust response of a high solidity stator vane row is examined in terms of the fundamental gust modeling assumptions with particular attention given to the effects near an acoustic resonance. A series of experiments was performed with gusts generated by rotors comprised of perforated plates and airfoils. It is concluded that, for both the perforated plate and airfoil wake generated gusts, the unsteady pressure responses do not agree with the linear-theory gust predictions near an acoustic resonance. The effects of the acoustic resonance phenomena are clearly evident on the airfoil surface unsteady pressure responses. The transition of the measured lift coefficients across the acoustic resonance from the subresonant regime to the superresonant regime occurs in a simple linear fashion.

  14. 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.

  15. Suction power output and the inertial cost of rotating the neurocranium to generate suction in fish.

    PubMed

    Van Wassenbergh, Sam; Day, Steven W; Hernández, L Patricia; Higham, Timothy E; Skorczewski, Tyler

    2015-05-01

    To expand the buccal cavity, many suction-feeding fishes rely on a considerable contribution from dorsal rotation of the dorsal part of the head including the brains, eyes, and several bones forming the braincase and skull roof (jointly referred to as the neurocranium). As the neurocranium takes up a large part of the total mass of the head, this rotation may incur a considerable inertial cost. If so, this would suggest a significant selective pressure on the kinematics and mass distribution of the neurocranium of suction feeders. Here, an inverse dynamic model is formulated to calculate the instantaneous power required to rotate the neurocranium, approximated by a quarter ellipsoid volume of homogeneous density, as well as to calculate the instantaneous suction power based on intra-oral pressure and head volume quantifications. We applied this model to largemouth bass (Micropterus salmoides) and found that the power required to rotate the neurocranium accounts for only about 4% of the power required to suck water into the mouth. Furthermore, recovery of kinetic energy from the rotating neurocranium converted into suction work may be possible during the phase of neurocranial deceleration. Thus, we suggest that only a negligible proportion of the power output of the feeding muscles is lost as inertial costs in the largemouth bass. Consequently, the feeding performance of piscivorous suction feeders with generalised morphology, comparable to our model species, is not limited by neurocranial motion during head expansion. This suggests that it is thus not likely to be a factor of importance in the evolution of cranial shape and size. PMID:25769945

  16. 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.

  17. 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.

  18. 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%.

  19. 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.

  20. Flight tests of a supersonic natural laminar flow airfoil

    NASA Astrophysics Data System (ADS)

    Frederick, M. A.; Banks, D. W.; Garzon, G. A.; Matisheck, J. R.

    2015-06-01

    A flight test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80 inch (203 cm) chord and 40 inch (102 cm) span article mounted on the centerline store location of an F-15B airplane. The test article was designed with a leading edge sweep of effectively 0° to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate that the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, was similar to that of subsonic natural laminar flow wings.

  1. 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.

  2. Flight Tests of a Supersonic Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

    Frederick, M. A.; Banks, D. W.; Garzon, G. A.; Matisheck, J. R.

    2014-01-01

    A flight test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline store location of an F-15B airplane. The wing was designed with a leading edge sweep of effectively 0 deg to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, were similar to that of subsonic natural laminar flow wings.

  3. Flight Tests of a Supersonic Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

    Frederick, Michael A.; Banks, Daniel W.; Garzon, G. A.; Matisheck, J. R.

    2015-01-01

    A flight-test campaign of a supersonic natural laminar flow airfoil has been recently completed. The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline store location of an F-15B airplane (McDonnell Douglas Corporation, now The Boeing Company, Chicago, Illinois). The test article was designed with a leading edge sweep of effectively 0 deg to minimize boundary layer crossflow. The test article surface was coated with an insulating material to avoid significant heat transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared camera system was used to determine boundary layer transition and the extent of laminar flow. The tests were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, were similar to that of subsonic natural laminar flow wings.

  4. Measurements in Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions

    NASA Technical Reports Server (NTRS)

    Volino, Ralph J.; Hultgren, Lennart .

    2000-01-01

    Detailed velocity measurements were made along a flat plate subject to the same dimensionless pressure gradient as the suction side of a modern low-pressure turbine airfoil. Reynolds numbers based on wetted plate length and nominal exit velocity were varied from 50,000 to 300,000, covering cruise to takeoff conditions. Low and high inlet free-stream turbulence intensities (0.2% and 7%) were set using passive grids. The location of boundary-layer separation does not depend strongly on the free-stream turbulence level or Reynolds number, as long as the boundary layer remains non-turbulent prior to separation. Strong acceleration prevents transition on the upstream part of the plate in all cases. Both free-stream turbulence and Reynolds number have strong effects on transition in the adverse pressure gradient region. Under low free-stream turbulence conditions transition is induced by instability waves in the shear layer of the separation bubble. Reattachment generally occurs at the transition start. At Re = 50,000 the separation bubble does not close before the trailing edge of the modeled airfoil. At higher Re, transition moves upstream, and the boundary layer reattaches. With high free-stream turbulence levels, transition appears to occur in a bypass mode, similar to that in attached boundary layers. Transition moves upstream, resulting in shorter separation regions. At Re above 200,000, transition begins before separation. Mean velocity, turbulence and intermittency profiles are presented.

  5. 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.

  6. The effects of Reynolds number, rotor incidence angle, and surface roughness on the heat transfer distribution in a large-scale turbine rotor passage

    NASA Technical Reports Server (NTRS)

    Blair, Michael F.; Anderson, Olof L.

    1989-01-01

    A combined experimental and computational program was conducted to examine the heat transfer distribution in a turbine rotor passage geometrically similiar to the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP). Heat transfer was measured and computed for both the full-span suction and pressure surfaces of the rotor airfoil as well as for the hub endwall surface. The primary objective of the program was to provide a benchmark-quality data base for the assessment of rotor passage heat transfer computational procedures. The experimental portion of the study was conducted in a large-scale, ambient temperature, rotating turbine model. Heat transfer data were obtained using thermocouple and liquid-crystal techniques to measure temperature distributions on the thin, electrically-heated skin of the rotor passage model. Test data were obtained for various combinations of Reynolds number, rotor incidence angle and model surface roughness. The data are reported in the form of contour maps of Stanton number. These heat distribution maps revealed numerous local effects produced by the three-dimensional flows within the rotor passage. Of particular importance were regions of local enhancement produced on the airfoil suction surface by the main-passage and tip-leakage vortices and on the hub endwall by the leading-edge horseshoe vortex system. The computational portion consisted of the application of a well-posed parabolized Navier-Stokes analysis to the calculation of the three-dimensional viscous flow through ducts simulating the a gas turbine passage. These cases include a 90 deg turning duct, a gas turbine cascade simulating a stator passage, and a gas turbine rotor passage including Coriolis forces. The calculated results were evaluated using experimental data of the three-dimensional velocity fields, wall static pressures, and wall heat transfer on the suction surface of the turbine airfoil and on the end wall. Particular attention was paid to an

  7. 21 CFR 880.5740 - Suction snakebite kit.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... 21 Food and Drugs 8 2010-04-01 2010-04-01 false Suction snakebite kit. 880.5740 Section 880.5740 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES GENERAL HOSPITAL AND PERSONAL USE DEVICES General Hospital and Personal Use Therapeutic Devices § 880.5740 Suction snakebite kit....

  8. Transient bacteremia following endotracheal suctioning in ventilated newborns.

    PubMed

    Storm, W

    1980-03-01

    Endotracheal suctioning is a routine procedure in ventilated newborns. A study of ten neonates demonstrates the association of transient bacteremia with endotracheal suctioning. This complication in ventilated newborns, with colonization of the respiratory tract by the same organism, must be considered in the pathogenesis of systemic infection.

  9. 21 CFR 870.4430 - Cardiopulmonary bypass intracardiac suction control.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... 21 Food and Drugs 8 2010-04-01 2010-04-01 false Cardiopulmonary bypass intracardiac suction control. 870.4430 Section 870.4430 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND....4430 Cardiopulmonary bypass intracardiac suction control. (a) Identification. A cardiopulmonary...

  10. 21 CFR 874.5350 - Suction antichoke device.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ... 21 Food and Drugs 8 2014-04-01 2014-04-01 false Suction antichoke device. 874.5350 Section 874.5350 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES EAR, NOSE, AND THROAT DEVICES Therapeutic Devices § 874.5350 Suction antichoke device....

  11. 21 CFR 874.5350 - Suction antichoke device.

    Code of Federal Regulations, 2013 CFR

    2013-04-01

    ... 21 Food and Drugs 8 2013-04-01 2013-04-01 false Suction antichoke device. 874.5350 Section 874.5350 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES EAR, NOSE, AND THROAT DEVICES Therapeutic Devices § 874.5350 Suction antichoke device....

  12. 21 CFR 874.5350 - Suction antichoke device.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ... 21 Food and Drugs 8 2012-04-01 2012-04-01 false Suction antichoke device. 874.5350 Section 874.5350 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES EAR, NOSE, AND THROAT DEVICES Therapeutic Devices § 874.5350 Suction antichoke device....

  13. Swimming muscles power suction feeding in largemouth bass.

    PubMed

    Camp, Ariel L; Roberts, Thomas J; Brainerd, Elizabeth L

    2015-07-14

    Most aquatic vertebrates use suction to capture food, relying on rapid expansion of the mouth cavity to accelerate water and food into the mouth. In ray-finned fishes, mouth expansion is both fast and forceful, and therefore requires considerable power. However, the cranial muscles of these fishes are relatively small and may not be able to produce enough power for suction expansion. The axial swimming muscles of these fishes also attach to the feeding apparatus and have the potential to generate mouth expansion. Because of their large size, these axial muscles could contribute substantial power to suction feeding. To determine whether suction feeding is powered primarily by axial muscles, we measured the power required for suction expansion in largemouth bass and compared it to the power capacities of the axial and cranial muscles. Using X-ray reconstruction of moving morphology (XROMM), we generated 3D animations of the mouth skeleton and created a dynamic digital endocast to measure the rate of mouth volume expansion. This time-resolved expansion rate was combined with intraoral pressure recordings to calculate the instantaneous power required for suction feeding. Peak expansion powers for all but the weakest strikes far exceeded the maximum power capacity of the cranial muscles. The axial muscles did not merely contribute but were the primary source of suction expansion power and generated up to 95% of peak expansion power. The recruitment of axial muscle power may have been crucial for the evolution of high-power suction feeding in ray-finned fishes.

  14. 21 CFR 880.5740 - Suction snakebite kit.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ... 21 Food and Drugs 8 2014-04-01 2014-04-01 false Suction snakebite kit. 880.5740 Section 880.5740 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) MEDICAL DEVICES GENERAL HOSPITAL AND PERSONAL USE DEVICES General Hospital and Personal Use Therapeutic Devices § 880.5740 Suction snakebite kit....

  15. 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.

  16. 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.

  17. Compressor airfoil tip clearance optimization system

    DOEpatents

    Little, David A.; Pu, Zhengxiang

    2015-08-18

    A compressor airfoil tip clearance optimization system for reducing a gap between a tip of a compressor airfoil and a radially adjacent component of a turbine engine is disclosed. The turbine engine may include ID and OD flowpath boundaries configured to minimize compressor airfoil tip clearances during turbine engine operation in cooperation with one or more clearance reduction systems that are configured to move the rotor assembly axially to reduce tip clearance. The configurations of the ID and OD flowpath boundaries enhance the effectiveness of the axial movement of the rotor assembly, which includes movement of the ID flowpath boundary. During operation of the turbine engine, the rotor assembly may be moved axially to increase the efficiency of the turbine engine.

