Science.gov

Sample records for airfoil chord length

  1. Exact Chord-Length Distribution For SEU Calculations

    NASA Technical Reports Server (NTRS)

    Buehler, Martin G.; Luke, Keung L.

    1990-01-01

    Computed rates of SEU's more accurate. Exact integral chord-length distribution derived for use in calculations of rates of single-event upsets (SEU's) (changes in logic states) caused by impingement of cosmic rays or other ionizing radiation on electronic logic circuits.

  2. Techniques for ensuring the correct length of new mitral chords.

    PubMed

    Duran, Carlos M G; Pekar, Filip

    2003-03-01

    Repair of degenerative mitral insufficiency has extensively been shown to be superior to replacement. In the majority of cases, the culprit lesion is limited to the posterior mitral leaflet (PML), which is treated with quadrangular resection of the prolapsing PML, annular plication of the corresponding segment of the annulus, and prosthetic annuloplasty. Anterior mitral leaflet (AML) prolapse is less common and is not always considered an indication for repair despite availability of a variety of surgical maneuvers specifically designed for its treatment. Although reliable if properly performed, chordal shortening at the papillary muscle level is technically demanding. Chordal transfer from the PML with the 'flip-over' technique is highly reproducible, but limited by the very frequent presence of an abnormal PML. Although feasible, transfer of an anterior basal chord to the prolapsing free edge assumes that the basal chords can be sectioned with impunity. More recently, chord replacement with expanded polytetrafluoroethylene (PTFE) sutures has become increasingly popular because of its availability, theoretical simplicity, and demonstrated long-term durability. Although papillary and leaflet anchoring of the neo-chord has not been shown to be a problem, the determination of its appropriate length remains intuitive and based on personal experience. Here, simple surgical maneuvers designed to ensure safe and reproducible results of single or multiple chord replacement with PTFE sutures are described. PMID:12701786

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

    NASA Technical Reports Server (NTRS)

    Harris, Thomas A; Recant, Isidore G

    1941-01-01

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

  4. The Effect of Variable Chord Length on Transonic Axial Rotor Performance

    NASA Technical Reports Server (NTRS)

    Roberts, William B.; Suder, Kenneth L.; Thorp, Scott A.; Strazisar, Anthony J.; Armin, Albert; Kassaseya, George

    2002-01-01

    Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

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

    NASA Technical Reports Server (NTRS)

    Fischel, Jack; Riebe, John M

    1944-01-01

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

  6. Airfoil System for Cruising Flight

    NASA Technical Reports Server (NTRS)

    Shams, Qamar A. (Inventor); Liu, Tianshu (Inventor)

    2014-01-01

    An airfoil system includes an airfoil body and at least one flexible strip. The airfoil body has a top surface and a bottom surface, a chord length, a span, and a maximum thickness. Each flexible strip is attached along at least one edge thereof to either the top or bottom surface of the airfoil body. The flexible strip has a spanwise length that is a function of the airfoil body's span, a chordwise width that is a function of the airfoil body's chord length, and a thickness that is a function of the airfoil body's maximum thickness.

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

    NASA Technical Reports Server (NTRS)

    Jernell, L. S.

    1976-01-01

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

  8. Chord-length and free-path distribution functions for many-body systems

    SciTech Connect

    Lu, B. ); Torquato, S. )

    1993-04-15

    We study fundamental morphological descriptors of disordered media (e.g., heterogeneous materials, liquids, and amorphous solids): [ital the] [ital chord]-[ital length] [ital distribution] [ital function] [ital p]([ital z]) and the [ital free]-[ital path] [ital distribution] [ital function] [ital p]([ital z],[ital a]). For concreteness, we will speak in the language of heterogeneous materials composed of two different materials or phases.'' The probability density function [ital p]([ital z]) describes the distribution of chord lengths in the sample and is of great interest in stereology. For example, the first moment of [ital p]([ital z]) is the mean intercept length'' or mean chord length.'' The chord-length distribution function is of importance in transport phenomena and problems involving discrete free paths'' of point particles (e.g., Knudsen diffusion and radiative transport). The free-path distribution function [ital p]([ital z],[ital a]) takes into account the finite size of a simple particle of radius [ital a] undergoing discrete free-path motion in the heterogeneous material and we show that it is actually the chord-length distribution function for the system in which the pore space'' is the space available to a finite-sized particle of radius [ital a]. Thus it is shown that [ital p]([ital z])=[ital p]([ital z],0). We demonstrate that the functions [ital p]([ital z]) and [ital p]([ital z],[ital a]) are related to another fundamentally important morphological descriptor of disordered media, namely, the so-called lineal-path function [ital L]([ital z]) studied by us in previous work [Phys. Rev. A [bold 45], 922 (1992)]. The lineal path function gives the probability of finding a line segment of length [ital z] wholly in one of the phases'' when randomly thrown into the sample.

  9. Wind-tunnel investigation of an NACA 23012 airfoil with 30 percent-chord venetian-blind flaps

    NASA Technical Reports Server (NTRS)

    Rogallo, F M; Spano, Bartholomew S

    1942-01-01

    Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a NACA 23012 airfoil with 30-percent-chord venetian-blind flaps having one, two, three, and four slats of Clark y section. The three-slat arrangements was aerodynamically the best of those tested but showed practically no improvement over the comparable arrangement used in the preliminary tests published in NACA Technical Report No. 689. The multiple-slat flaps gave slightly higher lift coefficients than the one-slat (Fowler) flap but gave considerably greater pitching-moment coefficients. An analysis of test data indicates that substitution of a thicker and more cambered section for the Clark y slats should improve the aerodynamic and the structural characteristics of the venetian-blind flap.

  10. Computation of Turbulent Trailing-edge Noise for a Finite-chord Airfoil

    NASA Astrophysics Data System (ADS)

    Oberai, Assad; Roknaldin, Farzam; Hughes, Thomas

    2001-11-01

    We have considered the application of the variational formulation of Lighthill's acoustic analogy to the trailing-edge noise problem. We have used this formulation to study the effect of the finiteness of the chord and the variation of the far-field pressure directivity with frequency. To validate our results, we have made comparisons with analytical solutions for certain limiting cases. We have also used this methodology to determined the far-field acoustic pressure for a low Mach number turbulent flow. To determine the acoustic sources for this problem we have performed a large eddy simulation of the incompressible Navier-Stokes equations. an attachment.

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

    NASA Technical Reports Server (NTRS)

    Stivers, Louis S., Jr.

    1947-01-01

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

  12. Implementation of Chord Length Sampling for Transport Through a Binary Stochastic Mixture

    SciTech Connect

    T.J. Donovan; T.M. Sutton; Y. Danon

    2002-11-18

    Neutron transport through a special case stochastic mixture is examined, in which spheres of constant radius are uniformly mixed in a matrix material. A Monte Carlo algorithm previously proposed and examined in 2-D has been implemented in a test version of MCNP. The Limited Chord Length Sampling (LCLS) technique provides a means for modeling a binary stochastic mixture as a cell in MCNP. When inside a matrix cell, LCLS uses chord-length sampling to sample the distance to the next stochastic sphere. After a surface crossing into a stochastic sphere, transport is treated explicitly until the particle exits or is killed. Results were computed for a simple model with two different fixed neutron source distributions and three sets of material number densities. Stochastic spheres were modeled as black absorbers and varying degrees of scattering were introduced in the matrix material. Tallies were computed using the LCLS capability and by averaging results obtained from multiple realizations of the random geometry. Results were compared for accuracy and figures of merit were compared to indicate the efficiency gain of the LCLS method over the benchmark method. Results show that LCLS provides very good accuracy if the scattering optical thickness of the matrix is small ({le} 1). Comparisons of figures of merit show an advantage to LCLS varying between factors of 141 and 5. LCLS efficiency and accuracy relative to the benchmark both decrease as scattering is increased in the matrix.

  13. Modeling the increase in aerodynamic efficiency for a thick (37.5% chord) airfoil with slot suction in vortex cells with allowance for the compressibility effect

    NASA Astrophysics Data System (ADS)

    Isaev, S. A.; Baranov, P. A.; Sudakov, A. G.; Ermakov, A. M.

    2015-01-01

    The Reynolds equations closed using the Menter shear-stress-transfer model modified with allowance for the curvature of flow lines have been numerically solved using multiblock computational technologies. The obtained solution has been used to analyze subsonic flow past a thick (37.5% chord) airfoil with slot suction in circular vortex cells intended for the Ecology and Progress (Ekologiya i Progress, EKIP) aircraft project in comparison to a distributed (from the central body surface) suction at fixed values of the total volume flow rate (0.02121) and Reynolds number (105) in a range of Mach numbers from 0 to 0.4. This analysis revealed a significant (up to tenfold) decrease in the bow drag (determined with allowance for the energy losses) and a large (by an order of magnitude) increase in the aerodynamic efficiency of the thick airfoil containing vortex cells with slot suction, which reached up to 160.

  14. Shockless airfoils with thicknesses of 20.6 and 20.7 percent chord analytically designed for a Mach number of 0.68 and a lift coefficient of 0.40

    NASA Technical Reports Server (NTRS)

    Allison, D. O.

    1976-01-01

    A 20.8 percent-thick airfoil shape was designed to have shockless inviscid flow at a Mach number of 0.68 and a lift coefficient of 0.40. In order to determine the actual airfoils which would yield this same shockless flow when viscous effects are included, boundary layer displacement thicknesses were subtracted from the inviscid shape for Reynolds numbers of 100 and 35 million. This process yielded airfoils with thicknesses of 20.7 and 20.6 percent, respectively. Subtraction of boundary layer displacement thicknesses for Reynolds numbers below 35 million yielded nonphysical airfoils, that is airfoils with negative thicknesses near tHe trailing edge. The pitching moment about the quarter-chord point at the design condition was -0.082 for the inviscid shape and, consequently, for both airfoils. Off-design calculations for the two airfoils were made using a computer program which provides for the interaction of the inviscid flow and boundary layer solutions. The pressure distributions of the airfoils were shockless for conditions from the design point to lower Mach numbers and lift coefficients. No boundary layer separation was predicted except in the last 3 percent chord on the upper surface.

  15. Airfoil structure

    DOEpatents

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

    1998-01-13

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

  16. Airfoil structure

    DOEpatents

    Frey, Gary A.; Twardochleb, Christopher Z.

    1998-01-01

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

  17. Aerodynamic study of a blade with sine variation of chord length along the height for Darrieus wind turbine

    NASA Astrophysics Data System (ADS)

    Crunteanu, D. E.; Constantinescu, S. G.; Niculescu, M. L.

    2013-10-01

    The wind energy is deemed as one of the most durable energetic variants of the future because the wind resources are immense. Furthermore, one predicts that the small wind turbines will play a vital role in the urban environment. Unfortunately, the complexity and the price of pitch regulated small horizontal-axis wind turbines represent ones of the main obstacles to widespread the use in populated zones. In contrast to these wind turbines, the Darrieus wind turbines are simpler and their price is lower. Unfortunately, their blades run at high variations of angles of attack, in stall and post-stall regimes, which can induce significant vibrations, fatigue and even the wind turbine failure. For this reason, the present paper deals with a blade with sine variation of chord length along the height because it has better behavior in stall and post-stall regimes than the classic blade with constant chord length.

  18. Impact of Spherical Inclusion Mean Chord Length and Radius Distribution on Three-Dimensional Binary Stochastic Medium Particle Transport

    SciTech Connect

    Brantley, P S; Martos, J N

    2011-03-02

    We describe a parallel benchmark procedure and numerical results for a three-dimensional binary stochastic medium particle transport benchmark problem. The binary stochastic medium is composed of optically thick spherical inclusions distributed in an optically thin background matrix material. We investigate three sphere mean chord lengths, three distributions for the sphere radii (constant, uniform, and exponential), and six sphere volume fractions ranging from 0.05 to 0.3. For each sampled independent material realization, we solve the associated transport problem using the Mercury Monte Carlo particle transport code. We compare the ensemble-averaged benchmark fiducial tallies of reflection from and transmission through the spatial domain as well as absorption in the spherical inclusion and background matrix materials. For the parameter values investigated, we find a significant dependence of the ensemble-averaged fiducial tallies on both sphere mean chord length and sphere volume fraction, with the most dramatic variation occurring for the transmission through the spatial domain. We find a weaker dependence of most benchmark tally quantities on the distribution describing the sphere radii, provided the sphere mean chord length used is the same in the different distributions. The exponential distribution produces larger differences from the constant distribution than the uniform distribution produces. The transmission through the spatial domain does exhibit a significant variation when an exponential radius distribution is used.

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

    NASA Technical Reports Server (NTRS)

    Platt, Robert C; Abbott, Ira H

    1937-01-01

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

  20. Wind-tunnel investigation of an NACA 23012 airfoil with several arrangements of slotted flaps with extended lips

    NASA Technical Reports Server (NTRS)

    Lowry, John G

    1941-01-01

    An investigation was made in the NACA 7- by 10-foot wind tunnel to determine the effect of slot-lip location on the aerodynamic section characteristics of an NACA 23012 airfoil with a 30-percent-chord slotted flap. Tests were made with slot lips located at 90 and 100 percent of the airfoil chord and with two different flap shapes. The results are compared with a slotted flap previously developed by the National advisory Committee for Aeronautics with a slot lip located at 83 percent of the airfoil chord. The extension of the slot lip to the rear increased the section lift and pitching-moment coefficients. Comparisons made on a basis of pitching moment for a given tail length show that the Fowler type flap, lip extended to trailing edge of the airfoil, has the greatest section lift coefficient. For moderate tail lengths, 2 to 3 chord lengths, there was only a slight difference between the previously developed slotted flap and the slotted flap with slot lip extended to 90 percent of the airfoil chord. Of the three flaps tested, the Fowler flap had the lowest drag coefficient at high lift coefficients. The extension of the lower surface at the leading edge of the slot had a negligible effect on the profile drag of the airfoil-flap arrangement with the flap deflected when the lip terminated at 90 percent of the airfoil chord.

  1. Airfoil

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

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

  2. High fidelity numerical simulation of airfoil thickness and kinematics effects on flapping airfoil propulsion

    NASA Astrophysics Data System (ADS)

    Yu, Meilin; Wang, Z. J.; Hu, Hui

    2013-10-01

    High-fidelity numerical simulations with the spectral difference (SD) method are carried out to investigate the unsteady flow over a series of oscillating NACA 4-digit airfoils. Airfoil thickness and kinematics effects on the flapping airfoil propulsion are highlighted. It is confirmed that the aerodynamic performance of airfoils with different thickness can be very different under the same kinematics. Distinct evolutionary patterns of vortical structures are analyzed to unveil the underlying flow physics behind the diverse flow phenomena associated with different airfoil thickness and kinematics and reveal the synthetic effects of airfoil thickness and kinematics on the propulsive performance. Thickness effects at various reduced frequencies and Strouhal numbers for the same chord length based Reynolds number (=1200) are then discussed in detail. It is found that at relatively small Strouhal number (=0.3), for all types of airfoils with the combined pitching and plunging motion (pitch angle 20°, the pitch axis located at one third of chord length from the leading edge, pitch leading plunge by 75°), low reduced frequency (=1) is conducive for both the thrust production and propulsive efficiency. Moreover, relatively thin airfoils (e.g. NACA0006) can generate larger thrust and maintain higher propulsive efficiency than thick airfoils (e.g. NACA0030). However, with the same kinematics but at relatively large Strouhal number (=0.45), it is found that airfoils with different thickness exhibit diverse trend on thrust production and propulsive efficiency, especially at large reduced frequency (=3.5). Results on effects of airfoil thickness based Reynolds numbers indicate that relative thin airfoils show superior propulsion performance in the tested Reynolds number range. The evolution of leading edge vortices and the interaction between the leading and trailing edge vortices play key roles in flapping airfoil propulsive performance.

  3. On the use of area-averaged void fraction and local bubble chord length entropies as two-phase flow regime indicators

    NASA Astrophysics Data System (ADS)

    Hernández, Leonor; Juliá, J. Enrique; Paranjape, Sidharth; Hibiki, Takashi; Ishii, Mamoru

    2010-11-01

    In this work, the use of the area-averaged void fraction and bubble chord length entropies is introduced as flow regime indicators in two-phase flow systems. The entropy provides quantitative information about the disorder in the area-averaged void fraction or bubble chord length distributions. The CPDF (cumulative probability distribution function) of void fractions and bubble chord lengths obtained by means of impedance meters and conductivity probes are used to calculate both entropies. Entropy values for 242 flow conditions in upward two-phase flows in 25.4 and 50.8-mm pipes have been calculated. The measured conditions cover ranges from 0.13 to 5 m/s in the superficial liquid velocity j f and ranges from 0.01 to 25 m/s in the superficial gas velocity j g. The physical meaning of both entropies has been interpreted using the visual flow regime map information. The area-averaged void fraction and bubble chord length entropies capability as flow regime indicators have been checked with other statistical parameters and also with different input signals durations. The area-averaged void fraction and the bubble chord length entropies provide better or at least similar results than those obtained with other indicators that include more than one parameter. The entropy is capable to reduce the relevant information of the flow regimes in only one significant and useful parameter. In addition, the entropy computation time is shorter than the majority of the other indicators. The use of one parameter as input also represents faster predictions.

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

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

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

  5. Airfoil

    DOEpatents

    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.

  6. A low speed two-dimensional study of flow separation on the GA(W)-1 airfoil with 30-percent chord Fowler flap

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    Measurements of flow fields with low speed turbulent boundary layers were made for the GA(W)-1 airfoil with a 0.30 c Fowler flap deflected 40 deg at angles of attack of 2.7 deg, 7.7 deg, and 12.8 deg, at a Reynolds number of 2.2 million, and a Mach number of 0.13. Details of velocity and pressure fields associated with the airfoil flap combination are presented for cases of narrow, optimum and wide slot gaps. Extensive flow field turbulence surveys were also conducted employing hot-film anemometry. For the optimum gap setting, the boundaries of the regions of flow reversal within the wake were determined by this technique for two angles of attack. Local skin friction distributions for the basic airfoil and the airfoil with flap (optimum gap) were obtained using the razor blade technique.

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

    NASA Astrophysics Data System (ADS)

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

    2010-04-01

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

  8. Porous airfoil and process

    NASA Technical Reports Server (NTRS)

    Hartwich, Peter M. (Inventor)

    1992-01-01

    A porous airfoil having venting cavities with contoured barrier walls, formed by a core piece, placed beneath a porous upper and lower surface area that stretches over the nominal chord of an airfoil is employed, to provide an airfoil configuration that becomes self-adaptive to very dissimilar flow conditions to thereby improve the lift and drag characteristics of the airfoil at both subcritical and supercritical conditions.

  9. 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.

  10. 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.

  11. Additional flow field studies of the GA(W)-1 airfoil with 30-percent chord Fowler flap including slot-gap variations and cove shape modifications

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    Experimental measurements were made to determine the effects of slot gap opening and flap cove shape on flap and airfoil flow fields. Test model was the GA(W)-1 airfoil with 0.30c Fowler flap deflected 35 degrees. Tests were conducted with optimum, wide and narrow gaps, and with three cove shapes. Three test angles were selected, corresponding to pre-stall and post-stall conditions. Reynolds number was 2,200,000 and Mach number was 0.13. Force, surface pressure, total pressure, and split-film turbulence measurements were made. Results were compared with theory for those parameters for which theoretical values were available.

  12. 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.

  13. Control of laminar separation over airfoils by acoustic excitation

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

    The effect of acoustic excitation in reducing laminar separation over two-dimensional airfoils at low angles of attack is investigated experimentally. Airfoils of two different cross sections, each with two different chord lengths, are studied in the chord Reynolds number range of 25,000 is less than R sub c is less than 100,000. While keeping the amplitude of the excitation induced velocity perturbation a constant, it is found that the most effective frequency scales as U (sup 3/2)(sub infinity). The parameter St/R (sup 1/2)(sub c), corresponding to the most effective f sub p for all the cases studied, falls in the range of 0.02 to 0.03, St being the Strouhal number based on the chord.

  14. 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.

  15. 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. PMID:27347517

  16. 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

  17. Acoustic performance of low pressure axial fan rotors with different blade chord length and radial load distribution

    NASA Astrophysics Data System (ADS)

    Carolus, Thomas

    The paper examines the acoustic and aerodynamic performance of low-pressure axial fan rotors with a hub/tip ratio of 0.45. Six rotors were designed for the same working point by means of the well-known airfoil theory. The condition of an equilibrium between the static pressure gradient and the centrifugal forces is maintained. All rotors have unequally spaced blades to diminish tonal noise. The rotors are tested in a short cylindrical housing without guide vanes. All rotors show very similar flux-pressure difference characteristics. The peak efficiency and the noise performance is considerably influenced by the chosen blade design. The aerodynamically and acoustically optimal rotor is the one with the reduced load at the hub and increased load in the tip region under satisfied equilibrium conditions. It runs at the highest aerodynamic efficiency, and its noise spectrum is fairly smooth. The overall sound pressure level of this rotor is up to 8 dB (A) lower compared to the other rotors under consideration.

  18. An exact, closed-form expression of the integral chord-length distribution for the calculation of single-event upsets induced by cosmic rays

    NASA Technical Reports Server (NTRS)

    Luke, Keung L.; Buehler, Martin G.

    1988-01-01

    This paper presents a derivation of an exact closed-form expression of the integral chord-length distribution for the calculation of single-event upsets (SEUs) in an electronic memory cell, caused by cosmic rays. Results computed for two rectangular parallelepipeds using this exact expression are compared with those computed with Bradford's (1979) semiexact expression of C(x). It is found that the values of C(x) are identical for x equal or smaller than b but are vastly different for x greater than b. Moreover, while C(x) of Bradford gives reasonably accurate values of SEU rate for certain sets of computational parameters, it gives values more than 10 times larger than the correct values for other sets of parameters.

  19. Stall flutter of NACA 0012 airfoil at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Bhat, Shantanu S.; Govardhan, Raghuraman N.

    2013-08-01

    In the present work, we experimentally study and demarcate the stall flutter boundaries of a NACA 0012 airfoil at low Reynolds numbers (Re˜104) by measuring the forces and flow fields around the airfoil when it is forced to oscillate. The airfoil is placed at large mean angle of attack (αm), and is forced to undergo small amplitude pitch oscillations, the amplitude (Δα) and frequency (f) of which are systematically varied. The unsteady loads on the oscillating airfoil are directly measured, and are used to calculate the energy transfer to the airfoil from the flow. These measurements indicate that for large mean angles of attack of the airfoil (αm), there is positive energy transfer to the airfoil over a range of reduced frequencies (k=πfc/U), indicating that there is a possibility of airfoil excitation or stall flutter even at these low Re (c=chord length). Outside this range of reduced frequencies, the energy transfer is negative and under these conditions the oscillations would be damped. Particle Image Velocimetry (PIV) measurements of the flow around the oscillating airfoil show that the shear layer separates from the leading edge and forms a leading edge vortex, although it is not very clear and distinct due to the low oscillation amplitudes. On the other hand, the shear layer formed after separation is found to clearly move periodically away from the airfoil suction surface and towards it with a phase lag to the airfoil oscillations. The phase of the shear layer motion with respect to the airfoil motions shows a clear difference between the exciting and the damping case.

  20. 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

  1. Numerical Investigations of an Optimized Airfoil with a Rotary Cylinder

    NASA Astrophysics Data System (ADS)

    Gada, Komal; Rahai, Hamid

    2015-11-01

    Numerical Investigations of an optimized thin airfoil with a rotary cylinder as a control device for reducing separation and improving lift to drag ratio have been performed. Our previous investigations have used geometrical optimization for development of an optimized airfoil with increased torque for applications in a vertical axis wind turbine. The improved performance was due to contributions of lift to torque at low angles of attack. The current investigations have been focused on using the optimized airfoil for micro-uav applications with an active flow control device, a rotary cylinder, to further control flow separation, especially during wind gust conditions. The airfoil has a chord length of 19.66 cm and a width of 25 cm with 0.254 cm thickness. Previous investigations have shown flow separation at approximately 85% chord length at moderate angles of attack. Thus the rotary cylinder with a 0.254 cm diameter was placed slightly downstream of the location of flow separation. The free stream mean velocity was 10 m/sec. and investigations have been performed at different cylinder's rotations with corresponding tangential velocities higher than, equal to and less than the free stream velocity. Results have shown more than 10% improvement in lift to drag ratio when the tangential velocity is near the free stream mean velocity. Graduate Assistant, Center for Energy and Environmental Research and Services (CEERS), College of Engineering, California State University, Long Beach.

  2. Chord Splicing & Joining Detail; Chord & CrossBracing Joint Details; ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Chord Splicing & Joining Detail; Chord & Cross-Bracing Joint Details; Cross Bracing Center Joint Detail; Chord & Diagonal Joint Detail - Vermont Covered Bridge, Highland Park, spanning Kokomo Creek at West end of Deffenbaugh Street (moved to), Kokomo, Howard County, IN

  3. Arch & Chord Joint Detail; Crossbracing Center Joint Detail; Chord, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Arch & Chord Joint Detail; Crossbracing Center Joint Detail; Chord, Panel Post, Tie & Diagonal Brace Joint Detail; Chord, Panel Post, Tie & Crossbracing Joint Detail - Dunlapsville Covered Bridge, Spanning East Fork Whitewater River, Dunlapsville, Union County, IN

  4. 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.

  5. 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).

  6. 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.

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

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed

    1999-01-01

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

  8. Drag reduction of a blunt trailing-edge airfoil

    NASA Astrophysics Data System (ADS)

    Baker, Jonathon Paul

    Wind-tunnel experimentation and Reynolds-averaged Navier--Stokes simulations were used to analyze simple, static trailing-edge devices applied to an FB-3500-1750 airfoil, a 35% thick airfoil with a 17.5% chord blunt trailing edge, in order to mitigate base drag. The drag reduction devices investigated include Gurney-type tabs, splitter plates, base cavities, and offset cavities. The Gurney-type tabs consisted of small tabs, attached at the trailing edge and distributed along the span, extending above the upper and lower surfaces of the airfoil. The Gurney-type devices were determined to have little drag reduction capabilities for the FB-3500-1750 airfoil. Splitter plates, mounted to the center of the trailing edge, with lengths between 50% and 150% of the trailing-edge thickness and various plate angles (0° and +/-10° from perpendicular) were investigated and shown to influence the lift and drag characteristics of the baseline airfoil. Drag reductions of up to 50% were achieved with the addition of a splitter plate. The base cavity was created by adding two plates perpendicular to the trailing edge, extending from the upper and lower surfaces of the airfoil. The base cavity demonstrated possible drag reductions of 25%, but caused significant changes to lift, primarily due to the method of device implementation. The offset cavity, created by adding two splitter plates offset from the upper and lower surfaces by 25% of the trailing-edge thickness, was shown to improve on the drag reductions of the splitter plate, while also eliminating unsteady vortex shedding prior to airfoil stall.

  9. Airfoil design for variable RPM horizontal axis wind turbines

    NASA Astrophysics Data System (ADS)

    Bjoerck, Anders

    1990-01-01

    The design criteria for new airfoils for a variable speed horizontal axis wind turbine are described. The two series of airfoils developed are characterized by high design lift coefficients in order to achieve small blade chords, high lift drag ratios for the airfoil sections designed for the outer part of the blade, performance insensitivity to surface roughness, and a gentle stall at an angle of attack in order to reduce excessive loads. Each series consists of airfoils with varying thickness to chord ratios for different radial stations. Interpolation between the two series is possible.

  10. Impingement of water droplets on wedges and double-wedge airfoils at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Serafini, John S

    1954-01-01

    An analytical solution has been obtained for the equations of motion of water droplets impinging on a wedge in a two-dimensional supersonic flow field with a shock wave attached to the wedge. The closed-form solution yields analytical expressions for the equation of the droplet trajectory, the local rate of impingement and the impingement velocity at any point on the wedge surface, and the total rate of impingement. The analytical expressions are utilized to determine the impingement on the forward surfaces of diamond airfoils in supersonic flow fields with attached shock waves. The results presented include the following conditions: droplet diameters from 2 to 100 microns, pressure altitudes from sea level to 30,000 feet, free-stream static temperatures from 420 degrees r, free stream Mach numbers from 1.1 to 2.0, semiapex angles for the wedge from 1.14 degrees to 7.97 degrees, thickness-to-chord ratios for the diamond airfoil from 0.02 to 0.14, chord lengths from 1 to 20 feet, and angles of attack from zero to the inverse tangent of the airfoil thickness-to-chord ratio.

  11. 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.

  12. 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.

  13. The effects of actuation frequency on the separation control over an airfoil using a synthetic jet

    NASA Astrophysics Data System (ADS)

    Abe, Y.; Okada, K.; Nonomura, T.; Fujii, K.

    2015-06-01

    The simulation of separation control using a synthetic jet (SJ) is conducted around an NACA (National Advisory Committee for Aeronautics) 0015 airfoil by large-eddy simulation (LES) with a compact difference scheme. The synthetic jet is installed at the leading edge of the airfoil and the effects of an actuation frequency F+ (normalized by chord length and velocity of freestream) are observed. The lift-drag coefficient is recovered the most for F+ = 6. The relationship between momentum addition by turbulent mixing and large vortex structures is investigated using a phase-averaging procedure based on F+. The Reynolds shear stress is decomposed into periodic and turbulent components where the turbulent components are found to be dominant on the airfoil. The strong turbulent components appear near the large vortex structures that are observed in phase- and span-averaged flow fields.

  14. Darrieus wind-turbine airfoil configurations

    NASA Astrophysics Data System (ADS)

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

    1982-06-01

    The purpose was to determine what aerodynamic performance improvement, if any, could be achieved by judiciously choosing the airfoil sections for Darrieus wind turbine blades. 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. 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.

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

    NASA Astrophysics Data System (ADS)

    Zhang, Qiang

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

  16. 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

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

    NASA Technical Reports Server (NTRS)

    Platt, Robert C; Shortal, Joseph A

    1937-01-01

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

  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. Airfoil design: Finding the balance between design lift and structural stiffness

    NASA Astrophysics Data System (ADS)

    Bak, Christian; Gaudern, Nicholas; Zahle, Frederik; Vronsky, Tomas

    2014-06-01

    When upscaling wind turbine blades there is an increasing need for high levels of structural efficiency. In this paper the relationships between the aerodynamic characteristics; design lift and lift-drag ratio; and the structural characteristics were investigated. Using a unified optimization setup, airfoils were designed with relative thicknesses between 18% and 36%, a structural box height of 85% of the relative thickness, and varying box widths in chordwise direction between 20% and 40% of the chord length. The results from these airfoil designs showed that for a given flapwise stiffness, the design lift coefficient increases if the box length reduces and at the same time the relative thickness increases. Even though the conclusions are specific to the airfoil design approach used, the study indicated that an increased design lift required slightly higher relative thickness compared to airfoils with lower design lift to maintain the flapwise stiffness. Also, the study indicated that the lift-drag ratio as a function of flapwise stiffness was relatively independent of the airfoil design with a tendency that the lift-drag ratio decreased for large box lengths. The above conclusions were supported by an analysis of the three airfoil families Riso-C2, DU and FFA, where the lift-drag ratio as a function of flapwise stiffness was decreasing, but relatively independent of the airfoil design, and the design lift coefficient was varying depending on the design philosophy. To make the analysis complete also design lift and lift- drag ratio as a function of edgewise and torsional stiffness were shown.

  20. 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.

  1. Wind-tunnel investigation of an NACA 23030 airfoil with various arrangements of slotted flaps

    NASA Technical Reports Server (NTRS)

    Recant, I G

    1940-01-01

    AN investigation was made of a large-chord NACA 23030 airfoil with a 40- and a 25.66 percent-chord slotted flap to determine the section aerodynamic characteristics of the airfoil affected by flap chord, slot shape, flap position, and flap deflection. The flap positions for maximum lift, the position for minimum drag at moderate and high lift coefficients, and the complete section aerodynamic characteristics of selected optimum arrangements are given. Envelope polar of various flap arrangements are included. The relative merits of slotted flaps of different chords on the NACA 23030 airfoil are discussed, and a comparison is made of each flap size with a corresponding flap size on the NACA 23021 and 23012 airfoils. The lowest profile drags at moderate lift coefficients were obtained with an easy entrance to the slot. The 25.66-percent-chord slotted flap gave lower drag than the 40-percent-chord flap for lift coefficients less than 1.8, but the 40-percent-chord flap gave considerably lower drag for lift coefficients. The drag coefficients at moderate and high lift coefficients were greater with both sizes of flap on the NACA 23030 airfoil than on either the NACA 23021 or the NACA 23012 airfoil. The maximum lift coefficient for the deflections tested with either flap was practically independent of airfoil.