  18. Flow-induced noise of a wall-mounted finite airfoil at low-to-moderate Reynolds number

    NASA Astrophysics Data System (ADS)

    Moreau, Danielle J.; Prime, Zebb; Porteous, Ric; Doolan, Con J.; Valeau, Vincent

    2014-12-01

    This paper presents an experimental investigation of the flow-induced noise created by a wall-mounted finite airfoil at low-to-moderate Reynolds number and zero angle of attack. Far-field noise measurements have been taken at a single observer location and with two perpendicular microphone arrays in an anechoic wind tunnel at Reynolds numbers of Rec=9.2×104-1.6×105, based on chord, and for a variety of airfoil aspect ratios (length to chord ratio of L/C=0.2-2, corresponding to length to thickness ratio of L/T=1.7-16.7). Additionally, surface oil-film visualisation images and unsteady velocity measurements taken in the near trailing edge wake are related to far-field noise measurements to determine the flow mechanisms responsible for noise generation. The results show that the wall-mounted finite airfoil radiates noise similar to a two-dimensional airfoil when L/T>8.3. Despite the incoming boundary layer height at the junction being 1.30≤δ/T≤1.46, junction and tip flow suppresses tonal noise production for airfoil's up to L/T=8.3 at Rec=9.2×104-1.2×105. Trailing edge noise is found to be the dominant airfoil noise generation mechanism at frequencies above 1 kHz with the position of the noise source along the trailing edge determined by the proportion of the airfoil span influenced by flow at the airfoil-wall junction.

  19. Transition Flight Experiments on a Swept Wing With Suction

    NASA Technical Reports Server (NTRS)

    Maddalon, D. V.; Collier, F. S., Jr.; Montoya, L. C.; Land, C. K.

    1989-01-01

    Flight experiments were conducted on a 30 degree swept wing with a perforated leading edge by systematically varying the location and amount of suction over a range of Mach number and Reynolds number. Suction was varied chordwise ahead of the front spar from either the front or rear direction by sealing spanwise perforated strips. Transition from laminar to turbulent flow was due to leading edge turbulence contamination or crossflow disturbance growth and/or Tollmien-Schlichting disturbance growth-depending on the test configuration, flight condition, and suction location. A state-of-the-art linear stability theory which accounts for body and streamline curvature and compressibility was used to study the boundary layer stability as suction location and magnitude varied. N-factor correlations with transition location were made for various suction configurations.

  20. Suction socket suspension for below-knee amputees.

    PubMed

    Roberts, R A

    1986-03-01

    In this study the current use of suction suspension for below-knee prostheses is examined by means of two questionnaire surveys. The experience of 56 below-knee (B-K) amputees wearing suction socket prostheses is evaluated comparing suction prostheses with previously worn limbs. A high degree of satisfaction was found, with amputees on the whole reporting improved skin condition, diminished pain, and increased activity levels compared to previous prosthetic history. The experience and opinions of 466 certified prosthetist members of the American Orthotist Prosthetist Association are examined in the second survey, including degree of contact, success, and evaluation of problems in using suction suspension for the B-K amputee. This survey indicated limited contact and familiarity with B-K suction suspension, with only 22% stating they had made this type of prosthesis. Prosthetists cited characteristics of the B-K residual limb as the chief deterrent to a successful fitting.

  1. Interferometric investigations of compressible dynamic stall over a transiently pitching airfoil

    NASA Technical Reports Server (NTRS)

    Chandrasekhara, M. S.; Carr, L. W.; Wilder, M. C.

    1993-01-01

    The dynamic stall flow field over NACA 0012 airfoil pitching transiently from 0 - 60 at a constant rate under compressible flow conditions has been studied using the real-time technique of point diffraction interferometry. This investigation using nonintrusive diagnostics provides a quantitative description of the overall flow field, including the finer details of dynamic stall vortex formation, growth and the concomitant changes in the pressure distribution. Analysis of several hundred interferograms obtained for a range of flow conditions shows that the peak leading edge suction pressure coefficient that stall is nearly constant for a given free stream Mach number at all nondimensional pitch rates. Also, this value is below that seen in steady flow at static stall for the same Mach number, indicating that dynamic effects significantly effect the separation behavior. Further, for a given Mach number, the dynamic stall vortex seems to form rapidly at nearly the same angle of attack for all pitch rates studied. As the vortex is shed, it induces an anti-clockwise trailing edge vortex, which grows in a manner similar to that of a starting vortex. The measured peak suction pressure coefficient drops as the free stream Mach number increases. For free stream Mach numbers above 0.4, small multiple shocks appear near the leading edge.

  2. 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.

  3. Origins, Innovations, and Diversification of Suction Feeding in Vertebrates.

    PubMed

    Wainwright, Peter C; McGee, Matthew D; Longo, Sarah J; Hernandez, L Patricia

    2015-07-01

    We review the origins, prominent innovations, and major patterns of diversification in suction feeding by vertebrates. Non-vertebrate chordates and larval lamprey suspension-feed by capturing small particles in pharyngeal mucous. In most of these lineages the gentle flows that transport particles are generated by buccal cilia, although larval lamprey and thaliacean urochordates have independently evolved a weak buccal pump to generate an oscillating flow of water that is powered by elastic recovery of the pharynx following compression by buccal muscles. The evolution of jaws and the hyoid facilitated powerful buccal expansion and high-performance suction feeding as found today throughout aquatic vertebrates. We highlight three major innovations in suction feeding. Most vertebrate suction feeders have mechanisms that occlude the corners of the open mouth during feeding. This produces a planar opening that is often nearly circular in shape. Both features contribute to efficient flow of water into the mouth and help direct the flow to the area directly in front of the mouth's aperture. Among several functions that have been identified for protrusion of the upper jaw, is an increase in the hydrodynamic forces that suction feeders exert on their prey. Protrusion of the upper jaw has evolved five times in ray-finned fishes, including in two of the most successful teleost radiations, cypriniforms and acanthomorphs, and is found in about 60% of living teleost species. Diversification of the mechanisms of suction feeding and of feeding behavior reveals that suction feeders with high capacity for suction rarely approach their prey rapidly, while slender-bodied predators with low capacity for suction show the full range of attack speeds. We hypothesize that a dominant axis of diversification among suction feeders involves a trade-off between the forces that are exerted on prey and the volume of water that is ingested. PMID:25920508

  4. The concentration distribution around a growing gas bubble in a bio tissue under the effect of suction process.

    PubMed

    Mohammadein, S A

    2014-07-01

    The concentration distribution around a growing nitrogen gas bubble in the blood and other bio tissues of divers who ascend to surface too quickly is obtained by Mohammadein and Mohamed model (2010) for variant and constant ambient pressure through the decompression process. In this paper, the growing of gas bubbles and concentration distribution under the effect of suction process are studied as a modification of Mohammadein and Mohamed model (zero suction). The growth of gas bubble is affected by ascent rate, tissue diffusivity, initial concentration difference, surface tension and void fraction. Mohammadein and Mohamed model (2010) is obtained as a special case from the present model. Results showed that, the suction process activates the systemic blood circulation and delay the growth of gas bubbles in the bio tissues to avoid the incidence of decompression sickness (DCS).

  5. 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.

  6. 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.

  7. Unsteady Thick Airfoil Aerodynamics: Experiments, Computation, and Theory

    NASA Technical Reports Server (NTRS)

    Strangfeld, C.; Rumsey, C. L.; Mueller-Vahl, H.; Greenblatt, D.; Nayeri, C. N.; Paschereit, C. O.

    2015-01-01

    An experimental, computational and theoretical investigation was carried out to study the aerodynamic loads acting on a relatively thick NACA 0018 airfoil when subjected to pitching and surging, individually and synchronously. Both pre-stall and post-stall angles of attack were considered. Experiments were carried out in a dedicated unsteady wind tunnel, with large surge amplitudes, and airfoil loads were estimated by means of unsteady surface mounted pressure measurements. Theoretical predictions were based on Theodorsen's and Isaacs' results as well as on the relatively recent generalizations of van der Wall. Both two- and three-dimensional computations were performed on structured grids employing unsteady Reynolds-averaged Navier-Stokes (URANS). For pure surging at pre-stall angles of attack, the correspondence between experiments and theory was satisfactory; this served as a validation of Isaacs theory. Discrepancies were traced to dynamic trailing-edge separation, even at low angles of attack. Excellent correspondence was found between experiments and theory for airfoil pitching as well as combined pitching and surging; the latter appears to be the first clear validation of van der Wall's theoretical results. Although qualitatively similar to experiment at low angles of attack, two-dimensional URANS computations yielded notable errors in the unsteady load effects of pitching, surging and their synchronous combination. The main reason is believed to be that the URANS equations do not resolve wake vorticity (explicitly modeled in the theory) or the resulting rolled-up un- steady flow structures because high values of eddy viscosity tend to \\smear" the wake. At post-stall angles, three-dimensional computations illustrated the importance of modeling the tunnel side walls.

  8. Technology for pressure-instrumented thin airfoil models, phase 1

    NASA Technical Reports Server (NTRS)

    Wigley, D. A.

    1985-01-01

    A network of channels was chemically milled into one surface of a pair of matched plates having bond planes which were neither planar or profiled to match the contour of the trailing edge of a supercritical airfoil for testing in cryogenic wind tunnels. Vacuum brazing bonded the plates together to create a network of pressure passages without blockages or cross leaks. The greatest success was achieved with the smaller samples and planar bonding surfaces. In larger samples, problems were encountered due to warpage created by the relief of residual stresses. Successful bonds were formed by brazing A286, Nitronic 40 and 300 series stainless steels at 1065 C using AMS 4777B brazing alloy, but excessive grain growth occurred in samples of 200 grade 18 nickel maraging steels. Good bonds were obtained with maraging steel using a 47 percent Nickel-47 percent Palladium-6 percent Silicon alloy and brazing at 927 C. Electro-Discharge-Machining was an effective method of cutting profiled bond planes and airfoil contours. Orifices of good definition were obtained when the EDM wire cut passed through predrilled holes. Possible configurations for joints between small segments and the larger main wing were also studied.

  9. A Separation Control CFD Validation Test Case. Part 1; Baseline and Steady Suction

    NASA Technical Reports Server (NTRS)

    Greenblatt, David; Paschal, Keith B.; Yao, Chung-Sheng; Harris, jerome; Schaeffler, Norman W.; Washburn, Anthony E.

    2004-01-01

    Low speed flow separation over a wall-mounted hump, and its control using steady suction, were studied experimentally in order to generate a data set for a workshop aimed at validating CFD turbulence models. The baseline and controlled data sets comprised static and dynamic surface pressure measurements, flow field measurements using Particle Image Velocimetry (PIV) and wall shear stress obtained via oil-film interferometry. In addition to the specific test cases studied, surface pressures for a wide variety of conditions were reported for different Reynolds numbers and suction rates. Stereoscopic PIV and oil-film flow visualization indicated that the baseline separated flow field was mainly two-dimensional. With the application of control, some three-dimensionality was evident in the spanwise variation of pressure recovery, reattachment location and spanwise pressure fluctuations. Part 2 of this paper, under preparation for the AIAA Meeting in Reno 2005, considers separation control by means of zero-efflux oscillatory blowing.

  10. Numerical simulation of compressible fluid flow in an ultrasonic suction pump.

    PubMed

    Wada, Yuji; Koyama, Daisuke; Nakamura, Kentaro

    2016-08-01

    Characteristics of an ultrasonic suction pump that uses a vibrating piston surface and a pipe are numerically simulated and compared with experimental results. Fluid analysis based on the finite-difference time-domain (FDTD) routine is performed, where the nonlinear term and the moving fluid-surface boundary condition are considered. As a result, the suction mechanism of the pump is found to be similar to that of a check valve, where the gap is open during the inflow phase, and it is nearly closed during the outflow phase. The effects of Reynolds number, vibration amplitude and gap thickness on the pump performance are analyzed. The calculated result is in good agreement with the previously measured results.

  11. Numerical simulation of compressible fluid flow in an ultrasonic suction pump.

    PubMed

    Wada, Yuji; Koyama, Daisuke; Nakamura, Kentaro

    2016-08-01

    Characteristics of an ultrasonic suction pump that uses a vibrating piston surface and a pipe are numerically simulated and compared with experimental results. Fluid analysis based on the finite-difference time-domain (FDTD) routine is performed, where the nonlinear term and the moving fluid-surface boundary condition are considered. As a result, the suction mechanism of the pump is found to be similar to that of a check valve, where the gap is open during the inflow phase, and it is nearly closed during the outflow phase. The effects of Reynolds number, vibration amplitude and gap thickness on the pump performance are analyzed. The calculated result is in good agreement with the previously measured results. PMID:27183101

  12. 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.

  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. Maximizing collection and minimizing risk: does vacuum suction sampling increase the likelihood for misinterpretation of food web connections?

    PubMed

    Chapman, Eric G; Romero, Susan A; Harwood, James D

    2010-11-01

    Molecular tools that characterize the structure of complex food webs and identify trophic connectedness in the field have become widely adopted in recent years. However, characterizing the intensity of predator-prey interactions can be prone to error. Maximizing collection success of small, fast-moving predators with vacuum suction samplers has the potential to increase the likelihood of prey DNA detection either through surface-level contamination with damaged prey or direct consumption within the sampling device. In this study, we used PCR to test the hypothesis that vacuum suction sampling will not cause an erroneous increase in the detection of 'predation', thereby incorrectly assigning trophic linkages when evaluating food web structure. We utilized general (1) Aphidoidea and (2) Collembola primers to measure the predation rates of Glenognatha foxi (Araneae: Tetragnathidae) on these prey collected by hand versus those sampled with a vacuum suction device. With both primer pairs, there was no significant increase in predators screening positive for prey DNA when sampled by vacuum suction versus those predators collected, in parallel, by hand. These results clearly validate the application of vacuum suction sampling during molecular gut-content analysis of predator-prey feeding linkages in the field. Furthermore, we found no evidence that predation was occurring inside the suction sampler because specimens collected were never observed to be feeding nor did they screen positive at greater frequencies than hand-collected individuals. Therefore, it can be concluded that the use of vacuum suction sampling devices (in this case a Modified CDC Backpack Aspirator Model 1412) is suitable for molecular gut-content analysis.

  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. Body ram, not suction, is the primary axis of suction-feeding diversity in spiny-rayed fishes.

    PubMed

    Longo, Sarah J; McGee, Matthew D; Oufiero, Christopher E; Waltzek, Thomas B; Wainwright, Peter C

    2016-01-01

    Suction-feeding fishes exhibit diverse prey-capture strategies that vary in their relative use of suction and predator approach (ram), which is often referred to as the ram-suction continuum. Previous research has found that ram varies more than suction distance among species, such that ram accounts for most differences in prey-capture behaviors. To determine whether these findings hold at broad evolutionary scales, we collected high-speed videos of 40 species of spiny-rayed fishes (Acanthomorpha) feeding on live prey. For each strike, we calculated the contributions of suction, body ram (swimming) and jaw ram (mouth movement relative to the body) to closing the distance between predator and prey. We confirm that the contribution of suction distance is limited even in this phylogenetically and ecologically broad sample of species, with the extreme suction area of prey-capture space conspicuously unoccupied. Instead of a continuum from suction to ram, we find that variation in body ram is the major factor underlying the diversity of prey-capture strategies among suction-feeding fishes. Independent measurement of the contribution of jaw ram revealed that it is an important component of diversity among spiny-rayed fishes, with a number of ecomorphologies relying heavily on jaw ram, including pivot feeding in syngnathiforms, extreme jaw protruders and benthic sit-and-wait ambush predators. A combination of morphological and behavioral innovations has allowed fish to invade the extreme jaw ram area of prey-capture space. We caution that while two-species comparisons may support a ram-suction trade-off, these patterns do not speak to broader patterns across spiny-rayed fishes.

  17. Body ram, not suction, is the primary axis of suction-feeding diversity in spiny-rayed fishes.

    PubMed

    Longo, Sarah J; McGee, Matthew D; Oufiero, Christopher E; Waltzek, Thomas B; Wainwright, Peter C

    2016-01-01

    Suction-feeding fishes exhibit diverse prey-capture strategies that vary in their relative use of suction and predator approach (ram), which is often referred to as the ram-suction continuum. Previous research has found that ram varies more than suction distance among species, such that ram accounts for most differences in prey-capture behaviors. To determine whether these findings hold at broad evolutionary scales, we collected high-speed videos of 40 species of spiny-rayed fishes (Acanthomorpha) feeding on live prey. For each strike, we calculated the contributions of suction, body ram (swimming) and jaw ram (mouth movement relative to the body) to closing the distance between predator and prey. We confirm that the contribution of suction distance is limited even in this phylogenetically and ecologically broad sample of species, with the extreme suction area of prey-capture space conspicuously unoccupied. Instead of a continuum from suction to ram, we find that variation in body ram is the major factor underlying the diversity of prey-capture strategies among suction-feeding fishes. Independent measurement of the contribution of jaw ram revealed that it is an important component of diversity among spiny-rayed fishes, with a number of ecomorphologies relying heavily on jaw ram, including pivot feeding in syngnathiforms, extreme jaw protruders and benthic sit-and-wait ambush predators. A combination of morphological and behavioral innovations has allowed fish to invade the extreme jaw ram area of prey-capture space. We caution that while two-species comparisons may support a ram-suction trade-off, these patterns do not speak to broader patterns across spiny-rayed fishes. PMID:26596534

  18. Formation of Three-Dimensional Stall Cells on Two-Dimensional Airfoils

    NASA Astrophysics Data System (ADS)

    Sivaneri, Victor; Tuna, Burak; Demauro, Edward; Amitay, Michael

    2014-11-01

    Stall cells are a pattern of three-dimensional mushroom-shaped structures that form within the separated region of stalled, thick airfoils within a certain range of Reynolds numbers. The occurrence and number of stall cells are dependent on the wing camber, aspect ratio, angle of attack, and Reynolds number. While much work within the literature has been conducted to visualize and measure this phenomenon, to date a comprehensive explanation for their existence remains elusive. The present work aims to identify these structures, quantify them, and understand the mechanisms by which they are formed. This was conducted using oil flow visualization and stereoscopic particle image velocimetry (SPIV) on a two-dimensional NACA 0015 airfoil, pitched to 18° angle of attack, at Reynolds numbers ranging from 160,000 to 400,000. Oil flow visualization was used to qualitatively identify the signature of the stall cells on the airfoil surface and resolve the associated skin friction vector fields. In addition, SPIV measurements were taken in order to quantify the flow field in the presence and absence of stall cells within the region of separated flow above the surface of the airfoil. Results showed that the stall cells are highly sensitive to Reynolds number, with evidence of an apparent bi-stable state existing at a Reynolds number of 320,000.

  19. 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.

  20. Experimental and computational investigation of lift-enhancing tabs on a multi-element airfoil

    NASA Technical Reports Server (NTRS)

    Ashby, Dale

    1996-01-01

    An experimental and computational investigation of the effect of lift enhancing tabs on a two-element airfoil was conducted. The objective of the study was to develop an understanding of the flow physics associated with lift enhancing tabs on a multi-element airfoil. A NACA 63(sub 2)-215 ModB airfoil with a 30 percent chord Fowler flap was tested in the NASA Ames 7 by 10 foot wind tunnel. Lift enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computer results predict all of the trends in the experimental data quite well. When the flow over the flap upper surface is attached, tabs mounted at the main element trailing edge (cove tabs) produce very little change in lift. At high flap deflections. however, the flow over the flap is separated and cove tabs produce large increases in lift and corresponding reductions in drag by eliminating the separated flow. Cove tabs permit high flap deflection angles to be achieved and reduce the sensitivity of the airfoil lift to the size of the flap gap. Tabs attached to the flap training edge (flap tabs) are effective at increasing lift without significantly increasing drag. A combination of a cove tab and a flap tab increased the airfoil lift coefficient by 11 percent relative to the highest lift tab coefficient achieved by any baseline configuration at an angle of attack of zero percent and the maximum lift coefficient was increased by more than 3 percent. A simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift enhancing tabs work. The tabs were modeled by a point vortex at the training edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift enhancing tabs on a multi-element airfoil. Results of the modeling

  1. How to perform open tracheal suction via an endotracheal tube.

    PubMed

    Credland, Nicola

    2016-04-27

    Rationale and key points Tracheal suction involves the removal of pulmonary secretions from the respiratory tract using negative pressure under sterile conditions. Practitioners should be aware of the indications for, and risks associated with, open tracheal suction via an endotracheal tube. ▶ Respiratory assessment of the patient should be carried out to identify when tracheal suction is required. ▶ A suction pressure of 80-120mmHg is recommended, and suction should last no longer than 15 seconds. ▶ Reassurance and support should be given to the patient to minimise any discomfort and distress that might result from tracheal suction. Reflective activity Clinical skills articles can help update your practice and ensure it remains evidence-based. Apply this article to your practice. Reflect on and write a short account of: 1. How you think this article will change your practice when performing open tracheal suction via an endotracheal tube. 2. How you could use this resource to educate your colleagues. Subscribers can upload their reflective accounts at: rcni.com/portfolio .

  2. 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.

  3. 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.

  4. 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.

  5. Influence of airfoil geometry on delta wing leading-edge vortices and vortex-induced aerodynamics at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.

    1992-01-01

    An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.

  6. Aerodynamic performance and pressure distributions for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Jenkins, Renaldo V.; Hill, Acquilla S.; Ray, Edward J.

    1988-01-01

    This report presents in graphic and tabular forms the aerodynamic coefficient and surface pressure distribution data for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The test was another in a series of tests involved in the joint NASA/U.S. Industry Advanced Technology Airfoil Tests program. This 14% thick supercritical airfoil was tested at Mach numbers from 0.6 to 0.76 and angles of attack from -2.0 to 6.0 degrees. The test Reynolds numbers were 4 million, 6 million, 10 million, 15 million, 30 million, 40 million, and 45 million.

  7. 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.

  8. Dynamic Stall in Pitching Airfoils: Aerodynamic Damping and Compressibility Effects

    NASA Astrophysics Data System (ADS)

    Corke, Thomas C.; Thomas, Flint O.

    2015-01-01

    Dynamic stall is an incredibly rich fluid dynamics problem that manifests itself on an airfoil during rapid, transient motion in which the angle of incidence surpasses the static stall limit. It is an important element of many manmade and natural flyers, including helicopters and supermaneuverable aircraft, and low-Reynolds number flapping-wing birds and insects. The fluid dynamic attributes that accompany dynamic stall include an eruption of vorticity that organizes into a well-defined dynamic stall vortex and massive excursions in aerodynamic loads that can couple with the airfoil structural dynamics. The dynamic stall process is highly sensitive to surface roughness that can influence turbulent transition and to local compressibility effects that occur at free-stream Mach numbers that are otherwise incompressible. Under some conditions, dynamic stall can result in negative aerodynamic damping that leads to limit-cycle growth of structural vibrations and rapid mechanical failure. The mechanisms leading to negative damping have been a principal interest of recent experiments and analysis. Computational fluid dynamic simulations and low-order models have not been good predictors so far. Large-eddy simulation could be a viable approach although it remains computationally intensive. The topic is technologically important owing to the desire to develop next-generation rotorcraft that employ adaptive rotor dynamic stall control.