  2. Effect of Variation of Chord and Span of Ailerons on Rolling and Yawing Moments at Several Angles of Pitch

    NASA Technical Reports Server (NTRS)

    Heald, R H; Strother, D H; Monish, B H

    1931-01-01

    This report presents the results of an extension to higher angles of attack of the investigation of the rolling and yawing moments due to ailerons of various chords and spans on two airfoils having the Clark Y and U. S. A. 27 wings. The measurements were made at various angles of pitch but at zero angle of roll and yaw, the wing chord being set at an angle of +4 degrees to the fuselage axis. In the case of the Clark Y airfoil the measurements have been extended to a pitch angle of 40 degrees, using ailerons of span equal to 67 per cent of the wing semispan and chord equal to 20 and 30 per cent of the wing chord. The work was conducted on wing models of 60-inch span and 10-inch chord.

  3. 16. Typical end post, top chord, bottom chord, vertical lattice ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    16. Typical end post, top chord, bottom chord, vertical lattice and diagonal connections for the 2nd and 3rd spans. View is of south side of 3rd span just west of its connection with the 4th span. - Cleves Bridge, Spanning Great Miami River on U.S. Highway 50, Cleves, Hamilton County, OH

  4. 27. Typical end post, top chord, bottom chord, vertical lattice ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    27. Typical end post, top chord, bottom chord, vertical lattice and diagonal connections for the 1st and 4th spans. View is of south side of 4th span just east of its connection with the 3rd span. - Cleves Bridge, Spanning Great Miami River on U.S. Highway 50, Cleves, Hamilton County, OH

  5. Time resolved PIV analysis of flow over a NACA 0015 airfoil with Gurney flap

    NASA Astrophysics Data System (ADS)

    Troolin, D. R.; Longmire, E. K.; Lai, W. T.

    2006-08-01

    A NACA 0015 airfoil with and without a Gurney flap was studied in a wind tunnel with Re c = 2.0 × 105 in order to examine the evolving flow structure of the wake through time-resolved PIV and to correlate this structure with time-averaged measurements of the lift coefficient. The Gurney flap, a tab of small length (1-4% of the airfoil chord) that protrudes perpendicular to the chord at the trailing edge, yields a significant and relatively constant lift increment through the linear range of the C L versus α curve. Two distinct vortex shedding modes were found to exist and interact in the wake downstream of flapped airfoils. The dominant mode resembles a Kàrmàn vortex street shedding behind an asymmetric bluff body. The second mode, which was caused by the intermittent shedding of fluid recirculating in the cavity upstream of the flap, becomes more coherent with increasing angle of attack. For a 4% Gurney flap at α = 8°, the first and second modes corresponded with Strouhal numbers based on flap height of 0.18 and 0.13. Comparison of flow around ‘filled’ and ‘open’ flap configurations suggested that the second shedding mode was responsible for a significant portion of the overall lift increment.

  6. 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.

  7. Angle-of-attack validation of a new zonal CFD method for airfoil simulations

    NASA Technical Reports Server (NTRS)

    Yoo, Sungyul; Summa, J. Michael; Strash, Daniel J.

    1990-01-01

    The angle-of-attack validation of a new concept suggested by Summa (1990) for coupling potential and viscous flow methods has been investigated for two-dimensional airfoil simulations. The fully coupled potential/Navier-Stokes code, ZAP2D (Zonal Aerodynamics Program 2D), has been used to compute the flow field around an NACA 0012 airfoil for a range of angles of attack up to stall at a Mach number of 0.3 and a Reynolds number of 3 million. ZAP2D calculation for various domain sizes from 25 to 0.12 chord lengths are compared with the ARC2D large domain solution as well as with experimental data.

  8. Mach number validation of a new zonal CFD method (ZAP2D) for airfoil simulations

    NASA Technical Reports Server (NTRS)

    Strash, Daniel J.; Summa, Michael; Yoo, Sungyul

    1991-01-01

    A closed-loop overlapped velocity coupling procedure has been utilized to combine a two-dimensional potential-flow panel code and a Navier-Stokes code. The fully coupled two-zone code (ZAP2D) has been used to compute the flow past a NACA 0012 airfoil at Mach numbers ranging from 0.3 to 0.84 near the two-dimensional airfoil C(lmax) point for a Reynolds number of 3 million. For these cases, the grid domain size can be reduced to 3 chord lengths with less than 3-percent loss in accuracy for freestream Mach numbers through 0.8. Earlier validation work with ZAP2D has demonstrated a reduction in the required Navier-Stokes computation time by a factor of 4 for subsonic Mach numbers. For this more challenging condition of high lift and Mach number, the saving in CPU time is reduced to a factor of 2.

  9. Comparison of sound power radiation from isolated airfoils and cascades in a turbulent flow.

    PubMed

    Blandeau, Vincent P; Joseph, Phillip F; Jenkins, Gareth; Powles, Christopher J

    2011-06-01

    An analytical model of the sound power radiated from a flat plate airfoil of infinite span in a 2D turbulent flow is presented. The effects of stagger angle on the radiated sound power are included so that the sound power radiated upstream and downstream relative to the fan axis can be predicted. Closed-form asymptotic expressions, valid at low and high frequencies, are provided for the upstream, downstream, and total sound power. A study of the effects of chord length on the total sound power at all reduced frequencies is presented. Excellent agreement for frequencies above a critical frequency is shown between the fast analytical isolated airfoil model presented in this paper and an existing, computationally demanding, cascade model, in which the unsteady loading of the cascade is computed numerically. Reasonable agreement is also observed at low frequencies for low solidity cascade configurations. PMID:21682379

  10. 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.

  11. The NASA Langley laminar flow control airfoil experiment

    NASA Technical Reports Server (NTRS)

    Harvey, W. D.; Pride, J. D.

    1982-01-01

    A large chord swept supercritical LFC airfoil has been constructed for NASA-Langley's research program to determine the compatibility of supercritical airfoils with suction laminarization and to establish a technology base for future transport designs. Features include a high design Mach number and shock-free flow, as well as the minimization of the laminarization suction through a choice of airfoil geometry and pressure distribution. Two suction surface concepts and a variety of hybrid suction concepts involving combinations of natural and forced laminar flow are to be investigated. The test facility has been modified to insure achievement of required flow quality and transonic interference-free flow over the yawed LFC airfoil.

  12. 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.

  13. 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.

  14. Ground effect on the aerodynamics of a two-dimensional oscillating airfoil

    NASA Astrophysics Data System (ADS)

    Lu, H.; Lua, K. B.; Lim, T. T.; Yeo, K. S.

    2014-07-01

    This paper reports results of an experimental investigation into ground effect on the aerodynamics of a two-dimensional elliptic airfoil undergoing simple harmonic translation and rotational motion. Ground clearance ( D) ranging from 1 c to 5 c (where c is the airfoil chord length) was investigated for three rotational amplitudes ( α m) of 30°, 45° and 60° (which respectively translate to mid-stroke angle of attack of 60°, 45° and 30°). For the lowest rotational amplitude of 30°, results show that an airfoil approaching a ground plane experiences a gradual decrease in cycle-averaged lift and drag coefficients until it reaches D ≈ 2.0 c, below which they increase rapidly. Corresponding DPIV measurement indicates that the initial force reduction is associated with the formation of a weaker leading edge vortex and the subsequent force increase below D ≈ 2.0 c may be attributed to stronger wake capture effect. Furthermore, an airfoil oscillating at higher amplitude lessens the initial force reduction when approaching the ground and this subsequently leads to lift distribution that bears striking resemblance to the ground effect on a conventional fixed wing in steady translation.

  15. Direct simulations of trailing-edge noise generation from two-dimensional airfoils at low Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Ikeda, Tomoaki; Atobe, Takashi; Takagi, Shohei

    2012-01-01

    The aeroacoustic sound generated from the flow around two NACA four-digit airfoils is investigated numerically, at relatively low Reynolds numbers that do not prompt boundary-layer transition. By using high-order finite-difference schemes to discretize compressible Navier-Stokes equations, the sound scattered on airfoil surface is directly resolved as an unsteady pressure fluctuation. As the wavelength of an emitted noise is shortened compared to the airfoil chord, the diffraction effect on non-compact chord length appears more noticeable, developing multiple lobes in directivity. The instability mechanism that produces sound sources, or unsteady vortical motions, is quantitatively examined, also by using a linear stability theory. While the evidence of boundary-layer instability waves is captured in the present result, the most amplified frequency in the boundary shear layer does not necessarily agree with the primary frequency of a trailing-edge noise, when wake instability is dominant in laminar flow. This contradicts the observation of other trailing-edge noise studies at higher Reynolds numbers. However, via acoustic disturbances, the boundary-layer instability may become more significant, through the resonance with the wake instability, excited by increasing a base-flow Mach number. Evidence suggests that this would correspond to the onset of an acoustic feedback loop. The wake-flow frequencies derived by an absolute-instability analysis are compared with the frequencies realized in flow simulations, to clarify the effect of an acoustic feedback mechanism, at a low Reynolds number.

  16. Pressure Distribution Over Airfoils with Fowler Flaps

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Anderson, Walter B

    1938-01-01

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

  17. Chord distribution along a line in the local Universe

    NASA Astrophysics Data System (ADS)

    Zaninetti, L.

    2013-04-01

    A method is developed to compute the chord length distribution along a line which intersects a cellular universe. The cellular universe is here modeled by the Poissonian Voronoi Tessellation (PVT) and by a non-Poissonian Voronoi Tessellation (NPVT). The distribution of the spheres is obtained from common approximations used in modeling the volumes of Voronoi Diagrams. We give analytical formulas for the distributions of the lengths of chords in both the PVT and NPVT. The astrophysical applications are made to the real ESO Slice Project and to an artificial slice of galaxies which simulates the 2dF Galaxy Redshift Survey.

  18. PIC Detector for Piano Chords

    NASA Astrophysics Data System (ADS)

    Barbancho, Ana M.; Tardón, Lorenzo J.; Barbancho, Isabel

    2010-12-01

    In this paper, a piano chords detector based on parallel interference cancellation (PIC) is presented. The proposed system makes use of the novel idea of modeling a segment of music as a third generation mobile communications signal, specifically, as a CDMA (Code Division Multiple Access) signal. The proposed model considers each piano note as a CDMA user in which the spreading code is replaced by a representative note pattern. The lack of orthogonality between the note patterns will make necessary to design a specific thresholding matrix to decide whether the PIC outputs correspond to the actual notes composing the chord or not. An additional stage that performs an octave test and a fifth test has been included that improves the error rate in the detection of these intervals that are specially difficult to detect. The proposed system attains very good results in both the detection of the notes that compose a chord and the estimation of the polyphony number.

  19. 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.

  20. 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.

  1. A computer program for the design and analysis of low-speed airfoils, supplement

    NASA Technical Reports Server (NTRS)

    Eppler, R.; Somers, D. M.

    1980-01-01

    Three new options were incorporated into an existing computer program for the design and analysis of low speed airfoils. These options permit the analysis of airfoils having variable chord (variable geometry), a boundary layer displacement iteration, and the analysis of the effect of single roughness elements. All three options are described in detail and are included in the FORTRAN IV computer program.

  2. Determination of Boundary-Layer Transition on Three Symmetrical Airfoils in the NACA Full-Scale Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Silverstein, Abe; Becker, John V

    1938-01-01

    For the purpose of studying the transition from laminar to turbulent flow, boundary-layer measurements were made in the NACA full-scale wind tunnel on three symmetrical airfoils of NACA 0009, 0012, and 0018 sections. The effects of variations in lift coefficient, Reynolds number, and airfoil thickness on transition were investigated. Air speed in the boundary layer was measured by total-head tubes and by hot wires; a comparison of transition as indicated by the two techniques was obtained. The results indicate no unique value of Reynolds number for the transition, whether the Reynolds number is based upon the distance along the chord or upon the thickness of the boundary layer at the transition point. In general, the transition is not abrupt and occurs in a region that varies in length as a function of the test conditions.

  3. A bistable mechanism for chord extension morphing rotors

    NASA Astrophysics Data System (ADS)

    Johnson, Terrence; Frecker, Mary; Gandhi, Farhan

    2009-03-01

    Research efforts have shown that helicopter rotor blade morphing is an effective means to improve flight performance. Previous example of rotor blade morphing include using smart-materials for trailing deflection and rotor blade twist and tip twist, the development of a comfortable airfoil using compliant mechanisms, the use of a Gurney flap for air-flow deflection and centrifugal force actuated device to increase the span of the blade. In this paper we explore the use of a bistable mechanism for rotor morphing, specifically, blade chord extension using a bistable arc. Increasing the chord of the rotor blade is expected to generate more lift-load and improve helicopter performance. Bistable or "snap through" mechanisms have multiple stable equilibrium states and are a novel way to achieve large actuation output stroke. Bistable mechanisms do not require energy input to maintain a stable equilibrium state as both states do not require locking. In this work, we introduce a methodology for the design of bistable arcs for chord morphing using the finite element analysis and pseudo-rigid body model, to study the effect of different arc types, applied loads and rigidity on arc performance.

  4. Holography and LDV techniques, their status and use in airfoil research

    NASA Technical Reports Server (NTRS)

    Johnson, D. A.; Bachalo, W. D.

    1978-01-01

    The measurement capabilities of laser velocimetry and holographic interferometry in transonic airfoil testing were demonstrated. Presented are representative results obtained with these two nonintrusive techniques on a 15.24 cm chord airfoil section. These results include the density field about the airfoil, flow angles in the inviscid flow and viscous flow properties including the turbulent Reynolds stresses. The accuracies of the density fields obtained by interferometry were verified from comparisons with surface pressure and laser velocimeter measurements.

  5. Notes on the theoretical characteristics of two-dimensional supersonic airfoils

    NASA Technical Reports Server (NTRS)

    Ivey, H Reese

    1947-01-01

    The shock expansion method of the NACA TN No. 1143 was used to determine the principal aerodynamic characteristics of two-dimensional supersonic airfoils. A discussion is given of the effect of thickness ratio, free-stream Mach number, angle of attack, camber, thickness distribution, and aileron deflection. The calculations indicated that the minimum drag of supersonic airfoils is obtained when the maximum thickness is behind the 0.50 chord. The center of pressure obtained for a symmetrical supersonic airfoil was found to be ahead of the 0.50 chord.

  6. Development of a computer program to obtain ordinates for NACA-6 and 6A-series airfoils

    NASA Technical Reports Server (NTRS)

    Ladson, C. L.; Brooks, C. W., Jr.

    1974-01-01

    A computer program was developed to produce the ordinates for airfoils of any thickness, thickness distribution, or camber in the NACA 6- and 6A-series. For the 6-series and for all but the leading edge of the 6A-series, agreement between the ordinates obtained from the new program and previously published values is generally within .00005 chord. Near the leading edge of the 6A-series airfoils, differences up to .00035 chord are found.

  7. Correlations Among Ice Measurements, Impingement Rates Icing Conditions, and Drag Coefficients for Unswept NACA 65A004 Airfoil

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H.

    1958-01-01

    An empirical relation has been obtained by which the change in drag coefficient caused by ice formations on an unswept NACA 65AO04 airfoil section can be determined from the following icing and operating conditions: icing time, airspeed, air total temperature, liquid-water content, cloud droplet impingement efficiencies, airfoil chord length, and angles of attack. The correlation was obtained by use of measured ice heights and ice angles. These measurements were obtained from a variety of ice formations, which were carefully photographed, cross-sectioned, and weighed. Ice weights increased at a constant rate with icing time in a rime icing condition and at progressively increasing rates in glaze icing conditions. Initial rates of ice collection agreed reasonably well with values predicted from droplet impingement data. Experimental droplet impingement rates obtained on this airfoil section agreed with previous theoretical calculations for angles of attack of 40 or less. Disagreement at higher angles of attack was attributed to flow separation from the upper surface of the experimental airfoil model.

  8. Models of Lift and Drag Coefficients of Stalled and Unstalled Airfoils in Wind Turbines and Wind Tunnels

    NASA Technical Reports Server (NTRS)

    Spera, David A.

    2008-01-01

    Equations are developed with which to calculate lift and drag coefficients along the spans of torsionally-stiff rotating airfoils of the type used in wind turbine rotors and wind tunnel fans, at angles of attack in both the unstalled and stalled aerodynamic regimes. Explicit adjustments are made for the effects of aspect ratio (length to chord width) and airfoil thickness ratio. Calculated lift and drag parameters are compared to measured parameters for 55 airfoil data sets including 585 test points. Mean deviation was found to be -0.4 percent and standard deviation was 4.8 percent. When the proposed equations were applied to the calculation of power from a stall-controlled wind turbine tested in a NASA wind tunnel, mean deviation from 54 data points was -1.3 percent and standard deviation was 4.0 percent. Pressure-rise calculations for a large wind tunnel fan deviated by 2.7 percent (mean) and 4.4 percent (standard). The assumption that a single set of lift and drag coefficient equations can represent the stalled aerodynamic behavior of a wide variety of airfoils was found to be satisfactory.

  9. 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.

  10. A Theory of Unstaggered Airfoil Cascades in Compressible Flow

    NASA Technical Reports Server (NTRS)

    Spurr, Robert A.; Allen, H. Julian

    1947-01-01

    By use of the methods of thin airfoil theory, which include effects of compressibility, rela.tio^as are developed which permit the rapid determination of the pressure distribution over an unstaggered cascade of airfoils of a given profile, and the determination of the profile shape necessary to yield a given pressure distribution for small chord gap ratios, For incompressible flow the results of the theory are compared with available examples obtained by the more exact method of conformal transformation. Although the theory is developed for small chord/gap ratios, these comparisons show that it may be extended to chord/gap ratios of order unity, at least for low speed flows. Choking of cascades, a phenomenon of particular importance in compressor design, is considered.

  11. Aerodynamic effects of simulated ice shapes on two-dimensional airfoils and a swept finite tail

    NASA Astrophysics Data System (ADS)

    Alansatan, Sait

    An experimental study was conducted to investigate the effect of simulated glaze ice shapes on the aerodynamic performance characteristics of two-dimensional airfoils and a swept finite tail. The two dimensional tests involved two NACA 0011 airfoils with chords of 24 and 12 inches. Glaze ice shapes computed with the LEWICE code that were representative of 22.5-min and 45-min ice accretions were simulated with spoilers, which were sized to approximate the horn heights of the LEWICE ice shapes. Lift, drag, pitching moment, and surface pressure coefficients were obtained for a range of test conditions. Test variables included Reynolds number, geometric scaling, control deflection and the key glaze ice features, which were horn height, horn angle, and horn location. For the three-dimensional tests, a 25%-scale business jet empennage (BJE) with a T-tail configuration was used to study the effect of ice shapes on the aerodynamic performance of a swept horizontal tail. Simulated glaze ice shapes included the LEWICE and spoiler ice shapes to represent 9-min and 22.5-min ice accretions. Additional test variables included Reynolds number and elevator deflection. Lift, drag, hinge moment coefficients as well as boundary layer velocity profiles were obtained. The experimental results showed substantial degradation in aerodynamic performance of the airfoils and the swept horizontal tail due to the simulated ice shapes. For the two-dimensional airfoils, the largest aerodynamic penalties were obtained when the 3-in spoiler-ice, which was representative of 45-min glaze ice accretions, was set normal to the chord. Scale and Reynolds effects were not significant for lift and drag. However, pitching moments and pressure distributions showed great sensitivity to Reynolds number and geometric scaling. For the threedimensional study with the swept finite tail, the 22.5-min ice shapes resulted in greater aerodynamic performance degradation than the 9-min ice shapes. The addition of 24

  12. Pressure distribution over NACA 23012 airfoil with a slotted and a split flap

    NASA Technical Reports Server (NTRS)

    Harris, Thomas A; Lowry, John G

    1941-01-01

    A pressure-distribution investigation has been conducted in the NACA 4 by 6-foot vertical wind tunnel to determine the air loads on an NACA 23012 airfoil in combination with a 25.66-percent-chord slotted flap and a 20-percent-chord split flap. Pressures were measured on both the upper and the lower surfaces of the main airfoil and the flaps for several angles of attack and at several flap settings. The data, presented as pressure diagrams and as graphs of the section coefficients for the flap alone and for the airfoil-flap combinations, are applicable to rib and flap design for a combination of a thick airfoil and a slotted or a split flap. The results of previous tests of a NACA 23012 airfoil with a slotted flap are compared with the present results.

  13. Direct numerical simulations of curvature effects on shear layer transition over airfoils

    NASA Astrophysics Data System (ADS)

    Zhang, Wei; Cheng, Wan; Qamar, Adnan; Gao, Wei; Samtaney, Ravi

    2013-11-01

    Shear layer transition and subsequent turbulent flow development over the leeward section of airfoils are affected by the surface curvature in terms of its associated effects, such as laminar flow separation, adverse pressure gradient, and the interactions between separated flow and wake vortices, etc. We present direct numerical simulations (DNS) of shear layer transitions over two airfoils, NACA 4412 and NACA 0012-64, at 10 deg. angle of attack, and Rec = 104 based on uniform inflow velocity and chord length. The two airfoils chosen are geometrically almost the same with identical maximum thickness along with chordwise position but different cambers and hence different curvature. The curvature effects on the flow are presented by the unsteady evolution patterns of laminar flow separation; shear layer detachment, breakdown to turbulence, turbulent boundary layer reattachment and vortex shedding, and quantitative results on the development of turbulent boundary layer are emphasized. This DNS database is generated with an energy conservative fourth-order incompressible Navier-Stokes code with O(109) mesh points. Supported by a KAUST funded project on large eddy simulation of turbulent flows. The IBM Blue Gene P Shaheen at KAUST was utilized for the simulations.

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

    NASA Astrophysics Data System (ADS)

    Murphy, Jeffery T.; Hu, Hui

    2010-08-01

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

  15. 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.

  16. Optimization of natural laminar flow airfoils for high section lift-to-drag ratios in the lower Reynolds number range

    NASA Technical Reports Server (NTRS)

    Pfenninger, Werner; Vemuru, Chandra S.

    1989-01-01

    Relatively thin natural-laminar-flow airfoils were arranged optimally for different design lift coefficients in the wing chord Reynolds number ranges of 200,000-600,00 and 0.875 x 10 to the 6th to 2 x 10 to the 6th. The 9.5 percent thick airfoil ASM-LRN-010, the 7.9 percent thick airfoil ASM-LRN-012, the 10.4 percent thick airfoil ASM-LRN-015, and the 8.2 percent thick airfoil ASM-LRN-017 were designed for high lift-to-drag ratios using Drela's design and analysis.

  17. Experimental studies of the Eppler 61 airfoil at low Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Burns, T. F.; Mueller, T. J.

    1982-01-01

    The results of an experimental study to document the effects of separation and transition on the performance of an airfoil designed for low Reynolds number operation are presented. Lift, drag and flow visualization data were obtained for the Eppler 61 airfoil section for chord Reynolds numbers from about 30,000 to over 200,000. Smoke flow visualization was employed to document the boundary layer behavior and was correlated with the Eppler airfoil design and analysis computer program. Laminar separation, transition and turbulent reattachment had significant effects on the performance of this airfoil.

  18. An experimental study of transonic flow about a supercritical airfoil. Static pressure and drag data obtained from tests of a supercritical airfoil and an NACA 0012 airfoil at transonic speeds, supplement

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    Surface static-pressure and drag data obtained from tests of two slightly modified versions of the original NASA Whitcomb airfoil and a model of the NACA 0012 airfoil section are presented. Data for the supercritical airfoil were obtained for a free-stream Mach number range of 0.5 to 0.9, and a chord Reynolds number range of 2 x 10 to the 6th power to 4 x 10 to the 6th power. The NACA 0012 airfoil was tested at a constant chord Reynolds number of 2 x 10 to the 6th power and a free-stream Mach number range of 0.6 to 0.8.

  19. Numerical analyses of passive and active flow control over a micro air vehicle with an optimized airfoil

    NASA Astrophysics Data System (ADS)

    Gada, Komal Kantilal

    Numerical investigations of an optimized thin airfoil with a passive and an active flow control device (riblets and rotary cylinder) have been performed. The objectives of the thesis were to investigate the tip vortices reduction using riblets and decrease in flow separation, using a rotary cylinder for improved lift-to-drag ratio. The investigations has application potentials in improving performances of Micro Air Vehicles (MAVs). The airfoil has a chord length of 19.66 cm and a span of 25 cm. with the free stream mean velocity was set at 20 m/s. The Reynolds number was calculated as 3 x 10 4. Investigations with base model of the airfoil have shown flow separation at approximately 85% chord length at an angle of attack of 17 degrees. For investigation using passive flow control device, i.e. riblets, investigations were performed for different radial sizes but at a fixed location. It was found that with 1 mm radial size riblet, the tip vortices were reduced by approximately 95%, as compared to the baseline model. Although negligible lift-to-drag improvement was seen, a faster dissipation rate in turbulent kinetic energy was observed. Furthermore, investigations were carried out using the active flow control device. The rotary cylinder with a 0.51 cm in diameter was placed slightly downstream of the location of flow separation, i.e. at x/c = 0.848. Investigations were performed at different cylinder's rotations, corresponding to different tangential velocities of being higher than, equal to and less than the free stream mean velocity. Results have shown approximately 10% improvement in lift to drag ratio when the tangential velocity is near the free stream mean velocity. Further investigation may include usage of the riblets and the rotary cylinder combined, to increase the stability as well as the lift-to-drag ratio of the MAVs.

  20. N1 component reflects difference of terminal chords in three-chord sequences.

    PubMed

    Dekio-Hotta, Shoko; Kojima, Tomoyuki; Karino, Shotaro; Yamasoba, Tatsuya; Dekio, Itaru; Ito, Yoichi M; Satake, Hiroyuki; Kaga, Kimitaka

    2009-02-18

    We understand from experience that musical contexts are formed when chords are combined according to the rules of harmony. In this study, the N1 component of the auditory-evoked potential was measured using comparable three-chord sequences; these sequences were constructed as a consecutive task (cf. C-C-C vs. Cm-Cm-Cm; control) and a cadence task (cf. C-G-C vs. C-G-Cm). In the cadence task, compared with cadences ending with a major chord (anticipated chord), those ending with a minor chord (unanticipated chord) showed a significantly larger amplitude of N1 waves. These components of auditory-evoked potentials reflect the effect of chord progression in musical perception and suggest that the musical context is recognized at least 100 ms after a chord is played. PMID:19188861

  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. Influence of self-adaptive hairy flaps on the stall delay of an airfoil in ramp-up motion

    NASA Astrophysics Data System (ADS)

    Brücker, Christoph; Weidner, Christoph

    2014-05-01

    It is known in the case of some birds that the coverts on the upper side of their wings pop-up under critical flight conditions such as the landing approach, thus acting like a brake on the spread of flow separation. Taking experimental investigations as its basis, this paper deals with the influence of various configurations of self-adaptable hairy flaplets located on the lower half of the wing and with chord-length c (dense rows of slender elastomeric flaps, L=0.05c, 0.1c, 0.2c) on the flow around an NACA0020 airfoil at low Reynolds number flow (Re=77×103). Flow evolution along the airfoil when in ramp-up motion (α0=0, αs=20°, reduced frequency k=0.12) was measured with and without hairy flaps, with growth in the chord-normal thickness of the separation region above the airfoil investigated in order to determine stall onset time Ts. Whereas small flaps with L=0.05c do not change the overall stall process, it was possible to use configurations with L=0.1c (double-row, triple-row configuration) to delay stall onset Ts by a factor of around 2-4 when compared with the clean airfoil. The motion of the flaps and the flow field were measured simultaneously at high temporal resolution using high-speed PIV. Correlation between flap motion and velocity distribution showed that backflow induced by vortex structures is indeed prevented by the hairy flaps. A significant difference was identified in the shear-layer roll-up process, which was almost regular and locked with the fundamental frequency on the covered airfoil with no signs of non-linear growth over longer periods. By way of contrast, in the case of the clean airfoil the early merging of the shear-layer vortices and a rapid increase in the thickness of the separation region were observed. It is therefore concluded that mode locking is achieved between flap rows with an interspacing of 0.15c-0.2c, while the fundamental shear-layer roll-up wavelength measured (λ0≈0.15c-0.2c) indicates the relevance of flap row

  3. A Generalization of the Parabolic Chord Property

    ERIC Educational Resources Information Center

    Mason, John

    2011-01-01

    The well known property of quadratic functions, that the tangents at either end of a chord of a parabola meet in a point aligned vertically with the midpoint of the chord is extended to polynomials of degree d. Given two distinct points on a polynomial of degree d, the Taylor polynomials of degree d - 1 at those points meet in d - 1 points whose…

  4. Japanese monkeys perceive sensory consonance of chords.

    PubMed

    Izumi, A

    2000-12-01

    Consonance/dissonance affects human perception of chords from early stages of development [e.g., Schellenberg and Trainor, J. Acoust. Soc. Am. 100, 3321-3328 (1996)]. To examine whether consonance has some role in audition of nonhumans, three Japanese monkeys (Macaca fuscata) were trained to discriminate simultaneous two-tone complexes (chords). The task was serial discrimination (AX procedure) with repetitive presentation of background stimuli. Each tone in a chord was comprised of six harmonics, and chords with complex ratios of fundamental frequency (e.g., frequency ratio of 8:15 in major seventh) resulted in dissonance. The chords were transposed for each presentation to make monkeys attend to cues other than the absolute frequency of a component tone. Monkeys were initially trained to detect changes from consonant (octave) to dissonant (major seventh). Following the successful acquisition of the task, transfer tests with novel chords were conducted. In these transfer tests, the performances with detecting changes from consonant to dissonant chords (perfect fifth to major seventh; perfect fourth to major seventh) were better than those with detecting reverse changes. These results suggested that the consonance of chords affected the performances of monkeys. PMID:11144600

  5. Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Viken, J. K.; Pfenninger, W.; Beasley, W. D.; Harvey, W. D.

    1984-01-01

    Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.

  6. 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.

  7. Statistical error in a chord estimator of correlation dimension: The rule of five''

    SciTech Connect

    Theiler, J. ); Lookman, T. . Dept. of Applied Mathematics)

    1992-09-01

    The statistical precision of a chord method for estimating dimension from a correlation integral is derived. The optimal chord length is determined, and a comparison is made to other estimators. The simple chord estimator is only 25% less precise than the optimal estimator which uses the full resolution and full range of the correlation integral. The analytic calculations are based on the hypothesis that all pairwise distances between the points in the embedding space are statistically independent. The adequacy of this approximation is assessed numerically, and a surprising result is observed in which dimension estimators can be anomalously precise for sets with reasonably uniform (nonfractal) distributions.

  8. Status of NASA advanced LFC airfoil high-lift study

    NASA Technical Reports Server (NTRS)

    Applin, Z. T.

    1982-01-01

    The design of a high lift system for the NASA advanced LFC airfoil designed by Pfenninger is described. The high lift system consists of both leading and trailing edge flaps. A 3 meter semispan, 1 meter chord wing model using the above airfoil and high lift system is under construction and will be tested in the NASA Langley 4 by 7 meter tunnel. This model will have two separate full span leading edge flaps (0.10c and 0.12c) and one full span trailing edge flap (0.25c). The performance of this high lift system was predicted by the NASA two dimensional viscous multicomponent airfoil program. This program was also used to predict the characteristics of the LFC airfoils developed by the Douglas Aircraft Company and Lockheed-Georgia Aircraft Company.

  9. The structure of separated flow regions occurring near the leading edge of airfoils - including transition

    NASA Technical Reports Server (NTRS)

    1989-01-01

    Laser Doppler Velocimeter data, static pressure data, and smoke flow visualization data was obtained and analyzed to correlate with separation bubble data. The Eppler 387 airfoil was focused on at a chord Reynolds number of 100,000 and an angle of attack of 2 deg. Additional data was also obtained from the NACA 663-018 airfoil at a chord Reynolds number of 160,000 and an angle of attack of 12 deg. The structure and behavior of the transition separation bubble was documented along with the redeveloping boundary layer after reattachment over an airfoil at low Reynolds numbers. The understanding of the complex flow phenomena was examined so that analytic methods for predicting their formation and development can be improved. These analytic techniques have applications in the design and performance prediction of airfoils operating in the low Reynolds number flight regime.