  9. 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.

  10. Measurement and Formation of Three-Dimensional Stall Cells on Two-Dimensional Airfoils

    NASA Astrophysics Data System (ADS)

    Sivaneri, Victor

    The present work aims to identify, quantify, and understand the physics by which three-dimensional stall cells are formed on two-dimensional airfoils, using oil flow visualization, load cells, and Stereoscopic Particle Image Velocimetry (SPIV). The oil flow visualizations were conducted on two-dimensional NACA 0015 and NACA 0009 airfoils, at angles of attack ranging from 14° to 20°, Reynolds numbers ranging from 1.70x105to 4.20x105, and aspect ratios of 4, 6.67, and 13.33. Load cell data was conducted on a NACA 0015 airfoil, at angles of attack ranging from 0° to 20°, Reynolds numbers ranging from 1.70x105 to 4.20x105, and an aspect ratio of 4. SPIV experiments were conducted on a on a stalled two-dimensional NACA-0015, pitched to 18° angle of attack, at four Reynolds numbers ranging from 1.70x105 to 4.20x105. Oil flow visualizations were used for both qualitatively identify the stall cells and to calculate the near-surface skin friction field. The results showed that, the angle of attack, Reynolds number, aspect ratio, and the airfoil shape and thickness had a pronounced effect on the formation of stall cells. In order to understand the formation of stall cells, the effect of adding a passive disturbance, a zig-zag tape that attached to the airfoil's surface, was explored. The results showed that under some conditions, adding either a two-dimensional or a localized disturbance could alter the shape of the separation and yield the formation of stall cells. In addition, wall-mounted load cells were used to measure the lift and drag on the airfoil, in the presence or absence of stall cells. Finally, the SPIV measurements correlated the flow field over the airfoil with the oil flow visualization on the surface.

  11. Development of Cutting and Suction Device with Twist Blade Screw for Minimally Invasive Surgery: Evaluation of Suction Performance

    PubMed Central

    Fujii, Yusuke; Suzuki, Takashi; Tamura, Manabu; Muragaki, Yoshihiro; Iseki, Hiroshi

    2015-01-01

    In this study, we aim to develop a narrow-diameter and long-bore device for minimally invasive surgery that achieves the simultaneous cutting and suction of body tissue such as the diseased part of an organ. In this paper, we propose a screw made of a thin metal plate, and we developed a prototype device using this screw. For smooth operation, the suction performance must be superior to the cutting performance. Therefore, we performed experiments and evaluated the suction performance of the developed device assuming the crushed tissue pieces correspond to a highly viscous fluid. From the results, we confirmed that the suction volume is almost proportional to the rotation speed of the screw in the low speed range, and the device has an upper limit of suction volume at a certain rotation speed. Considering practical use, its proportional speed range is suitable for the device controllability of cutting and suction volume, and the size of the device tip needs to be 1 mm or more. Based on these conditions, we are planning to examine the shape of the cutting edge for realizing efficient cutting and suction and we will complete the device. PMID:26132592

  12. Development of Cutting and Suction Device with Twist Blade Screw for Minimally Invasive Surgery: Evaluation of Suction Performance.

    PubMed

    Fujii, Yusuke; Suzuki, Takashi; Tamura, Manabu; Muragaki, Yoshihiro; Iseki, Hiroshi

    2015-01-01

    In this study, we aim to develop a narrow-diameter and long-bore device for minimally invasive surgery that achieves the simultaneous cutting and suction of body tissue such as the diseased part of an organ. In this paper, we propose a screw made of a thin metal plate, and we developed a prototype device using this screw. For smooth operation, the suction performance must be superior to the cutting performance. Therefore, we performed experiments and evaluated the suction performance of the developed device assuming the crushed tissue pieces correspond to a highly viscous fluid. From the results, we confirmed that the suction volume is almost proportional to the rotation speed of the screw in the low speed range, and the device has an upper limit of suction volume at a certain rotation speed. Considering practical use, its proportional speed range is suitable for the device controllability of cutting and suction volume, and the size of the device tip needs to be 1 mm or more. Based on these conditions, we are planning to examine the shape of the cutting edge for realizing efficient cutting and suction and we will complete the device. PMID:26132592

  13. Turbulent intensity and Reynolds number effects on an airfoil at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Wang, S.; Zhou, Y.; Alam, Md. Mahbub; Yang, H.

    2014-11-01

    This work investigates the aerodynamics of a NACA 0012 airfoil at the chord-based Reynolds numbers (Rec) from 5.3 × 103 to 2.0 × 104. The lift and drag coefficients, CL and CD, of the airfoil, along with the flow structure, were measured as the turbulent intensity Tu of oncoming flow varies from 0.6% to 6.0%. The analysis of the present data and those in the literature unveils a total of eight distinct flow structures around the suction side of the airfoil. Four Rec regimes, i.e., the ultra-low (<1.0 × 104), low (1.0 × 104-3.0 × 105), moderate (3.0 × 105-5.0 × 106), and high Rec (>5.0 × 106), are proposed based on their characteristics of the CL-Rec relationship and the flow structure. It has been observed that Tu has a more pronounced effect at lower Rec than at higher Rec on the shear layer separation, reattachment, transition, and formation of the separation bubble. As a result, CL, CD, CL/CD and their dependence on the airfoil angle of attack all vary with Tu. So does the critical Reynolds number Rec,cr that divides the ultra-low and low Rec regimes. It is further noted that the effect of increasing Tu bears similarity in many aspects to that of increasing Rec, albeit with differences. The concept of the effective Reynolds number Rec,eff advocated for the moderate and high Rec regimes is re-evaluated for the low and ultra-low Rec regimes. The Rec,eff treats the non-zero Tu effect as an addition of Rec and is determined based on the presently defined Rec,cr. It has been found that all the maximum lift data from both present measurements and previous reports collapse into a single curve in the low and ultra-low Rec regimes if scaled with Rec,eff.

  14. 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.

  15. 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.

  16. Boundary-layer receptivity due to a wall suction and control of Tollmien-Schlichting waves

    NASA Technical Reports Server (NTRS)

    Bodonyi, R. J.; Duck, P. W.

    1990-01-01

    A numerical study of the generation of Tollmien-Schlichting (T-S) waves due to the interaction between a small free-stream disturbance and a small localized suction slot on an otherwise flat surface was carried out using finite difference methods. The nonlinear steady flow is of the viscous-inviscid interactive type while the unsteady disturbed flow is assumed to be governed by the Navier-Stokes equations linearized about this flow. Numerical solutions illustrate the growth or decay of T-S waves generated by the interaction between the free-stream disturbance and the suction slot, depending on the value of the scaled Strouhal number. An important result of this receptivity problem is the numerical determination of the amplitude of the T-S waves and the demonstration of the possible active control of the growth of T-S waves.

  17. Boundary-layer receptivity due to a wall suction and control of Tollmien-Schlichting waves

    NASA Technical Reports Server (NTRS)

    Bodonyi, R. J.; Duck, P. W.

    1992-01-01

    A numerical study of the generation of Tollmien-Schlichting (T-S) waves due to the interaction between a small free-stream disturbance and a small localized suction slot on an otherwise flat surface was carried out using finite difference methods. The nonlinear steady flow is of the viscous-inviscid interactive type while the unsteady disturbed flow is assumed to be governed by the Navier-Stokes equations linearized about this flow. Numerical solutions illustrate the growth or decay of T-S waves generated by the interaction between the free-stream disturbance and the suction slot, depending on the value of the scaled Strouhal number. An important result of this receptivity problem is the numerical determination of the amplitude of the T-S waves and the demonstration of the possible active control of the growth of T-S waves.

  18. Optically transparent multi-suction electrode arrays

    PubMed Central

    Nagarah, John M.; Stowasser, Annette; Parker, Rell L.; Asari, Hiroki; Wagenaar, Daniel A.

    2015-01-01

    Multielectrode arrays (MEAs) allow for acquisition of multisite electrophysiological activity with submillisecond temporal resolution from neural preparations. The signal to noise ratio from such arrays has recently been improved by substrate perforations that allow negative pressure to be applied to the tissue; however, such arrays are not optically transparent, limiting their potential to be combined with optical-based technologies. We present here multi-suction electrode arrays (MSEAs) in quartz that yield a substantial increase in the detected number of units and in signal to noise ratio from mouse cortico-hippocampal slices and mouse retina explants. This enables the visualization of stronger cross correlations between the firing rates of the various sources. Additionally, the MSEA's transparency allows us to record voltage sensitive dye activity from a leech ganglion with single neuron resolution using widefield microscopy simultaneously with the electrode array recordings. The combination of enhanced electrical signals and compatibility with optical-based technologies should make the MSEA a valuable tool for investigating neuronal circuits. PMID:26539078

  19. 21 CFR 878.4780 - Powered suction pump.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ...) MEDICAL DEVICES GENERAL AND PLASTIC SURGERY DEVICES Surgical Devices § 878.4780 Powered suction pump. (a... support system. The device may be used during surgery in the operating room or at the patient's...

  20. 21 CFR 878.4780 - Powered suction pump.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ...) MEDICAL DEVICES GENERAL AND PLASTIC SURGERY DEVICES Surgical Devices § 878.4780 Powered suction pump. (a... support system. The device may be used during surgery in the operating room or at the patient's...

  1. 21 CFR 878.4780 - Powered suction pump.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ...) MEDICAL DEVICES GENERAL AND PLASTIC SURGERY DEVICES Surgical Devices § 878.4780 Powered suction pump. (a... support system. The device may be used during surgery in the operating room or at the patient's...

  2. 21 CFR 878.4780 - Powered suction pump.

    Code of Federal Regulations, 2013 CFR

    2013-04-01

    ...) MEDICAL DEVICES GENERAL AND PLASTIC SURGERY DEVICES Surgical Devices § 878.4780 Powered suction pump. (a... support system. The device may be used during surgery in the operating room or at the patient's...

  3. 21 CFR 878.4780 - Powered suction pump.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ...) MEDICAL DEVICES GENERAL AND PLASTIC SURGERY DEVICES Surgical Devices § 878.4780 Powered suction pump. (a... support system. The device may be used during surgery in the operating room or at the patient's...

  4. A trial of suction drainage in inguinal hernia repair.

    PubMed

    Beacon, J; Hoile, R W; Ellis, H

    1980-08-01

    A prospective randomized trial was conducted on 301 adult males undergoing inguinal herniorrhaphy to assess the value of postoperative suction drainage. Hernias were classified into 'complicated' and 'simple'. In the 'complicated' group suction drainage for 24 h significantly reduced the incidence of wound haematoma, seroma or infection from 48.7 per cent to 17.6 per cent (P < 0.01); there was also a noticeable effect on the postoperative morbidity in the 'simple' hernias, although this just failed to achieve significance (4.5 per cent in the suction group compared with 9.8 per cent in the controls). It is concluded that suction drainage should be employed postoperatively following repair of hernias where dissection may be difficult or where other complicating factors are present.

  5. Effect of wakes from moving upstream rods on boundary layer separation from a high lift airfoil

    NASA Astrophysics Data System (ADS)

    Volino, Ralph J.

    2011-11-01

    Highly loaded airfoils in turbines allow power generation using fewer airfoils. High loading, however, can cause boundary layer separation, resulting in reduced lift and increased aerodynamic loss. Separation is affected by the interaction between rotating blades and stationary vanes. Wakes from upstream vanes periodically impinge on downstream blades, and can reduce separation. The wakes include elevated turbulence, which can induce transition, and a velocity deficit, which results in an impinging flow on the blade surface known as a ``negative jet.'' In the present study, flow through a linear cascade of very high lift airfoils is studied experimentally. Wakes are produced with moving rods which cut through the flow upstream of the airfoils, simulating the effect of upstream vanes. Pressure and velocity fields are documented. Wake spacing and velocity are varied. At low Reynolds numbers without wakes, the boundary layer separates and does not reattach. At high wake passing frequencies separation is largely suppressed. At lower frequencies, ensemble averaged velocity results show intermittent separation and reattachment during the wake passing cycle. Supported by NASA.

  6. Estimation of morphing airfoil shape and aerodynamic load using artificial hair sensors

    NASA Astrophysics Data System (ADS)

    Butler, Nathan S.; Su, Weihua; Thapa Magar, Kaman S.; Reich, Gregory W.

    2016-04-01

    An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape as well as the aerodynamic loads at all times. One approach is to utilize a new type of artificial hair sensors developed at the Air Force Research Laboratory to determine the flow conditions surrounding deformable airfoils. In this work, the hair sensor measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the hair sensor measurements. Such measurements will then be used in an artificial neural network based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Various aerodynamic and geometrical properties approximated from the artificial hair sensor and artificial neural network system will be compared with the results of XFoil in order to validate the approximation approach.

  7. Impulsive Start of a Symmetric Airfoil at High Angle of Attack

    NASA Technical Reports Server (NTRS)

    Katz, Joseph; Yon, Steven; Rogers, Stuart E.

    1996-01-01

    The fluid dynamic phenomena following the impulsive start of a NACA 0015 airfoil were studied by using a time accurate solution of the incompressible laminar Navier-Stokes equations. Angle of attack was set at 10 deg to simulate steady-state poststall conditions at a Reynolds number of 1.2 x 10(exp 4). The calculation revealed that large initial lift values can be obtained, immediately following the impulsive start, when a trapped vortex develops above the airfoil. Before the buildup of this trapped vortex and immediately after the airfoil was set into motion, the fluid is attached to the airfoil's surface and flows around the trailing edge, demonstrating the delay in the buildup of the classical Kutta condition. The transient of this effect is quite short and is followed by an attached How event that leads to the trapped vortex that has a longer duration. The just described initial phenomenon eventually transits into a fully developed separated flow pattern identifiable by an alternating, periodic vortex shedding.

  8. The semi-discrete Galerkin finite element modelling of compressible viscous flow past an airfoil

    NASA Technical Reports Server (NTRS)

    Meade, Andrew J., Jr.

    1992-01-01

    A method is developed to solve the two-dimensional, steady, compressible, turbulent boundary-layer equations and is coupled to an existing Euler solver for attached transonic airfoil analysis problems. The boundary-layer formulation utilizes the semi-discrete Galerkin (SDG) method to model the spatial variable normal to the surface with linear finite elements and the time-like variable with finite differences. A Dorodnitsyn transformed system of equations is used to bound the infinite spatial domain thereby permitting the use of a uniform finite element grid which provides high resolution near the wall and automatically follows boundary-layer growth. The second-order accurate Crank-Nicholson scheme is applied along with a linearization method to take advantage of the parabolic nature of the boundary-layer equations and generate a non-iterative marching routine. The SDG code can be applied to any smoothly-connected airfoil shape without modification and can be coupled to any inviscid flow solver. In this analysis, a direct viscous-inviscid interaction is accomplished between the Euler and boundary-layer codes, through the application of a transpiration velocity boundary condition. Results are presented for compressible turbulent flow past NACA 0012 and RAE 2822 airfoils at various freestream Mach numbers, Reynolds numbers, and angles of attack. All results show good agreement with experiment, and the coupled code proved to be a computationally-efficient and accurate airfoil analysis tool.

  9. 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.

  10. Oscillatory Excitation of Unsteady Compressible Flows over Airfoils at Flight Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Seifert, Avi; Pack, LaTunia G.

    1999-01-01

    An experimental investigation, aimed at delaying flow separation due to the occurrence of a shock-wave-boundary-layer interaction, is reported. The experiment was performed using a NACA 0012 airfoil and a NACA 0015 airfoil at high Reynolds number incompressible and compressible flow conditions. The effects of Mach and Reynolds numbers were identified, using the capabilities of the cryogenic-pressurized facility to maintain one parameter fixed and change the other. Significant Reynolds number effects were identified in the baseline compressible flow conditions even at Reynolds number of 10 and 20 million. The main objectives of the experiment were to study the effects of periodic excitation on airfoil drag-divergence and to alleviate the severe unsteadiness associated with shock-induced separation (known as "buffeting"). Zero-mass-flux oscillatory blowing was introduced through a downstream directed slot located at 10% chord on the upper surface of the NACA 0015 airfoil. The effective frequencies generated 2-4 vortices over the separated region, regardless of the Mach number. Even though the excitation was introduced upstream of the shock-wave, due to experimental limitations, it had pronounced effects downstream of it. Wake deficit (associated with drag) and unsteadiness (associated with buffeting) were significantly reduced. The spectral content of the wake pressure fluctuations indicates of steadier flow throughout the frequency range when excitation was applied. This is especially important at low frequencies which are more likely to interact with the airframe.

  11. Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics

    PubMed Central

    Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

    2015-01-01

    The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty. PMID:27347517

  12. Optimum Duty Cycle of Unsteady Plasma Aerodynamic Actuation for NACA0015 Airfoil Stall Separation Control

    NASA Astrophysics Data System (ADS)

    Sun, Min; Yang, Bo; Peng, Tianxiang; Lei, Mingkai

    2016-06-01

    Unsteady dielectric barrier discharge (DBD) plasma aerodynamic actuation technology is employed to suppress airfoil stall separation and the technical parameters are explored with wind tunnel experiments on an NACA0015 airfoil by measuring the surface pressure distribution of the airfoil. The performance of the DBD aerodynamic actuation for airfoil stall separation suppression is evaluated under DBD voltages from 2000 V to 4000 V and the duty cycles varied in the range of 0.1 to 1.0. It is found that higher lift coefficients and lower threshold voltages are achieved under the unsteady DBD aerodynamic actuation with the duty cycles less than 0.5 as compared to that of the steady plasma actuation at the same free-stream speeds and attack angles, indicating a better flow control performance. By comparing the lift coefficients and the threshold voltages, an optimum duty cycle is determined as 0.25 by which the maximum lift coefficient and the minimum threshold voltage are obtained at the same free-stream speed and attack angle. The non-uniform DBD discharge with stronger discharge in the positive half cycle due to electrons deposition on the dielectric slabs and the suppression of opposite momentum transfer due to the intermittent discharge with cutoff of the negative half cycle are responsible for the observed optimum duty cycle. supported by National Natural Science Foundation of China (No. 21276036), Liaoning Provincial Natural Science Foundation of China (No. 2015020123) and the Fundamental Research Funds for the Central Universities of China (No. 3132015154)

  13. Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

    NASA Technical Reports Server (NTRS)

    Shaw, L. M.; Boldman, D. R.; Buggele, A. E.; Buffum, D. H.

    1985-01-01

    Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

  14. A natural low frequency oscillation in the wake of an airfoil near stalling conditions

    NASA Technical Reports Server (NTRS)

    Zaman, K. B. M. Q.; Mckinzie, D. J.

    1988-01-01

    An unusually low frequency oscillation in the flow over an airfoil was explored experimentally. Wind tunnel measurements were carried out with a two dimensional airfoil model at a chord Reynolds number of 100,000. During deep stall the usual bluff-body shedding occurred at a Strouhal number. But at the onset of stall a low frequency periodic oscillation occurred, the corresponding Strouhal number being an order of magnitude lower. The phenomenon occurred in relatively unclean flow when the freestream turbulence was raised to 0.4 percent, but did not in the cleaner flow with turbulence intensity of 0.1 percent. It could also be produced by certain high frequency acoustic excitation. Details of the flow field are compared between a case of low frequency oscillation at alpha = 15 deg and a case of bluff-body shedding at alpha = 22.5 deg. The origin of the low frequency oscillation traces to the upper surface of the airfoil and is seemingly associated with the periodic formation and breakdown of a large separation bubble. The intense flow fluctuations impart significant unsteady forces to the airfoil but diminish rapidly within a distance of one chord from the trailing edge.

  15. A natural low frequency oscillation in the wake of an airfoil near stalling conditions

    NASA Technical Reports Server (NTRS)

    Zaman, K. B. M. Q.; Mckinzie, D. J.

    1987-01-01

    An unusually low frequency oscillation in the flow over an airfoil was explored experimentally. Wind tunnel measurements were carried out with a two dimensional airfoil model at a chord Reynolds number of 100,000. During deep stall the usual bluff-body shedding occurred at a Strouhal number. But at the onset of stall a low frequency periodic oscillation occurred, the corresponding Strouhal number being an order of magnitude lower. The phenomenon occurred in relatively unclean flow when the freestream turbulence was raised to 0.4 percent, but did not in the cleaner flow with turbulence intensity of 0.1 percent. It could also be produced by certain high frequency acoustic excitation. Details of the flow field are compared between a case of low frequency oscillation at alpha = 15 deg and a case of bluff-body shedding at alpha = 22.5 deg. The origin of the low frequency oscillation traces to the upper surface of the airfoil and is seemingly associated with the periodic formation and breakdown of a large separation bubble. The intense flow fluctuations impart significant unsteady forces to the airfoil but diminish rapidly within a distance of one chord from the trailing edge.

  16. Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics.

    PubMed

    Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

    2015-01-01

    The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty.