  10. 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.

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

    NASA Technical Reports Server (NTRS)

    Mueller, T. J.

    1984-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Adair, D.; Horne, W. C.

    1993-07-01

    Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analysed. 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% of flap chord. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 × 106 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.

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

    NASA Astrophysics Data System (ADS)

    Adair, D.; Horne, W. C.

    1989-09-01

    Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analysed. 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% of flap chord. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 × 106 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.

  14. 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.

  15. Comparison of NACA 6-series and 4-digit airfoils for Darrieus wind turbines

    SciTech Connect

    Migliore, P.G.

    1983-07-01

    The aerodynamic efficiency of Darrieus wind turbines as effected by blade airfoil geometry was investigated. Analysis was limited to curved-bladed machines having rotor solidities of 7 to 21% and operating at a Reynolds number of 3 X 10/sup 6/, Ten different airfoils, having thickness-to-chord ratios of 12, 15, and 18%, were studied. Performance estimates were made using a blade element/momentum theory approach. Results indicated that NACA 6-series airfoils yeild peak power coefficients as great as NACA 4-digit airfoils and have broader and flatter power coefficient-tip speed ratio curves. Sample calculations for an NACA 63/sub 2/-015 airfoil showed an annual energy output increase of 17-27%, depending on rotor solidity, compared to an NACA 0015 airfoil.

  16. Comparison of NACA 6-series and 4-digit airfoils for Darrieus wind turbines

    NASA Astrophysics Data System (ADS)

    Migliore, P. G.

    1983-08-01

    The aerodynamic efficiency of Darrieus wind turbines as effected by blade airfoil geometry was investigated. Analysis was limited to curved-bladed machines having rotor solidities of 7-21 percent and operating at a Reynolds number of 3 x 10 to the 6th. Ten different airfoils, having thickness-to-chord ratios of 12, 15, and 18 percent, were studied. Performance estimates were made using a blade element/momentum theory approach. Results indicated that NACA 6-series airfoils yield peak power coefficients as great as NACA 4-digit airfoils and have broader and flatter power coefficient-tip speed ratio curves. Sample calculations for an NACA 63(2)-015 airfoil showed an annual energy output increase of 17-27 percent, depending on rotor solidity, compared to an NACA 0015 airfoil.

  17. Calculated Low-Speed Steady and Time-Dependent Aerodynamic Derivatives for Some Airfoils Using a Discrete Vortex Method

    NASA Technical Reports Server (NTRS)

    Riley, Donald C.

    2015-01-01

    This paper contains a collection of some results of four individual studies presenting calculated numerical values for airfoil aerodynamic stability derivatives in unseparated inviscid incompressible flow due separately to angle-of-attack, pitch rate, flap deflection, and airfoil camber using a discrete vortex method. Both steady conditions and oscillatory motion were considered. Variables include the number of vortices representing the airfoil, the pitch axis / moment center chordwise location, flap chord to airfoil chord ratio, and circular or parabolic arc camber. Comparisons with some experimental and other theoretical information are included. The calculated aerodynamic numerical results obtained using a limited number of vortices provided in each study compared favorably with thin airfoil theory predictions. Of particular interest are those aerodynamic results calculated herein (such as induced drag) that are not readily available elsewhere.

  18. Reynolds number, thickness and camber effects on flapping airfoil propulsion

    NASA Astrophysics Data System (ADS)

    Ashraf, M. A.; Young, J.; Lai, J. C. S.

    2011-02-01

    The effect of varying airfoil thickness and camber on plunging and combined pitching and plunging airfoil propulsion at Reynolds number Re=200, 2000, 20 000 and 2×106 was studied by numerical simulations for fully laminar and fully turbulent flow regimes. The thickness study was performed on 2-D NACA symmetric airfoils with 6-50% thick sections undergoing pure plunging motion at reduced frequency k=2 and amplitudes h=0.25 and 0.5, and for combined pitching and plunging motion at k=2, h=0.5, phase ϕ=90°, pitch angle θo=15° and 30° and the pitch axis was located at 1/3 of chord from leading edge. At Re=200 for motions where positive thrust is generated, thin airfoils outperform thick airfoils. At higher Re significant gains could be achieved both in thrust generation and propulsive efficiency by using a thicker airfoil section for plunging and combined motion with low pitch amplitude. The camber study was performed on 2-D NACA airfoils with varying camber locations undergoing pure plunging motion at k=2, h=0.5 and Re=20 000. Little variation in thrust performance was found with camber. The underlying physics behind the alteration in propulsive performance between low and high Reynolds numbers has been explored by comparing viscous Navier-Stokes and inviscid panel method results. The role of leading edge vortices was found to be key to the observed performance variation.

  19. 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.

  20. 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.

  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. Chord, Tie Bar & Crossbracing Joint Detail in Plan; Crossbracing ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Chord, Tie Bar & Crossbracing Joint Detail in Plan; Crossbracing Center Joint Detail in Plan; Chord Joining Detail in Plan & Elevation; Chord, Panel Post, Tie Bar, & Diagonal Brace Joint Detail; Crossbracing Center Joint Detail in Section; Chord, Panel Post, Tie Bar & Horizontal Brace Joint Detail - Narrows Bridge, Spanning Sugar Creek at Old County Road 280 East, Marshall, Parke County, IN

  3. Redesigning a Film-Cooled Airfoil Trailing Edge using MRI Techniques

    NASA Astrophysics Data System (ADS)

    Benson, Michael; Elkins, Christopher; Eaton, John

    2011-11-01

    Trailing edges of modern gas turbine blades are film cooled through cutback slots on the airfoil pressure surface. The slots are spanwise divided, forming rectangular wall jets separated by tapered lands. The 3D wall jets mix rapidly with the mainstream flow reducing the cooling effectiveness. Experiments were conducted to document the 3D mean velocity and coolant concentration fields on a baseline configuration using Magnetic Resonance Imaging (MRI) in a water flow with Re = 110,000 based on airfoil chord length. Critical flow features causing rapid mixing were identified: a separation bubble behind the slot lip, and a pair of strong longitudinal vortices formed just downstream of the slot breakout. The geometry was modified to improve film cooling surface effectiveness obtained from the concentration field. The first redesign modified the slot lip and land shapes to minimize the slot lip separation bubble size and reduce 3D effects. The other redesigns modified the land shape to reduce the strength of the longitudinal vortices. These latter two designs produced a substantial reduction in the mixing rate of the coolant jet with the mainstream flow, improving the cooling system performance. The highly detailed concentration and velocity fields available with MRI-based experiments can be used to understand the flow physics and derive significant system improvements. This work was generously supported by GE Aviation under the GE-USA program and the Army Research Office.

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

    NASA Technical Reports Server (NTRS)

    Vandam, C. P.

    1984-01-01

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

  5. High Reynolds Number Configuration Development of a High-Lift Airfoil

    NASA Technical Reports Server (NTRS)

    Valarezo, Walter O.; Dominik, Chet J.; Mcghee, Robert J.; Goodman, Wesley L.

    1993-01-01

    An experimental program has been conducted to assess performance of a transport multielement airfoil at flight Reynolds numbers. The studies were performed at chord Reynolds numbers as high as 16 million in the NASA Langley Low Turbulence Pressure Tunnel. Sidewall boundary-layer control to enforce flow two dimensionality was provided via an endplate suction system. The basic airfoil was an 11.55 percent thick supercritical airfoil representative of the stall critical station of a new-generation transport aircraft wing. The multielement airfoil was configured as a three-element airfoil with slat and flap chord ratios of 14.48 percent and 30 percent respectively. Testing focused on the development of landing configurations with high maximum lift capability and the assessment of Reynolds and Mach number effects. Also assessed were high-lift performance effects due to devices such as drooped spoilers and trailing-edge wedges. The present experimental studies revealed significant effects on high-lift airfoil performance due to Reynolds and Mach number variations and favorable lift increments at approach angles of attack due to the use of drooped spoilers or trailing-edge wedges. However, no substantial improvements in maximum lift capability were identified. A recently developed high performance single-segment flap was also tested and results indicated considerable improvements in lift and drag performance over existing airfoils. Additionally, it was found that this new flap shape at its optimum rigging was less sensitive to Reynolds number variations than previous designs.

  6. Theoretical and Experimental Data for a Number of NACA 6A-Series Airfoil Sections

    NASA Technical Reports Server (NTRS)

    Loftin, Laurence K., Jr.

    1946-01-01

    The NACA 6A-series airfoil sections were designed to eliminate the trailing-edge cusp which is characteristic of the NACA 6-series sections. Theoretical data are presented for NACA 6A-series basic thickness forms having the position of minimum pressure at 30-, 40-, and 50-percent chord and with thickness ratios varying from 6 percent to 15 percent. Also presented are data for a mean line designed to maintain straight sides on the cambered sections. The experimental results of a two dimensional wind tunnel investigation of the aerodynamic characteristics of five NACA 64A-series airfoil sections and two NACA 63A-series airfoil sections are presented. An analysis of these results, which were obtained at Reynolds numbers of 3 x 10(exp 6), 6 x 10(exp 6), and 9 x 10(exp 6), indicates that the section minimum drag and maximum lift characteristics of comparable NACA 6-series and 6A-series airfoil sections are essentially the same. The quarter-chord pitching-moment coefficients and angles of zero lift of NACA 6A-series airfoil sections are slightly more negative than those of corresponding NACA 6-series airfoil sections. The position of the aerodynamic center and the lift-curve slope of smooth NACA 6-series sections. The addition of standard leading-edge roughness causes the lift-curve slope of the newer sections to decrease with increasing airfoil thickness ratio.

  7. Effects of long-chord acoustically treated stator vanes on fan noise. 2: Effect of acoustical treatment

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Scott, J. N.; Leonard, B. R.; Stakolich, E. G.

    1976-01-01

    A set of long chord stator vanes was designed to replace the vanes in an existing fan stage. The long chord stator vanes consisted of a turning section and axial extension pieces, all of which incorporated acoustic damping material. The long chord stator vanes were tested in two lengths, with the long version giving more noise reduction than the short, primarily because of the additional lining material. The noise reduction achieved with the acoustically treated long chord stator vanes was compared with the reduction achieved by an acoustically treated exhaust splitter. The long chord stator was at least as good as the splitter as a method for incorporating acoustic lining material. In addition, comparing an acoustic three ring inlet and an acoustic wall-only inlet discloses that the wall-only inlet could be used in an engine where the noise reduction requirements are not too stringent.

  8. Transonic airfoil design code

    NASA Technical Reports Server (NTRS)

    Bauer, F.; Garabedian, P.; Korn, D.

    1980-01-01

    Program aids in design of shockless airfoils, assists development of fuel-conserving, supercritical wings. Algorithm calculates approximate airfoil shape given prescribed pressure distribution. This allows design of families of transonic airfoils for use in aircraft wings or turbine and compressor blades. Program is written in FORTRAN IV for batch execution on CDC-6000.

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

    NASA Technical Reports Server (NTRS)

    Crane, Robert M; Holtzclaw, Ralph W

    1944-01-01

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

  10. Optimization of an Advanced Design Three-Element Airfoil at High Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Lin, John C.; Dominik, Chet J.

    1995-01-01

    New high-lift components have been designed for a three-element advanced high-lift research airfoil using a state-of-the-art computational method. The new components were designed with the aim to provide high maximum-lift values while maintaining attached flow on the single-segment flap at approach conditions. This three-element airfoil has been tested in the NASA Langley Low-Turbulence Pressure Tunnel at chord Reynolds number up to 16 million. The performance of the NASA research airfoil is compared to a reference advanced high-lift research airfoil. Effects of Reynolds number on slat and flap rigging have been studied experimentally. The performance trend of this new high-lift design is comparable to that predicted by the computational method over much of the angle of attack range. Nevertheless, the method did not accurately predict the airfoil performance or the configuration-based trends near maximum lift.

  11. Lift enhancement of an airfoil using a Gurney flap and vortex generators

    NASA Technical Reports Server (NTRS)

    Storms, Bruce L.; Jang, Cory S.

    1993-01-01

    The results of a low-speed wind tunnel test are presented for a single-element airfoil incorporating two lift-enhancing devices, namely a Gurney flap and vortex generators. The former consists of a small plate, on the order of one to two percent of the airfoil chord in height, located at the trailing edge perpendicular to the pressure side of the airfoil. The later consist of commercially-available, wishbone-shaped vortex generators. The test was conducted in the NASA Ames 7- by 10-foot Wind Tunnel with a full-span NACA 4412 airfoil. Measurements of surface pressure distributions and wake profiles were made to determine the lift, drag, and pitching-moment coefficients for the various airfoil configurations. The results indicate that the addition of a Gurney flap increased the maximum lift coefficient from 1.49 up to 1.96.

  12. Numerical simulation of the process of airfoil icing in the presence of large supercooled water drops

    NASA Astrophysics Data System (ADS)

    Prykhodko, O. A.; Alekseyenko, S. V.

    2014-10-01

    We have developed a software package and related methodology that can be used to simulate the process of airfoil icing during flight in the presence of large supercooled liquid water drops in the oncoming airflow. The motion of a carrier medium is described using the Navier-Stokes equations for a compressible gas. The motion of water drops is described using an inertial model. The process of water deposition and its subsequent freezing on an airfoil surface are described by the method of control volumes based on the equations of conservation of mass, momentum, and energy for each element of the surface. The main results of simulations are presented for the icing of an NACA 0012 airfoil profile with "barrier" ice formation in the absence and presence of heating of the leading edge. The influence of the ice-growth thickness and position on the airfoil chord on the pattern of airflow and aerodynamic characteristics of airfoil is analyzed.

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

  14. Computation of viscous transonic flow about a lifting airfoil

    NASA Technical Reports Server (NTRS)

    Walitt, L.; Liu, C. Y.

    1976-01-01

    The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.

  15. Response of a thin airfoil encountering strong density discontinuity

    SciTech Connect

    Marble, F.E.

    1993-12-01

    Airfoil theory for unsteady motion has been developed extensively assuming the undisturbed medium to be of uniform density, a restriction accurate for motion in the atmosphere. In some instances, notably for airfoil comprising fan, compressor and turbine blade rows, the undisturbed medium may carry density variations or ``spots``, resulting from non-uniformities in temperature or composition, of a size comparable to the blade chord. This condition exists for turbine blades, immediately downstream of the main burner of a gas turbine engine where the density fluctuations of the order of 50 percent may occur. Disturbances of a somewhat smaller magnitude arise from the ingestion of hot boundary layers into fans, and exhaust into hovercraft. Because these regions of non-uniform density convect with the moving medium, the airfoil experiences a time varying load and moment which the authors calculate.

  16. Perceptual integrality of major chord components.

    PubMed

    Acker, B E; Pastore, R E

    1996-07-01

    In the present study, an accuracy, rather than a reaction time, version of the Garner paradigm was used to evaluate the integrality or separability of major chord components. Tuned (prototype, or P) and mistuned (nonprototype, or NP) sets of root position C-major triads were constructed by holding the C constant in all stimuli and varying the E and G frequencies in 2- and 4-Hz steps. The P stimuli represent small systematic mistunings in the E and G notes relative to an equal-tempered C-major chord. The NP stimuli represent an equivalent range of frequency variation, but relative to a significantly out-of-tune C-major triad. In different experimental sessions, a same-different (AX) task was used to separately evaluate discrimination performance for the E and G frequencies as a function of whether the nontarget frequency (G or E) was fixed or varied in either a correlated or an orthogonal fashion (with the C frequency always held constant). Compared with a fixed baseline condition where only the target frequency changed, both chord components exhibited a significant redundancy gain in the correlated conditions and, to varying degrees, significant interference effects in the orthogonal condition, indicating that the chord components function largely in an integral fashion. Relative to the discrimination of G, discrimination of the E frequency was less influenced by variation in the nontarget (G) frequency, showing that attention, to some degree, could be selectively allocated to the E chord component. In addition, the results were consistent with previous findings that the functional prototype for the major chord category seems to act as a perceptual anchor, rather than as a magnet, and appears to be located in the physiologically defined area of just temperament, as opposed to the more experientially defined area of equal temperament. PMID:8710453

  17. On the Design of Lifting Airfoils with High Critical Mach Number Using Full Potential Theory

    NASA Astrophysics Data System (ADS)

    Kropinski, M. C. A.

    We wish to construct airfoils that have the highest free-stream Mach number for a given set of geometric constraints for which the flow is nowhere supersonic. Nonlifting airfoils that maximize the critical Mach number for a given cross-sectional area are known to possess long sonic segments at their critical speed. To construct lifting airfoils, we proceed under the conjecture that an airfoil with a high value of has the longest possible arc length of sonic velocity over its upper and lower surface. In Kropinski etal. (1995) the lifting problem was tackled in transonic small-disturbance theory. In this paper we numerically construct lifting airfoils with high using the full potential theory and we show that these airfoils have significantly higher than some standard airfoils. We also construct airfoils with higher values of the lift coefficient, by relaxing the speed constraint on the lower surface of the airfoil to have a value less than sonic.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G; Graham, Donald J

    1946-01-01

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

  20. High Reynolds number tests of the cast 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 2

    NASA Technical Reports Server (NTRS)

    Dress, D. A.; Stanewsky, E.; Mcguire, P. D.; Ray, E. J.

    1984-01-01

    Wind tunnel tests of an advanced technology airfoil, the CAST 10-2/DOA 2, were conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). This was the third of a series of tests conducted in a cooperative airfoil research program between the National Aeronautics and Space Administration and the Deutsche Forschungsund Versuchsanstalt fur Luft- und Raumfahrt e. V. For these tests, temperature was varied from 270 K to 110 K at pressures from 1.5 to 5.75 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 2 to 20 million. The aerodynamic data for the 7.62 cm chord airfoil model used in these tests is presented without analysis. Descriptions of the 0.3-m TCT, the airfoil model, the test instrumentation, and the testing procedures are included.

  1. A unique measurement technique to study laminar-separation bubble characteristics on an airfoil

    NASA Technical Reports Server (NTRS)

    Stack, J. P.; Mangalam, S. M.; Berry, S. A.

    1987-01-01

    A 'nonintrusive', multielement heat-transfer sensor was designed to study laminar-separation bubble characteristics on a NASA LRN (1)-1010 low-Reynolds number airfoil. The sensor consists of 30 individual nickel films, vacuum-deposited on a thin substrate (0.05 mm) that was bonded to the airfoil model with the sensor array placed streamwise on the airfoil upper surface. Experiments were conducted on a 15-cm chord model in the 50,000-300,000 chord Reynolds number range. Time history as well as spectral analysis of signals from surface film gauges were simultaneously obtained to determine the location of laminar separation and the subsequent behavior of the separated shear layer. In addition to the successful determination of laminar separation, a new phenomenon involving a large phase shift in dynamic shear stresses across the separation and reattachment points was observed.

  2. 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.

  3. Data from tests of a R4 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Jenkins, R. V.; Johnson, W. G., Jr.; Hill, A. S.; Mueller, R.; Redeker, G.

    1984-01-01

    Aerodynamic data for the DFVLR R4 airfoil are presented in both graphic and tabular form. The R4 was tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) at Mach number from 0.60 to 0.78 at angles of attack from -2.0 to 8.0 degrees. The airfoil was tested at Reynolds numbers of 4, 6, 10, 15, 30, and 40 million based on the 152.32 mm chord.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  6. 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.

  7. Design of the low-speed NLF(1)-0414F and the high-speed HSNLF(1)-0213 airfoils with high-lift systems

    NASA Technical Reports Server (NTRS)

    Viken, Jeffrey K.; Watson-Viken, Sally A.; Pfenninger, Werner; Morgan, Harry L., Jr.; Campbell, Richard L.

    1987-01-01

    The design and testing of Natural Laminar Flow (NLF) airfoils is examined. The NLF airfoil was designed for low speed, having a low profile drag at high chord Reynolds numbers. The success of the low speed NLF airfoil sparked interest in a high speed NLF airfoil applied to a single engine business jet with an unswept wing. Work was also conducted on the two dimensional flap design. The airfoil was decambered by removing the aft loading, however, high design Mach numbers are possible by increasing the aft loading and reducing the camber overall on the airfoil. This approach would also allow for flatter acceleration regions which are more stabilizing for cross flow disturbances. Sweep could then be used to increase the design Mach number to a higher value also. There would be some degradation of high lift by decambering the airfoil overall, and this aspect would have to be considered in a final design.

  8. Preliminary wind-tunnel investigation of an NACA 23012 airfoil with various arrangements of venetian-blind flaps

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Harris, Thomas A

    1940-01-01

    Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a large-chord NACA 23012 airfoil with several arrangements of venetian-blind flaps to determine the aerodynamic section characteristics as affected by the over-all flap chord, the chords of the slats used to form the flap, the slat spacing, the number of slats and the position of the flap with respect to the wing. Complete section data are given in the form of graphs for all the combinations tested.

  9. 13. UNDERSIDE OF THROUGHWAY SHOWING MAIN CHORDS, SUSPENSION EYEBAR PIN ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. UNDERSIDE OF THROUGHWAY SHOWING MAIN CHORDS, SUSPENSION EYE-BAR PIN CONNECTORS, LOWER CHORD EYEBARS AND LATERAL BRACING MEMBERS - Spruce Street Bridge, East Spruce Street, 500 Block, spanning Power Canal, Sault Ste. Marie, Chippewa County, MI

  10. 12. DETAIL VIEW OF BOTTOM CHORD CONNECTION AT THIRD PANAL ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. DETAIL VIEW OF BOTTOM CHORD CONNECTION AT THIRD PANAL POINT IN FROM ABUTMENT. NOTE THAT THE BOTTOM CHORD IS CONTINUOUS ACROSS THE CONNECTION - Poffenberger Road Bridge, Spanning Catoctin Creek, Middletown, Frederick County, MD

  11. 14. Detail, upper chord connection point on upstream side of ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    14. Detail, upper chord connection point on upstream side of truss, showing connection of upper chord, laced vertical compression member, strut, counters, and laterals. - Dry Creek Bridge, Spanning Dry Creek at Cook Road, Ione, Amador County, CA

  12. 24. PIN CONNECTION AT VERTICAL AND BOTTOM CHORD ON CAMELBACK ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    24. PIN CONNECTION AT VERTICAL AND BOTTOM CHORD ON CAMELBACK THROUGH TRUSS. VERTICAL AND BOTTOM CHORD MADE OF HAND-FORGED EYE BARS - New River Bridge, Spanning New River at State Route 623, Pembroke, Giles County, VA

  13. Chord, Horizontal Tie Bar & Crossbracing Joint Details; Crossbracing Center ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Chord, Horizontal Tie Bar & Crossbracing Joint Details; Crossbracing Center Joint Detail; Chord, Panel Posts, Braces & Counterbrace Joint Detail - Brownsville Covered Bridge, Spanning East Fork Whitewater River (moved to Eagle Creek Park, Indianapolis), Brownsville, Union County, IN

  14. An experimental investigation of flow surrounding an airfoil with a jet exhausting from the lower surface

    NASA Technical Reports Server (NTRS)

    Krothapalli, A.; Leopold, D.

    1982-01-01

    The aerodynamics of an airfoil with a rectangular jet issuing from the lower surface at seventy percent of the chord, with and without a ground plane was studied. Measurements include surface pressure on the airfoil and the total pressure profiles in the near wake. These measurements were made at jet to free stream velocity ratios ranging from 0.5 to 5.0. The measurements indicated a significant positive and negative pressure regions on the lower surface of the airfoil ahead and after the nozzle exit respectively. The extent and intensity of these regions increase with increase in velocity ratio for the range covered here. The upper surface pressure distribution with velocity ratio show no significant variation. The presence of the ground plane, for height h, greater than one chord seem to have little influence on the overall pressure distribution of the airfoil. The airfoil wake centerline moves up with velocity ratio as compared to that of the free airfoil (without the jet).

  15. Left auditory cortex specialization for vertical harmonic structure of chords.

    PubMed

    Passynkova, Natalia; Sander, Kerstin; Scheich, Henning

    2005-12-01

    The representation of consonant and dissonant chords in the auditory cortex was investigated using low-noise functional magnetic resonance imaging and different experimental paradigms to separate the effects of vertical harmony from those of other musical features. The results revealed higher activation by consonant compared with dissonant chords in the left posterior auditory cortex, suggesting contributions of mechanisms of encoding the acoustical chord structure rather than mechanisms based on sequential integration of chords. PMID:16597802

  16. 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)

  17. Determining the optimal smoothing length scale for actuator line models of wind turbine blades

    NASA Astrophysics Data System (ADS)

    Martinez, Luis; Meneveau, Charles

    2015-11-01

    The actuator line model (ALM) is a widely used tool for simulating wind turbines when performing Large-Eddy Simulations. The ALM uses a smearing kernel ηɛ = 1 /ɛ3π 3 / 2 exp (-r2 /ɛ2) , where r is the distance to an actuator point, and ɛ is the smoothing length scale which establishes the kernel width, to project the lift and drag forces onto the grid. In this work, we develop formulations to establish the optimum value of the smoothing length scale ɛ, based on physical arguments, instead of purely numerical constraints. This parameter has a very important role in the ALM, to provide a length scale, which may, for example, be related to the chord of the airfoil being studied. In the proposed approach, we compare features (such as vertical pressure gradient) of a potential flow solution for flow over a lifting surface with features of the solution of the Euler equations with a body force term. The potential flow solution over a lifting surface is used as a general representation of an airfoil. The method presented aims to minimize the difference between these features of the flow fields as a function of the smearing length scale (ɛ), in order to obtain the optimum value. This work is supported by NSF (IGERT and IIA-1243482) and computations use XSEDE resources.

  18. 7. View showing placement of timber deck placement on chord ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. View showing placement of timber deck placement on chord and built up construction of top chord and continuous construction through top panel points, eye bar construction on bottom chord - Bridge No. 2.4, Spanning Boiling Fork Creek at Railroad Milepost JC-2.4, Decherd, Franklin County, TN

  19. Chord Panel Post, Vertical X Bracing & Horizontal Tie Joint ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Chord Panel Post, Vertical X Bracing & Horizontal Tie Joint Detail; Chord Joining Block & Spacer Block Detail; Cross Bracing Joint Detail; Chord Panel Post Diagonal & Horizontal Tie Joint Detail - Jackson Covered Bridge, Spanning Sugar Creek, CR 775N (Changed from Spanning Sugar Creek), Bloomingdale, Parke County, IN

  20. 8. Comparison of construction of bottom and top chords and ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. Comparison of construction of bottom and top chords and pin connections, bottom chord second panel point, top chords showing third panel point. - Bridge No. 2.4, Spanning Boiling Fork Creek at Railroad Milepost JC-2.4, Decherd, Franklin County, TN

  1. Hodograph design of lifting airfoils with high critical mach numbers

    NASA Astrophysics Data System (ADS)

    Kropinski, M. C. A.; Schwendeman, D. W.; Cole, J. D.

    1995-05-01

    We wish to construct airfoils that have the highest free-stream Mach number M ∞ for a given set of geometric constraints for which the flow is nowhere supersonic. Nonlifting airfoils which maximize M ∞ for a given thickness ratio δ are known to possess long sonic segments at their critical speed. To construct lifting airfoils, we proceed under the conjecture that the optimal airfoil satisfying a given set of constraints is the one possessing the longest possible arc length of sonic velocity. A boundary-value problem is formulated in the hodograph plane using transonic small-disturbance theory whose solution determines an airfoil with long sonic arcs. For small lift coefficients, the hodograph domain covers two Riemann sheets and a finite-difference method is used to solve the boundary-value problem on this domain. A numerical integration of the solution around the boundary yields an airfoil shape, and three examples are discussed. The performance of these airfoils is compared with standard airfoils having the same lift coefficient and δ, and it is shown that the calculated airfoils have a 6% 10% increase in critical M ∞.

  2. Advanced natural laminar flow airfoil with high lift to drag ratio

    NASA Technical Reports Server (NTRS)

    Viken, Jeffrey K.; Pfenninger, Werner; Mcghee, Robert J.

    1986-01-01

    An experimental verification of a high performance natural laminar flow (NLF) airfoil for low speed and high Reynolds number applications was completed in the Langley Low Turbulence Pressure Tunnel (LTPT). Theoretical development allowed for the achievement of 0.70 chord laminar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence did not cause upstream movement of transition with increasing chord Reynolds number. With such a rearward pressure recovery, a concave type deceleration was implemented. Two-dimensional theoretical analysis indicated that a minimum profile drag coefficient of 0.0026 was possible with the desired laminar flow at the design condition. With the three-foot chord two-dimensional model constructed for the LTPT experiment, a minimum profile drag coefficient of 0.0027 was measured at c sub l = 0.41 and Re sub c = 10 x 10 to the 6th power. The low drag bucket was shifted over a considerably large c sub l range by the use of the 12.5 percent chord trailing edge flap. A two-dimensional lift to drag ratio (L/D) was 245. Surprisingly high c sub l max values were obtained for an airfoil of this type. A 0.20 chort split flap with 60 deg deflection was also implemented to verify the airfoil's lift capabilities. A maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million.

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

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Mineck, Raymond E.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Reissner, Eric

    1944-01-01

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

  5. The guitar chord-generating algorithm based on complex network

    NASA Astrophysics Data System (ADS)

    Ren, Tao; Wang, Yi-fan; Du, Dan; Liu, Miao-miao; Siddiqi, Awais

    2016-02-01

    This paper aims to generate chords for popular songs automatically based on complex network. Firstly, according to the characteristics of guitar tablature, six chord networks of popular songs by six pop singers are constructed and the properties of all networks are concluded. By analyzing the diverse chord networks, the accompaniment regulations and features are shown, with which the chords can be generated automatically. Secondly, in terms of the characteristics of popular songs, a two-tiered network containing a verse network and a chorus network is constructed. With this network, the verse and chorus can be composed respectively with the random walk algorithm. Thirdly, the musical motif is considered for generating chords, with which the bad chord progressions can be revised. This method can make the accompaniments sound more melodious. Finally, a popular song is chosen for generating chords and the new generated accompaniment sounds better than those done by the composers.

  6. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

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

    1945-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1945-01-01

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

  8. Drag Measurements at Transonic Speeds of NACA 65-009 Airfoils Mounted on a Freely Falling Body to Determine the Effects of Sweepback and Aspect Ratio

    NASA Technical Reports Server (NTRS)

    Mathews, Charles W.; Thompson, Jim Rogers

    1947-01-01

    Drag measurements at transonic speeds on rectangular airfoils and on airfoils swept back 450 are reported. These airfoils, which were mounted on cylindrical test bodies, are part of a series being tested in free drops from high altitude to determine the effect of variation of basic airfoil parameters on airfoil drag characteristics at transonic speeds. These rectangular and swept-back airfoils had the same span, airfoil section (NACA 65-009), and chord perpendicular to the leading edge. The tests were made to compare the drag of rectangular and sweptback airfoils at a higher aspect ratio than had been used in a similar comparison reported previously. The results showed that the drag of the swept-back airfoil was less than 0.15 that of the rectangular airfoil at a Mach number of 1.00 and less than 0.30 that of the rectangular airfoil at a Mach number of 1.17. A comparison of these swept-back airfoils with similar airfoils of lower aspect ratio previously tested by the same method indicated that in the investigated speed range reduction in aspect ratio results in increased drag. In the highest part of the investigated speed range, however, the drag coefficient of the high-aspect-ratio swept-back airfoils showed a tendency to approach that of the lower-aspect-ratio swept-back airfoils. A similar comparison for the rectangular airfoils showed that delay in the drag rise and a reduction in drag at supercritical speeds can be realIzed through reduction in aspect ratio. These results confirm those reported in NACA ACR No. L5J16.

  9. Inviscid analysis of two supercritical laminar-flow-control airfoils at design and off-design conditions

    NASA Technical Reports Server (NTRS)

    Allison, D. O.