  17. Investigation of a bio-inspired lift-enhancing effector on a 2D airfoil.

    PubMed

    Johnston, Joe; Gopalarathnam, Ashok

    2012-09-01

    A flap mounted on the upper surface of an airfoil, called a 'lift-enhancing effector', has been shown in wind tunnel tests to have a similar function to a bird's covert feathers, which rise off the wing's surface in response to separated flows. The effector, fabricated from a thin Mylar sheet, is allowed to rotate freely about its leading edge. The tests were performed in the NCSU subsonic wind tunnel at a chord Reynolds number of 4 × 10(5). The maximum lift coefficient with the effector was the same as that for the clean airfoil, but was maintained over an angle-of-attack range from 12° to almost 20°, resulting in a very gentle stall behavior. To better understand the aerodynamics and to estimate the deployment angle of the free-moving effector, fixed-angle effectors fabricated out of stiff wood were also tested. A progressive increase in the stall angle of attack with increasing effector angle was observed, with diminishing returns beyond the effector angle of 60°. Drag tests on both the free-moving and fixed effectors showed a marked improvement in drag at high angles of attack. Oil flow visualization on the airfoil with and without the fixed-angle effectors proved that the effector causes the separation point to move aft on the airfoil, as compared to the clean airfoil. This is thought to be the main mechanism by which an effector improves both lift and drag. A comparison of the fixed-effector results with those from the free-effector tests shows that the free effector's deployment angle is between 30° and 45°. When operating at and beyond the clean airfoil's stall angle, the free effector automatically deploys to progressively higher angles with increasing angles of attack. This slows down the rapid upstream movement of the separation point and avoids the severe reduction in the lift coefficient and an increase in the drag coefficient that are seen on the clean airfoil at the onset of stall. Thus, the effector postpones the stall by 4-8° and makes the

  18. Swimming muscles power suction feeding in largemouth bass.

    PubMed

    Camp, Ariel L; Roberts, Thomas J; Brainerd, Elizabeth L

    2015-07-14

    Most aquatic vertebrates use suction to capture food, relying on rapid expansion of the mouth cavity to accelerate water and food into the mouth. In ray-finned fishes, mouth expansion is both fast and forceful, and therefore requires considerable power. However, the cranial muscles of these fishes are relatively small and may not be able to produce enough power for suction expansion. The axial swimming muscles of these fishes also attach to the feeding apparatus and have the potential to generate mouth expansion. Because of their large size, these axial muscles could contribute substantial power to suction feeding. To determine whether suction feeding is powered primarily by axial muscles, we measured the power required for suction expansion in largemouth bass and compared it to the power capacities of the axial and cranial muscles. Using X-ray reconstruction of moving morphology (XROMM), we generated 3D animations of the mouth skeleton and created a dynamic digital endocast to measure the rate of mouth volume expansion. This time-resolved expansion rate was combined with intraoral pressure recordings to calculate the instantaneous power required for suction feeding. Peak expansion powers for all but the weakest strikes far exceeded the maximum power capacity of the cranial muscles. The axial muscles did not merely contribute but were the primary source of suction expansion power and generated up to 95% of peak expansion power. The recruitment of axial muscle power may have been crucial for the evolution of high-power suction feeding in ray-finned fishes. PMID:26100863

  19. Colonization of Yankauer suction catheters with pathogenic organisms.

    PubMed

    Brown, Melissa; Willms, David

    2005-10-01

    Oral suction devices may be fomites for nosocomial infections. This study was designed to evaluate the rate of contamination of Yankauer suction catheters. Among the 20 catheters tested, 16 (80%) yielded cultures for pathogens. Seven (35%) were colonized with multiple pathogens. Among the organisms encountered included methicillin-resistant Staphylococcus aureus (MRSA) and vancomycin-resistant Enterococci (VRE). These devices should be handled and stored with appropriate care.

  20. The generation of side force by distributed suction

    NASA Technical Reports Server (NTRS)

    Roberts, Leonard; Hong, John

    1993-01-01

    This report provides an approximate analysis of the generation of side force on a cylinder placed horizontal to the flow direction by the application of distributed suction on the rearward side of the cylinder. Relationships are derived between the side force coefficients and the required suction coefficients necessary to maintain attached flow on one side of the cylinder, thereby inducing circulation around the cylinder and a corresponding side force.

  1. Swimming muscles power suction feeding in largemouth bass

    PubMed Central

    Camp, Ariel L.; Roberts, Thomas J.; Brainerd, Elizabeth L.

    2015-01-01

    Most aquatic vertebrates use suction to capture food, relying on rapid expansion of the mouth cavity to accelerate water and food into the mouth. In ray-finned fishes, mouth expansion is both fast and forceful, and therefore requires considerable power. However, the cranial muscles of these fishes are relatively small and may not be able to produce enough power for suction expansion. The axial swimming muscles of these fishes also attach to the feeding apparatus and have the potential to generate mouth expansion. Because of their large size, these axial muscles could contribute substantial power to suction feeding. To determine whether suction feeding is powered primarily by axial muscles, we measured the power required for suction expansion in largemouth bass and compared it to the power capacities of the axial and cranial muscles. Using X-ray reconstruction of moving morphology (XROMM), we generated 3D animations of the mouth skeleton and created a dynamic digital endocast to measure the rate of mouth volume expansion. This time-resolved expansion rate was combined with intraoral pressure recordings to calculate the instantaneous power required for suction feeding. Peak expansion powers for all but the weakest strikes far exceeded the maximum power capacity of the cranial muscles. The axial muscles did not merely contribute but were the primary source of suction expansion power and generated up to 95% of peak expansion power. The recruitment of axial muscle power may have been crucial for the evolution of high-power suction feeding in ray-finned fishes. PMID:26100863

  2. Displacement of oropharyngeal structures during suction-swallowing cycles.

    PubMed

    Engelke, W; Glombek, J; Psychogios, M; Schneider, S; Ellenberger, D; Santander, P

    2014-07-01

    Suction ability plays an important role in supporting oral nutrition and needs special care following neurological disorders and tumor-associated defects. However, the details of suction are still poorly understood. The present study evaluates displacement of orofacial structures during suction and deglutition based on manometric controlled MRI. Nine healthy subjects were scanned wearing an intraoral mouthpiece for water intake by suction and subsequent swallowing. Suction-swallowing cycles were identified by intraoral negative pressure. Midsagittal MRI slices (3 T; temporal resolution 0.53 s) were analyzed at rest, suction and pharyngeal swallowing. The mandibular displacement was measured as the distance between the anterior nasal spine and the inferior point of the mandible. Following areas were defined: subpalatal compartment (SCA), retrolingual (RLA), epipharyngeal (EPA) and mouth floor area (MFA). During rest, an average distance of 7 cm was observed between the mandibular measurement points. The measured SCA was 3.67 cm(2), the RLA 6.98 cm(2), the EPA 9.00 cm(2) and the MFA 15.21 cm(2) (average values). At the end of suction, the mandibular distance reduces (to 6.88 cm), the SCA increases significantly (to 5.96 cm(2); p = 0.0002), the RLA decreases (to 6.45 cm(2)), the EPA increases (to 10.59 cm(2)) and the MFA decreases (to 15.02 cm(2)). During deglutition, the mandible lifted significantly (to 6.81 cm; p = 0.0276), the SCA reduced to zero, the RLA was not measurable, the EPA reduces significantly (to 3.01 cm(2); p < 0.0001) and the MFA increases (to 16.36 cm(2)). According to these observations, a combined displacement of the tongue in an anteroposterior direction with active tongue dorsum-velum contact appears to be the predominant activity during suction and responsible for the expansion of the subpalatal area.

  3. The generation of side force by distributed suction

    NASA Astrophysics Data System (ADS)

    Roberts, Leonard; Hong, John

    1993-05-01

    This report provides an approximate analysis of the generation of side force on a cylinder placed horizontal to the flow direction by the application of distributed suction on the rearward side of the cylinder. Relationships are derived between the side force coefficients and the required suction coefficients necessary to maintain attached flow on one side of the cylinder, thereby inducing circulation around the cylinder and a corresponding side force.

  4. Differing ERP patterns caused by suction and puff stimuli.

    PubMed

    Choi, Mi-Hyun; Kim, Hyung-Sik; Baek, Ji-Hye; Lee, Jung-Chul; Park, Sung-Jun; Jeong, Ul-Ho; Gim, Seon-Young; You, Ji Hye; Kim, Sung-Pil; Lim, Dae-Woon; Kim, Hyun-Jun; Chung, Soon-Cheol

    2015-05-01

    The present study compared event-related potential (ERP) patterns for two stimuli types, puff and suction, by applying these stimuli to the fingers; ERP patterns for the two stimuli were compared at C3, an area related to somatosensory perception, and at FC5, an area related to motor function. Participants were 12 healthy males in their 20s (mean age=23.1±2.0 years). One session consisted of a Control Phase (3s), a Stimulation Phase (3s), and a Rest Phase (9s). During the Stimulation Phase, a 4-psi suction or puff stimulus was applied to the first joint of the right index finger. After completion of the session, a subjective magnitude test was presented. In all phases, electroencephalography signals were recorded. We extracted maximum positive amplitude and minimum negative amplitude as well as relevant latency values for C3 and FC5 signals. Suction and puff stimuli had similar subjective magnitude scores. For both C3 and FC5, the maximum and minimum amplitude latency was reached earlier for the suction stimulus than for the puff stimulus. In conclusion, when suction and puff stimuli of the same intensity were applied to the fingers, the suction stimulus caused a more sensitive response in the somatosensory area (C3) and motor area (FC5) than did the puff stimulus.

  5. Endotracheal suctioning in intubated newborns: an integrative literature review.

    PubMed

    Gonçalves, Roberta Lins; Tsuzuki, Lucila Midori; Carvalho, Marcos Giovanni Santos

    2015-01-01

    Evidence-based practices search for the best available scientific evidence to support problem solving and decision making. Because of the complexity and amount of information related to health care, the results of methodologically sound scientific papers must be integrated by performing literature reviews. Although endotracheal suctioning is the most frequently performed invasive procedure in intubated newborns in neonatal intensive care units, few Brazilian studies of good methodological quality have examined this practice, and a national consensus or standardization of this technique is lacking. Therefore, the purpose of this study was to review secondary studies on the subject to establish recommendations for endotracheal suctioning in intubated newborns and promote the adoption of best-practice concepts when conducting this procedure. An integrative literature review was performed, and the recommendations of this study are to only perform endotracheal suctioning in newborns when there are signs of tracheal secretions and to avoid routinely performing the procedure. In addition, endotracheal suctioning should be conducted by at least two people, the suctioning time should be less than 15 seconds, the negative suction pressure should be below 100 mmHg, and hyperoxygenation should not be used on a routine basis. If indicated, oxygenation is recommended with an inspired oxygen fraction value that is 10 to 20% greater than the value of the previous fraction, and it should be performed 30 to 60 seconds before, during and 1 minute after the procedure. Saline instillation should not be performed routinely, and the standards for invasive procedures must be respected.

  6. Suction blister fluid as potential body fluid for biomarker proteins.

    PubMed

    Kool, Jeroen; Reubsaet, Léon; Wesseldijk, Feikje; Maravilha, Raquel T; Pinkse, Martijn W; D'Santos, Clive S; van Hilten, Jacobus J; Zijlstra, Freek J; Heck, Albert J R

    2007-10-01

    Early diagnosis is important for effective disease management. Measurement of biomarkers present at the local level of the skin could be advantageous in facilitating the diagnostic process. The analysis of the proteome of suction blister fluid, representative for the interstitial fluid of the skin, is therefore a desirable first step in the search for potential biomarkers involved in biological pathways of particular diseases. Here, we describe a global analysis of the suction blister fluid proteome as potential body fluid for biomarker proteins. The suction blister fluid proteome was compared with a serum proteome analyzed using identical protocols. By using stringent criteria allowing less than 1% false positive identifications, we were able to detect, using identical experimental conditions and amount of starting material, 401 proteins in suction blister fluid and 240 proteins in serum. As a major result of our analysis we construct a prejudiced list of 34 proteins, relatively highly and uniquely detected in suction blister fluid as compared to serum, with established and putative characteristics as biomarkers. We conclude that suction blister fluid might potentially serve as a good alternative biomarker body fluid for diseases that involve the skin.