    1983-01-01

    Inviscid transonic flow results are provided at design and off design conditions for two supercritical laminar flow control airfoils. The newer airfoil, with its lower suction requirements for full chord laminar flow, has a higher design Mach number, steeper pressure gradients, a more positive pressure level in the forward region of the lower surface, and a recovery to a less positive pressure at the trailing edge. The two dimensional design Mach numbers for the two airfoils are 0.755 and 0.730 at a common design lift coefficient of 0.60, and their thickness to chord ratios are 0.131 and 0.135, respectively. Off design shock formation characteristics are similar for the two airfoils over a range of Mach numbers between 0.6 and 0.8 and lift coefficients from 0.4 to 0.7. The newer airfoil is similar to the one used in a large chord swept model experiment designed for the Langley 8 Foot Transonic Pressure Tunnel.

  10. 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.

  11. 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.

  12. The Compressibility Burble and the Effect of Compressibility on Pressures and Forces Acting on a Airfoil

    NASA Technical Reports Server (NTRS)

    Stack, John; Lindsey, W F; Littell, Robert E

    1939-01-01

    Simultaneous air-flow photographs and pressure-distribution measurements were made of the NACA 4412 airfoil at high speeds to determine the physical nature of the compressibility burble. The tests were conducted in the NACA 24-inch high-speed wind tunnel. The flow photographs were obtained by the Schlieren method and the pressures were simultaneously measured for 54 stations in the 5-inch-chord airfoil by means of a multiple-tube manometer. Following the general program, a few measurements of total-pressure loss in the wake of the airfoil at high speeds were made to illustrate the magnitude of the losses involved and the extent of the disturbed region; and, finally, in order to relate this work to earlier force-test data, a force test of a 5-inch-chord NACA 4412 airfoil was made. The results show the general nature of the phenomenon known as the compressibility burble. The source of the increased drag is shown to be a compression shock that occurs on the airfoil as its speed approaches the speed of sound. Finally, it is indicated that considerable experimentation is needed in order to understand the phenomenon completely.

  13. Circulation Measurements About the Tip of an Airfoil During Flight Through a Gust

    NASA Technical Reports Server (NTRS)

    Kuethe, Arnold

    1939-01-01

    Measurements were made of the circulation about the rectangular tip of a short-span airfoil passing through an artificial gust of known velocity gradient. A Clark Y airfoil of 30-centimeter chord was mounted on a whirling arm and moved at a velocity of 29 meters per second over a vertical gust with a velocity of nearly 7 meters per second. Flow angles were measured with a hot-wire apparatus. The rate at which the lift at the tips of a wing entering a gust is realized was found to be in satisfactory agreement with that predicted on the basis of the two-dimensional theory of von Karman and Sears.

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

    SciTech Connect

    Dini, P.; Coiro, D.P.

    1997-12-31

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

  15. Aerodynamic Forces and Loadings on Symmetrical Circular-Arc Airfoils with Plain Leading-Edge and Plain Trailing-Edge Flaps

    NASA Technical Reports Server (NTRS)

    Cahill, Jones F; Underwood, William J; Nuber, Robert J; Cheesman, Gail A

    1953-01-01

    An investigation has been made in the Langley two-dimensional low-turbulence tunnel and in the Langley two-dimensional low-pressure tunnel of 6- and 10-percent-thick symmetrical circular-arc airfoil sections at low Mach numbers and several Reynolds numbers. The airfoils were equipped with 0.15-chord plain leading-edge flaps and 0.20-chord plan trailing-edge flaps. The section lift and pitching-moment characteristics were determined for both airfoils with the flaps deflected individually and in combination. The section drag characteristics were obtained for the 6-percent-thick airfoil with the flaps partly deflected as low-drag-control flaps and for airfoils with the flaps neutral. Surface pressures were measured on the 6-percent-thick airfoil section with the flaps deflected either individually or in appropriate combination to furnish flap load and hinge-moment data applicable to the structural design of the airfoil. A generalized method is developed that permits the determination of the chordwise pressure distribution over sharp-edge airfoils with plain leading-edge flaps and plain trailing-edge flaps of arbitrary size and deflection.

  16. 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.

  17. Prediction of the Effect of Vortex Generators on Airfoil Performance

    NASA Astrophysics Data System (ADS)

    Sørensen, Niels N.; Zahle, F.; Bak, C.; Vronsky, T.

    2014-06-01

    Vortex Generators (VGs) are widely used by the wind turbine industry, to control the flow over blade sections. The present work describes a computational fluid dynamic procedure that can handle a geometrical resolved VG on an airfoil section. After describing the method, it is applied to two different airfoils at a Reynolds number of 3 million, the FFA- W3-301 and FFA-W3-360, respectively. The computations are compared with wind tunnel measurements from the Stuttgart Laminar Wind Tunnel with respect to lift and drag variation as function of angle of attack. Even though the method does not exactly capture the measured performance, it can be used to compare different VG setups qualitatively with respect to chord- wise position, inter and intra-spacing and inclination of the VGs already in the design phase.

  18. Large eddy breakup devices as low Reynolds number airfoils

    NASA Technical Reports Server (NTRS)

    Anders, John B.

    1986-01-01

    Turbulent drag reduction downstream of large-eddy breakup (LEBU) devices is analyzed from the viewpoint of low-Reynolds number airfoil aerodynamics. It is argued that the variability of results between different research labs is primarily due to low Reynolds number 'phenomena' associated with unsteady separation/transition of the LEBU device boundary layer. LEBU drag reduction is shown to be an extremely sensitive function of device microgeometry at the low Reynolds numbers of all current investigations, and by analogy with conventional low-Reynolds number airfoil testing, the conclusion is drawn that the full potential for LEBU drag reduction must be explored at chord Reynolds numbers of 300,000 and above.

  19. 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.

  20. Numerical analysis of bio-inspired corrugated airfoil at low Reynolds number

    NASA Astrophysics Data System (ADS)

    Mondal, Partha Protim; Rahman, Md. Masudur; Hasan, A. B. M. Toufique

    2016-07-01

    A numerical study was conducted to investigate the aerodynamic performance of a bio-inspired corrugated airfoil at the chord Reynolds number of Rec=80,000 to explore the potential advantages of such airfoils at low Reynolds numbers. This study represents the transient nature of corrugated airfoils at low Reynolds number where flow is assumed to be laminar, unsteady, incompressible and two dimensional. The simulations include a sharp interface Cartesian grid based meshing employed with laminar viscous model. The flow field surrounding the corrugated airfoil has been analyzed using structured grid Finite Volume Method (FVM) based on Navier-Stokes equation. All parameters used in flow simulation are expressed in non-dimensional quantities for better understanding of flow behavior, regardless of dimensions or the fluid that is used. The simulated results revealed that the corrugated airfoil provides high lift with moderate drag and prevents large scale flow separation at higher angles of attack. This happens due to the negative shear drag produced by the recirculation zones which occurs in the valleys of the corrugated airfoils. The existence of small circulation bubbles sitting in the valleys prevents large scale flow separation thus increasing the aerodynamic performance of the corrugated airfoil.

  1. A two element laminar flow airfoil optimized for cruise. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Steen, Gregory Glen

    1994-01-01

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

  2. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

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

  4. 12. Detail, lower chord connection point on upstream side of ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. Detail, lower chord connection point on upstream side of truss, showing pinned connection of lower chord eye bars, laced vertical compression member, diagonal eye bar tension members, turnbuckled diagonal counters, and floor beam. Note also timber floor stringers supported by floor beam, and exposed ends of timber deck members visible at left above lower chord eye bar. View to northwest. - Red Bank Creek Bridge, Spanning Red Bank Creek at Rawson Road, Red Bluff, Tehama County, CA

  5. 15. Detail, lower chord connection point on downstream side, showing ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    15. Detail, lower chord connection point on downstream side, showing pinned connection of lower chord eye bars, laced vertical compression member, diagonal eye bar tension members, turnbuckled diagonal counters, and floor beam. Note also timber floor stringers supported by floor beam, and exposed ends of timber deck members visible at left above lower chord eye bar. View to northwest. - Dry Creek Bridge, Spanning Dry Creek at Cook Road, Ione, Amador County, CA

  6. The Development of Cambered Airfoil Sections Having Favorable Lift Characteristics at Supercritical Mach Numbers

    NASA Technical Reports Server (NTRS)

    Graham, Donald J

    1948-01-01

    Several groups of new airfoil sections, designated as the NACA 8-series, are derived analytically to have lift characteristics at supercritical Mach numbers which are favorable in the sense that the abrupt loss of lift, characteristic of the usual airfoil section at Mach numbers above the critical, is avoided. Aerodynamic characteristics determined, from two-dimensional windtunnel tests at Mach numbers up to approximately 0.9 are presented for each of the derived airfoils. Comparisons are made between the characteristics of these airfoils and the corresponding characteristics of representative NPiCA 6-series airfoils. The experimental results confirm the design expectations in demonstrating for the NACA S-series airfoils either no variation, or an Increase from the low-speed design value, In the lift coefficient at a constant angle of attack with increasing Mach number above the critical. It was not found possible to improve the variation with Mach number of the slope of the lift curve for these airfoils above that for the NACA 6-series airfoils. The drag characteristics of the new airfoils are somewhat inferior to those of the NACA 6- series with respect to divergence with Mach number, but the pitching-moment characteristics are more favorable for the thinner new sections In demonstrating somewhat smaller variations of moment coefficient with both angle of attack and Mach number. The effect on the aero&ynamic characteristics at high Mach numbers of removing the cusp from the trailing-edge regions of two 10-percent-chord-thick NACA 6-series airfoils is determined to be negligible.

  7. 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.

  8. Experimental investigation of the flowfield of an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Panda, J.; Zaman, K. B. M. Q.

    1992-01-01

    The flow field of an airfoil oscillated periodically over a wide range of reduced frequencies, 0 less than k less than 1.6, is studied experimentally at chord Reynolds numbers of R sub c = 22,000 and 44,000. The NACA0012 airfoil is pitched sinusoidally about one quarter chord between alpha of 5 deg and 25 deg. Detailed flow visualization and phase averaged vorticity measurements are carried out for k = 0.2 to document the evolution and the shedding of the dynamic stall vortex (DSV). In addition to the DSV, an intense vortex of opposite sign originates from the trailing edge just when the DSV is shed. After being shed into the wake, the two together take the shape of a large 'mushroom' while being convected away from the airfoil. The unsteady circulation around the airfoil and, therefore, the time varying component of the lift is estimated in a novel way from the shed vorticity flux and is found to be in good agreement with the lift variation reported by others. The delay in the shedding of the DSV with increasing k, as observed by previous researchers, is documented for the full range of k. The DSV, for example, is shed nearly at the maximum alpha of 25 deg at k = 0.2, but is shed at the minimum alpha of 5 deg at k = 0.8. At low k, the flowfield appears quasi-steady and the bluff body shedding corresponding to the maximum alpha (25 deg) dominates the unsteady fluctuations in the wake.

  9. Experimental investigation of the flowfield of an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Panda, J.; Zaman, K. B. M. Q.

    1992-01-01

    The flowfield of an airfoil oscillated periodically over a wide range of reduced frequencies, 0 less than or = k less than or = 1.6 is studied experimentally at chord Reynolds numbers of R sub c = 22,000 and 44,000. The NACA0012 airfoil is pitched sinusoidally about one quarter chord between angles of attack (alpha) of 5 and 25 degrees. Detailed flow visualization and phase averaged vorticity measurements are carried out for k = 0.2 to document the evolution and the shedding of the dynamic stall vortex (DSV). In addition to the DSV, an intense vortex of opposite sign originates from the trailing edge just when the DSV is shed. After being shed into the wake, the two together take the shape of a large 'mushroom' while being convected away from the airfoil. The unsteady circulation around the airfoil and, therefore, the time varying component of the lift is estimated in a novel way from the shed vorticity flux and is found to be in good agreement with the lift variation reported by others. The delay in the shedding of the DSV with increasing k, as observed by previous researchers, is documented for the full range of k. The DSV, for example, is shed nearly at the maximum alpha of 25 degrees at k = 0.2, but is shed at the minimum alpha of 5 degrees at k = 0.8. At low k, the flowfield appears quasi-steady and the bluff body shedding corresponding to the maximum alpha (25 degrees) dominates the unsteady fluctuations in the wake.

  10. Closed loop steam cooled airfoil

    SciTech Connect

    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.

  11. Computer Program to Obtain Ordinates for NACA Airfoils

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.; Brooks, Cuyler W., Jr.; Hill, Acquilla S.; Sproles, Darrell W.

    1996-01-01

    Computer programs to produce the ordinates for airfoils of any thickness, thickness distribution, or camber in the NACA airfoil series were developed in the early 1970's and are published as NASA TM X-3069 and TM X-3284. For analytic airfoils, the ordinates are exact. For the 6-series and all but the leading edge of the 6A-series airfoils, agreement between the ordinates obtained from the program and previously published ordinates is generally within 5 x 10(exp -5) chord. Since the publication of these programs, the use of personal computers and individual workstations has proliferated. This report describes a computer program that combines the capabilities of the previously published versions. This program is written in ANSI FORTRAN 77 and can be compiled to run on DOS, UNIX, and VMS based personal computers and workstations as well as mainframes. An effort was made to make all inputs to the program as simple as possible to use and to lead the user through the process by means of a menu.

  12. An experimental investigation of the low Reynolds number performance of the Lissaman 7769 airfoil

    NASA Technical Reports Server (NTRS)

    Conigliaro, P. E.

    1983-01-01

    A Lissaman 7769 airfoil, used on the Gossamer Condor and Gossamer Albatross human-powered aircraft, was tested in a low turbulence subsonic wind tunnel. Lift and drag data were collected at chord Reynolds numbers of 100,000, 150,000, 200,000, 250,000, and 300,000; at angles of attack from -10 to +20 deg by using an external strain gage force balance. Lift curves, drag curves, and drag polars were generated from both uncorrected data and data corrected for wind tunnel blockage effects. A flow visualization study was performed to correlate with the force data. The results of the investigation have shown that the airfoil exhibits a significant degradation in performance for chord Reynolds numbers below 150,000.

  13. Airfoil section characteristics as applied to the prediction of air forces and their distribution on wings

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Rhode, R V

    1938-01-01

    The results of previous reports dealing with airfoil section characteristics and span load distribution data are coordinated into a method for determining the air forces and their distribution on airplane wings. Formulas are given from which the resultant force distribution may be combined to find the wing aerodynamic center and pitching moment. The force distribution may also be resolved to determine the distribution of chord and beam components. The forces are resolved in such a manner that it is unnecessary to take the induced drag into account. An illustration of the method is given for a monoplane and a biplane for the conditions of steady flight and a sharp-edge gust. The force determination is completed by outlining a procedure for finding the distribution of load along the chord of airfoil sections.

  14. Some new airfoils

    NASA Technical Reports Server (NTRS)

    Eppler, R.

    1979-01-01

    A computer approach to the design and analysis of airfoils and some common problems concerning laminar separation bubbles at different lift coefficients are briefly discussed. Examples of application to ultralight airplanes, canards, and sailplanes with flaps are given.

  15. Turbine airfoil with controlled area cooling arrangement

    SciTech Connect

    Liang, George

    2010-04-27

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

  16. 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.

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

    NASA Astrophysics Data System (ADS)

    Adair, Desmond; Horne, W. Clifton

    1988-02-01

    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.

  18. Theoretical and experimental data for a number of NACA 6A-series airfoil sections

    NASA Technical Reports Server (NTRS)

    Loftin, Laurence K , Jr

    1948-01-01

    The NACA 6a-series airfoil sections were designed to eliminate the trailing-edge cusp which is characteristic of the NACA 6a-series sections. Theoretical data are presented for NACA 6a-series basic thickness forms having the position of minimum pressure of 30, 40, and 50 percent chord and with thickness ratios varying from 6 percent to 15 percent. Also presented are data for a mean line designed to maintain straight sides on the cambered sections.

  19. 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.