  7. Endotracheal suctioning in intubated newborns: an integrative literature review

    PubMed Central

    Gonçalves, Roberta Lins; Tsuzuki, Lucila Midori; Carvalho, Marcos Giovanni Santos

    2015-01-01

    Evidence-based practices search for the best available scientific evidence to support problem solving and decision making. Because of the complexity and amount of information related to health care, the results of methodologically sound scientific papers must be integrated by performing literature reviews. Although endotracheal suctioning is the most frequently performed invasive procedure in intubated newborns in neonatal intensive care units, few Brazilian studies of good methodological quality have examined this practice, and a national consensus or standardization of this technique is lacking. Therefore, the purpose of this study was to review secondary studies on the subject to establish recommendations for endotracheal suctioning in intubated newborns and promote the adoption of best-practice concepts when conducting this procedure. An integrative literature review was performed, and the recommendations of this study are to only perform endotracheal suctioning in newborns when there are signs of tracheal secretions and to avoid routinely performing the procedure. In addition, endotracheal suctioning should be conducted by at least two people, the suctioning time should be less than 15 seconds, the negative suction pressure should be below 100 mmHg, and hyperoxygenation should not be used on a routine basis. If indicated, oxygenation is recommended with an inspired oxygen fraction value that is 10 to 20% greater than the value of the previous fraction, and it should be performed 30 to 60 seconds before, during and 1 minute after the procedure. Saline instillation should not be performed routinely, and the standards for invasive procedures must be respected. PMID:26465249

  8. 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.

  9. 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.

  10. 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.

  11. Oronasopharyngeal suction versus no suction in normal, term and vaginally born infants: a prospective randomised controlled trial.

    PubMed

    Gungor, Sadettin; Teksoz, Ertan; Ceyhan, Temel; Kurt, Ercan; Goktolga, Umit; Baser, Iskender

    2005-10-01

    This prospective randomised controlled trial aimed to compare the effects of oronasopharyngeal suction with those of no suction in normal, term and vaginally born infants and was performed at a Turkish tertiary hospital from June 2003 to January 2004. A total of 140 newborns were enrolled in the trial (n = 70 per group). The no suction group showed lower mean heart rates through the 3rd and 6th minutes and higher SaO(2) values through the first 6 mins of life (P < 0.001). The maximum time to reach SaO2 of >or= 92% (6 vs. 11 min) and >or= 86% (5 vs. 8 min) were shorter in the no suction group (P < 0.001).

  12. Vapor deposition on doublet airfoil substrates: Control of coating thickness and microstructure

    SciTech Connect

    Rodgers, Theron M.; Zhao, Hengbei; Wadley, Haydn N. G.

    2015-11-15

    Gas jet assisted vapor deposition processes for depositing coatings are conducted at higher pressures than conventional physical vapor deposition methods, and have shown promise for coating complex shaped substrates including those with non-line-of-sight (NLS) regions on their surface. These regions typically receive vapor atoms at a lower rate and with a wider incident angular distribution than substrate regions in line-of-sight (LS) of the vapor source. To investigate the coating of such substrates, the thickness and microstructure variation along the inner (curved) surfaces of a model doublet airfoil containing both LS and NLS regions has been investigated. Results from atomistic simulations and experiments confirm that the coating's thickness is thinner in flux-shadowed regions than in other regions for all the coating processes investigated. They also indicated that the coatings columnar microstructure and pore volume fraction vary with surface location through the LS to NLS transition zone. A substrate rotation strategy for optimizing the thickness over the entire doublet airfoil surface was investigated, and led to the identification of a process that resulted in only small variation of coating thickness, columnar growth angle, and pore volume fraction on all doublet airfoil surfaces.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise

    NASA Technical Reports Server (NTRS)

    2010-01-01

    Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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)

  10. Estimation of morphing airfoil shapes and aerodynamic loads using artificial hair sensors

    NASA Astrophysics Data System (ADS)

    Butler, Nathan Scott

    An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape, as well as the aerodynamic loads, in real time. One approach is to utilize a new type of artificial hair sensors (AHS) developed at the Air Force Research Laboratory (AFRL) to determine the flow conditions surrounding deformable airfoils. In this study, AHS measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the AHS measurements. Such measurements will then be used in an artificial neural network (ANN) based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Additionally, an aerodynamic formulation based on the finite-state inflow theory has been developed to calculate the aerodynamic loads on thin airfoils with arbitrary camber deformations. Various aerodynamic properties approximated from the AHS/ANN system will be compared with the results of the finite-state inflow aerodynamic formulation in order to validate the approximation approach.

  11. 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.

  12. The decay of longitudinal vortices shed from airfoil vortex generators

    NASA Technical Reports Server (NTRS)

    Wendt, Bruce J.; Reichert, Bruce A.; Foster, Jeffry D.

    1995-01-01

    An experimental study is conducted to examine the crossplane structure and streamwise decay of vortices shed from airfoil-type vortex generators. The vortex generators are set in a counter-rotating array spanning the full circumference of a straight pipe. The span of the vortex generators above the duct surface, h, is approximately equal to the local turbulent boundary layer thickness, delta. Measurement of three-component mean flow velocity in downstream crossplanes are used to characterize the structure of the shed vortices. Measurements in adjacent crossplanes (closely spaced along the streamwise coordinate) characterize the interaction and decay of the embedded vortices. A model constructed by the superposition of Oseen vortices is compared to the data for one test case.

  13. Viscous effect on airfoils for unsteady transonic flows

    NASA Technical Reports Server (NTRS)

    Lee, S. C.

    1982-01-01

    The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.

  14. A turbulence model for iced airfoils and its validation

    NASA Technical Reports Server (NTRS)

    Shin, Jaiwon; Chen, Hsun H.; Cebeci, Tuncer

    1992-01-01

    A turbulence model based on the extension of the algebraic eddy viscosity formulation of Cebeci and Smith developed for two dimensional flows over smooth and rough surfaces is described for iced airfoils and validated for computed ice shapes obtained for a range of total temperatures varying from 28 to -15 F. The validation is made with an interactive boundary layer method which uses a panel method to compute the inviscid flow and an inverse finite difference boundary layer method to compute the viscous flow. The interaction between inviscid and viscous flows is established by the use of the Hilbert integral. The calculated drag coefficients compare well with recent experimental data taken at the NASA-Lewis Icing Research Tunnel (IRT) and show that, in general, the drag increase due to ice accretion can be predicted well and efficiently.

  15. Suction-generated noise in an anatomic silicon ear model.

    PubMed

    Luxenberger, Wolfgang; Lahousen, T; Walch, C

    2012-10-01

    The objectives of this study were to evaluate noise levels generated during micro-suction aural toilet using an anatomic silicon ear model. It is an experimental study. In an anatomic ear model made of silicone, the eardrum was replaced by a 1-cm diameter microphone of a calibrated sound-level measuring device. Ear wax was removed using the sucker of a standard ENT treatment unit (Atmos Servant 5(®)). Mean and peak sound levels during the suction procedure were recorded with suckers of various diameters (Fergusson-Frazier 2.7-4 mm as well as Rosen 1.4-2.5 mm). Average noise levels during normal suction in a distance of 1 cm in front of the eardrum ranged between 97 and 103.5 dB(A) (broadband noise). Peak noise levels reached 118 dB(A). During partial obstruction of the sucker by cerumen or dermal flakes, peak noise levels reached 146 dB(A). Peak noise levels observed during the so-called clarinet phenomena were independent of the diameter or type of suckers used. Although micro-suction aural toilet is regarded as an established, widespread and usually safe method to clean the external auditory canal, some caution seems advisable. The performance of long-lasting suction periods straight in front of the eardrum without sound-protecting earwax between sucker and eardrum should be avoided. In particular, when clarinet phenomena are occurring (as described above), the suction procedure should be aborted immediately. In the presence of dermal flakes blocking the auditory canal, cleaning with micro-forceps or other non-suctioning instruments might represent a reasonable alternative.

  16. Nebulous Art of Using Wind-Tunnel Airfoil Data for Predicting Rotor Performance: Preprint

    SciTech Connect

    Tangler, J. L.

    2002-01-01

    The objective of this study was threefold: to evaluate different two-dimensional S809 airfoil data sets in the prediction of rotor performance; to compare blade-element momentum rotor predicted results to lifting-surface, prescribed-wake results; and to compare the NASA Ames combined experiment rotor measured data with the two different performance prediction methods. The S809 airfoil data sets evaluated included those from Delft University of Technology, Ohio State University, and Colorado State University. The performance prediction comparison with NASA Ames data documents shortcomings of these performance prediction methods and recommends the use of the lifting-surface, prescribed-wake method over blade-element momentum theory for future analytical improvements.

  17. 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.

  18. Impingement of Cloud Droplets on 36.5-Percent-Thick Joukowski Airfoil at Zero Angle of Attack and Discussion of Use as Cloud Measuring Instrument in Dye-Tracer Technique

    NASA Technical Reports Server (NTRS)

    Brun, R. J.; Vogt, Dorothea E.

    1957-01-01

    The trajectories of droplets i n the air flowing past a 36.5-percent-thick Joukowski airfoil at zero angle of attack were determined. The amount of water i n 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 cover a large range of flight and atmospheric conditions. With the detailed impingement information available, the 36.5-percent-thick Joukowski airfoil can serve the dual purpose of use as the principal element in instruments for making measurements in clouds and of a basic shape for estimating impingement on a thick streamlined body. Methods and examples are presented for illustrating some limitations when the airfoil is used as the principal element in the dye-tracer technique.

  19. A method for predicting shock shapes and pressure distributions on two dimensional airfoils at large angles of attack

    NASA Technical Reports Server (NTRS)

    Kaattari, G. E.

    1973-01-01

    A method is presented for determining shock envelopes and pressure distributions for two-dimensional airfoils at angles of attack sufficiently large to cause shock detachment and subsonic flow over the windward surface of the airfoil. Correlation functions obtained from exact solutions are used to relate the shock standoff distance at the stagnation and sonic points of the body through a suitable choice for the shock shape. The necessary correlation functions were obtained from perfect gas solutions but may be extended to any gas flow for which the normal shock-density ratio can be specified.

  20. Design and construction of 2 transonic airfoil models for tests in the NASA Langley C.3-M TCT

    NASA Technical Reports Server (NTRS)

    Schaechterle, G.; Ludewig, K. H.; Stanewsky, E.; Ray, E. J.

    1982-01-01

    As part of a NASA/DFVLR cooperation program two transonic airfoils were tested in the NASA Langley 0.3-m TCT. Model design and construction was carried out by DFVLR. The models designed and constructed performed extremely well under cryogenic conditions. Essentially no permanent changes in surface quality and geometric dimensions occurred during the tests. The aerodynamic results from the TCT tests which demonstrate the large sensitivity of the airfoil CAST 10-Z/DOAZ to Reynolds number changes compared well with results from other facilities at ambient temperatures.

  1. The application of the gradient-based adjoint multi-point optimization of single and double shock control bumps for transonic airfoils

    NASA Astrophysics Data System (ADS)

    Mazaheri, K.; Nejati, A.; Chaharlang Kiani, K.; Taheri, R.

    2016-07-01

    A shock control bump (SCB) is a flow control method that uses local small deformations in a flexible wing surface to considerably reduce the strength of shock waves and the resulting wave drag in transonic flows. Most of the reported research is devoted to optimization in a single flow condition. Here, we have used a multi-point adjoint optimization scheme to optimize shape and location of the SCB. Practically, this introduces transonic airfoils equipped with the SCB that are simultaneously optimized for different off-design transonic flight conditions. Here, we use this optimization algorithm to enhance and optimize the performance of SCBs in two benchmark airfoils, i.e., RAE-2822 and NACA-64-A010, over a wide range of off-design Mach numbers. All results are compared with the usual single-point optimization. We use numerical simulation of the turbulent viscous flow and a gradient-based adjoint algorithm to find the optimum location and shape of the SCB. We show that the application of SCBs may increase the aerodynamic performance of an RAE-2822 airfoil by 21.9 and by 22.8 % for a NACA-64-A010 airfoil compared to the no-bump design in a particular flight condition. We have also investigated the simultaneous usage of two bumps for the upper and the lower surfaces of the airfoil. This has resulted in a 26.1 % improvement for the RAE-2822 compared to the clean airfoil in one flight condition.

  2. An Improved Version of the NASA-Lockheed Multielement Airfoil Analysis Computer Program

    NASA Technical Reports Server (NTRS)

    Brune, G. W.; Manke, J. W.

    1978-01-01

    An improved version of the NASA-Lockheed computer program for the analysis of multielement airfoils is described. The predictions of the program are evaluated by comparison with recent experimental high lift data including lift, pitching moment, profile drag, and detailed distributions of surface pressures and boundary layer parameters. The results of the evaluation show that the contract objectives of improving program reliability and accuracy have been met.

  3. Dielectric elastomer actuators for octopus inspired suction cups.

    PubMed

    Follador, M; Tramacere, F; Mazzolai, B

    2014-09-25

    Suction cups are often found in nature as attachment strategy in water. Nevertheless, the application of the artificial counterpart is limited by the dimension of the actuators and their usability in wet conditions. A novel design for the development of a suction cup inspired by octopus suckers is presented. The main focus of this research was on the modelling and characterization of the actuation unit, and a first prototype of the suction cup was realized as a proof of concept. The actuation of the suction cup is based on dielectric elastomer actuators. The presented device works in a wet environment, has an integrated actuation system, and is soft. The dimensions of the artificial suction cups are comparable to proximal octopus suckers, and the attachment mechanism is similar to the biological counterpart. The design approach proposed for the actuator allows the definition of the parameters for its development and for obtaining a desired pressure in water. The fabricated actuator is able to produce up to 6 kPa of pressure in water, reaching the maximum pressure in less than 300 ms.

  4. Dielectric elastomer actuators for octopus inspired suction cups.

    PubMed

    Follador, M; Tramacere, F; Mazzolai, B

    2014-01-01

    Suction cups are often found in nature as attachment strategy in water. Nevertheless, the application of the artificial counterpart is limited by the dimension of the actuators and their usability in wet conditions. A novel design for the development of a suction cup inspired by octopus suckers is presented. The main focus of this research was on the modelling and characterization of the actuation unit, and a first prototype of the suction cup was realized as a proof of concept. The actuation of the suction cup is based on dielectric elastomer actuators. The presented device works in a wet environment, has an integrated actuation system, and is soft. The dimensions of the artificial suction cups are comparable to proximal octopus suckers, and the attachment mechanism is similar to the biological counterpart. The design approach proposed for the actuator allows the definition of the parameters for its development and for obtaining a desired pressure in water. The fabricated actuator is able to produce up to 6 kPa of pressure in water, reaching the maximum pressure in less than 300 ms. PMID:25253019

  5. Rotary blood pump control strategy for preventing left ventricular suction.

    PubMed

    Wang, Yu; Koenig, Steven C; Slaughter, Mark S; Giridharan, Guruprasad A

    2015-01-01

    The risk for left ventricular (LV) suction while maintaining adequate perfusion over a range of physiologic conditions during continuous flow LV assist device (LVAD) support is a significant clinical concern. To address this challenge, we developed a suction prevention and physiologic control (SPPC) algorithm for use with axial and centrifugal LVADs. The SPPC algorithm uses two gain-scheduled, proportional-integral controllers that maintain a differential pump speed (ΔRPM) above a user-defined threshold to prevent LV suction, while maintaining an average reference differential pressure (ΔP) between the LV and aorta to provide physiologic perfusion. Efficacy and robustness of the proposed algorithm were evaluated in silico during simulated rest and exercise test conditions for (1) ΔP/ΔRPM excessive setpoint (ES); (2) rapid eightfold increase in pulmonary vascular resistance (PVR); and (3) ES and PVR. Hemodynamic waveforms (LV pressure and volume; aortic pressure and flow) were simulated and analyzed to identify suction event(s), quantify total flow output (pump + cardiac output), and characterize the performance of the SPPC algorithm. The results demonstrated that the proposed SPPC algorithm prevented LV suction while maintaining physiologic perfusion for all simulated test conditions, and warrants further investigation in vivo. PMID:25248043

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. Adaption of suction connectors for use in meconium aspiration syndrome.

    PubMed

    Green, David Anthony

    2010-01-01

    Severe meconium aspiration syndrome is difficult to manage and has a high mortality in developing countries. Guidelines are available for the initial management. If the infant has been born through particulate meconium and is not vigorous, an inspection of the vocal cords by laryngoscopy is recommended. If meconium is seen at the cords it should, ideally, be sucked out of the trachea using an endotracheal tube as a suction device. However, as this needs a way of applying suction directly to the endotracheal tube it can be problematic. Commercially available equipment does exist, but in a resource-scarce setting, its cost could be prohibitive. We have adapted cheap suction connectors which can be adapted for this purpose.

  11. Safety System for Controlling Fluid Flow into a Suction Line

    NASA Technical Reports Server (NTRS)

    England, John Dwight (Inventor); Kelley, Anthony R. (Inventor); Cronise, Raymond J. (Inventor)

    2015-01-01

    A safety system includes a sleeve fitted within a pool's suction line at the inlet thereof. An open end of the sleeve is approximately aligned with the suction line's inlet. The sleeve terminates with a plate that resides within the suction line. The plate has holes formed therethrough. A housing defining a plurality of distinct channels is fitted in the sleeve so that the distinct channels lie within the sleeve. Each of the distinct channels has a first opening on one end thereof and a second opening on another end thereof. The second openings reside in the sleeve. Each of the distinct channels is at least approximately three feet in length. The first openings are in fluid communication with the water in the pool, and are distributed around a periphery of an area of the housing that prevents coverage of all the first openings when a human interacts therewith.

  12. 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.

  13. 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.

  14. Cohesive suction-cup force in cell separation dynamics

    NASA Astrophysics Data System (ADS)

    Vasseur, H.

    2010-07-01

    When an external pulling force is applied onto a cell stuck to its substrate, a reacting "suction-cup" force, due to the slow penetration of the surrounding fluid between the cell and the substrate, opposes to the separation. It can overcome other known adhesive forces when the process is sufficiently violent (typically 105 N/m2). The physical origin of this effect may be compared with that leaning a suction-cup against a bathroom wall. We address the consequences of this effect on i) the separation energy, ii) the fluid motion surrounding the cell, and iii) the inhibition of cell motion.

  15. The flow over a thin airfoil subjected to elevated levels of freestream turbulence at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Ravi, Sridhar; Watkins, Simon; Watmuff, Jon; Massey, Kevin; Petersen, Phred; Marino, Matthew; Ravi, Anuradha

    2012-09-01

    Micro Air Vehicles (MAVs) can be difficult to control in the outdoor environment as they fly at relatively low speeds and are of low mass, yet exposed to high levels of freestream turbulence present within the Atmospheric Boundary Layer. In order to examine transient flow phenomena, two turbulence conditions of nominally the same longitudinal integral length scale (Lxx/c = 1) but with significantly different intensities (Ti = 7.2 % and 12.3 %) were generated within a wind tunnel; time-varying surface pressure measurements, smoke flow visualization, and wake velocity measurements were made on a thin flat plate airfoil. Rapid changes in oncoming flow pitch angle resulted in the shear layer to separate from the leading edge of the airfoil even at lower geometric angles of attack. At higher geometric angles of attack, massive flow separation occurred at the leading edge followed by enhanced roll up of the shear layer. This lead to the formation of large Leading Edge Vortices (LEVs) that advected at a rate much lower than the mean flow speed while imparting high pressure fluctuations over the airfoil. The rate of LEV formation was dependent on the angle of attack until 10° and it was independent of the turbulence properties tested. The fluctuations in surface pressures and consequently aerodynamic loads were considerably limited on the airfoil bottom surface due to the favorable pressure gradient.

  16. 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.

  17. Navier-Stokes calculations on multi-element airfoils using a chimera-based solver

    NASA Technical Reports Server (NTRS)

    Jasper, Donald W.; Agrawal, Shreekant; Robinson, Brian A.

    1993-01-01

    A study of Navier-Stokes calculations of flows about multielement airfoils using a chimera grid approach is presented. The chimera approach utilizes structured, overlapped grids which allow great flexibility of grid arrangement and simplifies grid generation. Calculations are made for two-, three-, and four-element airfoils, and modeling of the effect of gap distance between elements is demonstrated for a two element case. Solutions are obtained using the thin-layer form of the Reynolds averaged Navier-Stokes equations with turbulence closure provided by the Baldwin-Lomax algebraic model or the Baldwin-Barth one equation model. The Baldwin-Barth turbulence model is shown to provide better agreement with experimental data and to dramatically improve convergence rates for some cases. Recently developed, improved farfield boundary conditions are incorporated into the solver for greater efficiency. Computed results show good comparison with experimental data which include aerodynamic forces, surface pressures, and boundary layer velocity profiles.

  18. Study of the Aero-Acoustic and Aerodynamic Effects of Soft Coating upon Airfoil

    NASA Astrophysics Data System (ADS)

    Vad, János; Koscsó, Gábor; Gutermuth, Miklós; Kasza, Zsolt; Tábi, Tamás; Csörgo, Tibor

    Comparative acoustic and wind tunnel experiments were carried out on uncoated and coated isolated airfoils. The aim of the tests was to survey the airfoil noise reducing effect and the aerodynamic impact of the acoustically soft coating consisting of filaments, as a preliminary study in application of such coatings to axial flow turbomachinery bladings. It was found in the acoustic tests that the coating successfully reduces the sound pressure in the frequency range critical from the aspect of human audition. The wind tunnel experiments included laser Doppler anemometer studies on the development of the boundary layers and on the wake structure, and static pressure measurements on the blade surface and in the wake. The coating reduced the lift and increased the drag. A proposal has been made for further studies in order to retain the advantageous acoustic effects of the coating while avoiding the undesirable aerodynamic impact.

  19. A study of pitch oscillation and roughness on airfoils used for horizontal axis wind turbines

    SciTech Connect

    Gregorek, G.M.; Hoffmann, M.J.; Ramsay, R.R.; Janiszewska, J.M.

    1995-12-01

    Under subcontract XF-1-11009-3 the Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) with the National Renewable Energy Laboratory (NREL) developed an extensive database of empirical aerodynamic data. These data will assist in the development of analytical models and in the design of new airfoils for wind turbines. To accomplish the main objective, airfoil models were designed, built and wind tunnel tested with and without model leading edge grit roughness (LEGR). LEGR simulates surface irregularities due to the accumulation of insect debris, ice, and/or the aging process. This report is a summary of project project activity for Phase III, which encompasses the time period from September 17, 1 993 to September 6, 1 994.

  20. LDA measurement of the passage flow field in a 3-D airfoil cascade

    NASA Technical Reports Server (NTRS)

    Stauter, R. C.; Fleeter, S.

    1986-01-01

    Three-dimensional internal flow computational models are currently being developed to predict the flow through turbomachinery blade rows. For these codes to be of quantitative value, they must be verified with data obtained in experiments which model the fundamental flow phenomena. In this paper, the complete three-dimensional flow field through a subsonic annular cascade of cambered airfoils is experimentally quantified. In particular, detailed three-dimensional data are obtained to quantify the inlet velocity profile, the cascade passage velocity field, and the exit region flow field. The primary instrumentation for acquiring these data is a single-channel Laser Doppler Anemometer operating in the backscatter mode, with chordwise distributions of airfoil surface static pressure taps also utilized. Appropriate data are correlated with predictions from the MERIDL/TSONIC codes.