  20. Transesophageal echocardiography in NeoChord procedure

    PubMed Central

    Demetrio, Pittarello; Andrea, Colli; Gianclaudio, Falasco; Antonio, Marcassa; Gino, Gerosa; Carlo, Ori

    2015-01-01

    Background: Transapical off-pump mitral valve intervention with neochord implantation for degenerative mitral valve disease have been recently introduced in the surgical practice. The procedure is performed under 2D-3D transesophageal echocardiography guidance. Methods: The use of 3D real-time transesophageal echocardiography provides more accurate information than 2D echocardiography only in all the steps of the procedure. In particular 3D echocardiography is mandatory for preoperative assessment of the morphology of the valve, for correct positioning of the neochord on the diseased segment, for the final tensioning of the chordae and for the final evaluation of the surgical result. Result and Conclusion: This article is to outline the technical aspects of the transesophageal echocardiography guidance of the NeoChord procedure showing that the procedure can be performed only with a close and continuous interaction between the anesthesiologist and the cardiac surgeon. PMID:25849688

  1. To speak in chords about sexuality.

    PubMed

    Bermant, G

    1995-01-01

    Sexuality emerges from the interdependencies of biology, awareness, and the facts and artifacts of public life. A useful metaphor is that correct accountings of sexuality are not one-finger melodies--they are chords. Unfortunately, the physical vs. mental and nature vs. nurture controversies remain alive, well, and mischievous in regard to the correct understanding of human sexuality. Active political and legal disputes about homosexuality exemplify a continuing reliance on reductionistic models of the causes of conduct. Discourse relying on public misapprehension about biological causality can alter the course of subsequent science and public opinion and thus affect personal experience as well. Both dualistic and reductionistic models are traps and bar progress; the models should not be smuggled into accounts of sexuality. PMID:7630588

  2. R4 airfoil data corrected for sidewall boundary-layer effects in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Jenkins, R. V.

    1986-01-01

    Corrected aerodynamic data for the R4 airfoil at Mach numbers from 0.60 to 0.78 and angles of attack from -2.0 deg. to 4.5 deg. are presented. The test Reynolds numbers were 4 million, 6 million, 10 million, 15 million, 30 million, and 40 million based on the 152.32-mm chord of the airfoil. Corrections for data were previously published in NASA Technical Memorandum 85739. The design goal of a normal-force coefficient of 0.65 at a Mach number of 0.73 and a Reynolds number of 30 million was successfully obtained with this airfoil.

  3. 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.

  4. 16. Detail, lower chord connection point on downstream side at ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    16. Detail, lower chord connection point on downstream side at end panel showing lower chord eye bars, vertical tension eye bar, original and supplemental floor beams, turnbuckled lower laterals. View to northwest. - Dry Creek Bridge, Spanning Dry Creek at Cook Road, Ione, Amador County, CA

  5. 27. 100 foot through truss a typical lower chord ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    27. 100 foot through truss - a typical lower chord pin connection, located below the vertical member junction with the end post and upper chord. View shows one diagonal member. There are four of these per through truss for a total of 8, also shows the four inch conduit. - Weidemeyer Bridge, Spanning Thomes Creek at Rawson Road, Corning, Tehama County, CA

  6. Research on Chord Searching Algorithm Base on Cache Strategy

    NASA Astrophysics Data System (ADS)

    Jun, Guo; Chen, Chen

    How to improve search efficiency is a core problem in P2P network, Chord is a successful searching algorithm, but its lookup efficiency is lower because finger table has redundant information proposed the recently visited table and improved to gain more useful information in Chord. The simulation experiments show that approach can availably improve the routing efficiently.

  7. 13. Detail, upper chord connection point on upstream side of ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. Detail, upper chord connection point on upstream side of truss, showing connection of upper chord, laced vertical compression member, knee-braced strut, counters, and laterals. - Red Bank Creek Bridge, Spanning Red Bank Creek at Rawson Road, Red Bluff, Tehama County, CA

  8. Improved Techniques for Automatic Chord Recognition from Music Audio Signals

    ERIC Educational Resources Information Center

    Cho, Taemin

    2014-01-01

    This thesis is concerned with the development of techniques that facilitate the effective implementation of capable automatic chord transcription from music audio signals. Since chord transcriptions can capture many important aspects of music, they are useful for a wide variety of music applications and also useful for people who learn and perform…

  9. A study of high-lift airfoils at high Reynolds numbers in the Langley low-turbulence pressure tunnel

    NASA Technical Reports Server (NTRS)

    Morgan, Harry L., Jr.; Ferris, James C.; Mcghee, Robert J.

    1987-01-01

    An experimental study was conducted in the Langley Low Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of two supercritical type airfoils, one equipped with a conventional flap system and the other with an advanced high lift flap system. The conventional flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a small chord vane and a large chord aft flap. The advanced flap system consisted of a leading edge slat and a double slotted, trailing edge flap with a large chord vane and a small chord aft flap. Both models were tested with all elements nested to form the cruise airfoil and with the leading edge slat and with a single or double slotted, trailing edge flap deflected to form the high lift airfoils. The experimental tests were conducted through a Reynolds number range from 2.8 to 20.9 x 1,000,000 and a Mach number range from 0.10 to 0.35. Lift and pitching moment data were obtained. Summaries of the test results obtained are presented and comparisons are made between the observed aerodynamic performance trends for both models. The results showing the effect of leading edge frost and glaze ice formation is given.

  10. 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.

  11. 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.

  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. A frequency domain numerical method for airfoil broadband self-noise prediction

    NASA Astrophysics Data System (ADS)

    Zhou, Qidou; Joseph, Phillip

    2007-01-01

    This paper describes a numerical approach, based in the frequency domain, for predicting the broadband self-noise radiation due to an airfoil situated in a smooth mean flow. Noise is generated by the interaction between the boundary layer turbulence on the airfoil surface and the airfoil trailing edge. Thin airfoil theory is used to deduce the unsteady blade loading. In this paper, the important difference with much of the previous work dealing with trailing edge noise is that the integration of the surface sources for computation of the radiated sound field is evaluated on the actual airfoil surface rather than in the mean-chord plane. The assumption of flat plate geometry in the calculation of radiation is therefore avoided. Moreover, the solution is valid in both near and far fields and reduces to the analytic solution due to Amiet when the airfoil collapses to a flat plate with large span, and the measurement point is taken to the far field. Predictions of the airfoil broadband self-noise radiation presented here are shown to be in reasonable agreement with the predictions obtained using the Brooks approach, which are based on a comprehensive database of experimental data. Also investigated in this paper is the effect on the broadband noise prediction of relaxing the 'frozen-gust' assumption, whereby the turbulence at each frequency comprises a continuous spectrum of streamwise wavenumber components. It is shown that making the frozen gust assumption yields an under-prediction of the noise spectrum by approximately 2dB compared with that obtained when this assumption is relaxed, with the largest occurring at high frequencies. This paper concludes with a comparison of the broadband noise directivity for a flat-plat, a NACA 0012 and a NACA 0024 airfoil at non-zero angle of attack. Differences of up to 20 dB are predicted, with the largest difference occurring at a radiation angle of zero degrees relative to the airfoil mean centre line.

  15. 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.

  16. 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.

  17. 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.

  18. Prediction of high frequency gust response with airfoil thickness effects

    NASA Astrophysics Data System (ADS)

    Lysak, Peter D.; Capone, Dean E.; Jonson, Michael L.

    2013-05-01

    The unsteady lift forces that act on an airfoil in turbulent flow are an undesirable source of vibration and noise in many industrial applications. Methods to predict these forces have traditionally treated the airfoil as a flat plate. At higher frequencies, where the relevant turbulent length scales are comparable to the airfoil thickness, the flat plate approximation becomes invalid and results in overprediction of the unsteady force spectrum. This work provides an improved methodology for the prediction of the unsteady lift forces that accounts for the thickness of the airfoil. An analytical model was developed to calculate the response of the airfoil to high frequency gusts. The approach is based on a time-domain calculation with a sharp-edged gust and accounts for the distortion of the gust by the mean flow around the airfoil leading edge. The unsteady lift is calculated from a weighted integration of the gust vorticity, which makes the model relatively straightforward to implement and verify. For routine design calculations of turbulence-induced forces, a closed-form gust response thickness correction factor was developed for NACA 65 series airfoils.

  19. 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.

  20. 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.

  1. 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.

  2. An experimental study of the aerodynamics of a NACA 0012 airfoil with a simulated glaze ice accretion

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.

    1986-01-01

    An experimental study was conducted in the Ohio State University subsonic wind tunnel to measure the detailed aerodynamic characteristics of an airfoil with a simulated glaze ice accretion. A NACA 0012 model with interchangeable leading edges and pressure taps every one percent chord was used. Surface pressure and wake data were taken on the airfoil clean, with forced transition and with a simulated glaze ice shape. Lift and drag penalties due to the ice shape were found and the surface pressure clearly showed that large separation bubbles were present. Both total pressure and split-film probes were used to measure velocity profiles, both for the clean model and for the model with a simulated ice accretion. A large region of flow separation was seen in the velocity profiles and was correlated to the pressure measurements. Clean airfoil data were found to compare well to existing airfoil analysis methods.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  4. Design and experimental evaluation of a swept supercritical Laminar Flow Control (LFC) airfoil

    NASA Technical Reports Server (NTRS)

    Harvey, W. D.; Harris, C. D.; Brooks, C. W.; Clukey, P. G.; Stack, J. P.

    1986-01-01

    A large chord swept supercritical laminar flow control (LFC) airfoil was designed, constructed, and tested in the NASA Langley 8-ft Transonic Pressure Tunnel (TPT). The LFC airfoil experiment was established to provide basic information concerning the design and compatibility of high-performance supercritical airfoils with suction boundary layer control achieved through discrete fine slots or porous surface concepts. It was aimed at validating prediction techniques and establishing a technology base for future transport designs and drag reduction. Good agreement was obtained between measured and theoretically designed shockless pressure distributions. Suction laminarization was maintained over an extensive supercritical zone up to high Reynolds numbers before transition gradually moved forward. Full-chord laminar flow was maintained on the upper and lower surfaces at M sub infinity = 0.82 up to R sub c is less than or equal to 12 x 10 to the 6th power. When accounting for both the suction and wake drag, the total drag could be reducted by at least one-half of that for an equivalent turbulent airfoil. Specific objectives for the LFC experiment are given.

  5. Design and wind tunnel evaluation of a symmetric airfoil series for large wind turbine applications

    SciTech Connect

    Gregorek, G.M.

    1984-05-01

    Using the Eppler analytic design methodology, a family of symmetric airfoils with thickness to chord ratios of 0.21, 0.25, and 0.29 has been designed for applications to large wind turbines. The airfoil series, designed for Reynolds numbers above 4 x 10/sup 6/, has a contour that maintains its thickness to the 0.7 chord position, allowing use of a deep aft spar for structural integrity. Wind tunnel tests of the 21 percent thick airfoil conducted in The Ohio State University Airfoil Test Facilities at M = 0.25 and Re = 2, 4, and 8 x 10/sup 6/ showed good agreement with predicted surface pressures and lift and drag coefficients when the angle of attack was low and the Reynolds number was 8 x 10/sup 6/. At the lowest Reynolds number, both the measured lift curve slope and the maximum lift coefficient differ significantly from theory. In the post-stall region, agreement with theory and experiment is poor, at all Reynolds numbers, verifying the need for improved analytic methods in this important flow regime.

  6. High Reynolds number tests of the CAST 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 1

    NASA Technical Reports Server (NTRS)

    Dress, D. A.; Mcguire, P. D.; Stanewsky, E.; Ray, E. J.

    1983-01-01

    A wind tunnel investigation of an advanced technology airfoil, the CAST 10-2/DOA 2, was conducted in the Langley 0.3 meter Transonic Cryogenic Tunnel (0.3 m TCT). This was the first of a series of tests conducted in a cooperative National Aeronautics and Space Administration (NASA) and the Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e. V. (DFVLR) airfoil research program. Test temperature was varied from 280 K to 100 K to pressures from slightly above 1 to 5.8 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 4 x 10 to the 8th power to 45 x 10 to the 6th power. This report presents the experimental aerodynamic data obtained for the airfoil and includes descriptions of the airfoil model, the 0.3 m TCT, the test instrumentation, and the testing procedures.

  7. Exploratory investigation of the effectiveness of biplane wings with large-chord double slotted flaps in redirecting a propeller slipstream downward for vertical take-off

    NASA Technical Reports Server (NTRS)

    Kirby, Robert H

    1956-01-01

    Results are presented from static-force tests on biplane wings with large-chord double slotted flaps which turn the slipstream of a single counter rotating propeller downward for vertical take-off. The investigation provided information on the effect of chord length, flap deflection, propeller position, end plates, fuselage, and ground proximity on the efficiency of the wing system in turning the propeller slipstream.

  8. Aerodynamic data banks for Clark-Y, NACA 4-digit and NACA 16-series airfoil families

    NASA Technical Reports Server (NTRS)

    Korkan, K. D.; Camba, J., III; Morris, P. M.

    1986-01-01

    With the renewed interest in propellers as means of obtaining thrust and fuel efficiency in addition to the increased utilization of the computer, a significant amount of progress was made in the development of theoretical models to predict the performance of propeller systems. Inherent in the majority of the theoretical performance models to date is the need for airfoil data banks which provide lift, drag, and moment coefficient values as a function of Mach number, angle-of-attack, maximum thickness to chord ratio, and Reynolds number. Realizing the need for such data, a study was initiated to provide airfoil data banks for three commonly used airfoil families in propeller design and analysis. The families chosen consisted of the Clark-Y, NACA 16 series, and NACA 4 digit series airfoils. The various component of each computer code, the source of the data used to create the airfoil data bank, the limitations of each data bank, program listing, and a sample case with its associated input-output are described. Each airfoil data bank computer code was written to be used on the Amdahl Computer system, which is IBM compatible and uses Fortran.

  9. Parametric Investigation of a High-Lift Airfoil at High Reynolds Numbers

    NASA Technical Reports Server (NTRS)

    Lin, John C.; Dominik, Chet J.

    1997-01-01

    A new two-dimensional, three-element, advanced high-lift research airfoil has been tested in the NASA Langley Research Center s Low-Turbulence Pressure Tunnel at a chord Reynolds number up to 1.6 x 107. The components of this high-lift airfoil have been designed using a incompressible computational code (INS2D). The design was to provide high maximum-lift values while maintaining attached flow on the single-segment flap at landing conditions. The performance of the new NASA research airfoil is compared to a similar reference high-lift airfoil. On the new high-lift airfoil the effects of Reynolds number on slat and flap rigging have been studied experimentally, as well as the Mach number effects. The performance trend of the high-lift design is comparable to that predicted by INS2D over much of the angle-of-attack range. However, the code did not accurately predict the airfoil performance or the configuration-based trends near maximum lift where the compressibility effect could play a major role.

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

    NASA Astrophysics Data System (ADS)

    Grasso, Francesco

    2014-12-01

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

  11. 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.

  12. 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.

  13. Studies of friction drag and pressure drag of airfoils using the Eppler program

    NASA Technical Reports Server (NTRS)

    Phillips, William H.

    1988-01-01

    Most previous studies of the drag and two-dimensional airfoils consider only the total drag. The present report gives results of a study of three airfoils, using the Eppler program, to determine the distribution of friction drag along the chord and to obtain relative values of friction drag and pressure drag over a wide range of angle of attack and Reynolds number. The effects of boundary-layer suction in the turbulent region of the boundary layer of two of the airfoils are also investigated. The pressure drag is found to be an important component of the total drag, reaching values of 60 to 80 percent of the total drag near the stall. The use of suction producing a uniform inflow in the turbulent region of the boundary layer results in large increases in maximum lift, and increases the skin-friction drag but reduces or even changes the sign of the pressure drag.

  14. Status of advanced airfoil tests in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Ladson, C. L.; Ray, E. J.

    1984-01-01

    A joint NASA/U.S. industry program to test advanced technology airfoils in the Langley 0.3-meter Transonic Tunnel (TCT) was formulated under the Langley ACEE Project Office. The objectives include providing U.S. industry an opportunity to compare their most advanced airfoils to the latest NASA designs by means of high Reynolds number tests in the same facility. At the same time, industry would again experience in the design and construction of cryogenic test techniques. The status and details of the test program are presented. Typical aerodynamic results obtained, to date, are presented at chord Reynolds number up to 45 x 10(6) and are compared to results from other facilities and theory. Details of a joint agreement between NASA and the Deutsche Forschungs- und Versuchsantalt fur Luft- and Raumfahrt e.V. (DFVLR) for tests of two airfoils are also included. Results of these tests will be made available as soon as practical.

  15. Computation aspects and results of low-speed viscous flow about multicomponent airfoils

    NASA Astrophysics Data System (ADS)

    Oskam, B.

    1980-08-01

    The viscous flow about multicomponent airfoils was calculated solving incompressible potential flow and boundary layer problems iteratively. The presence of the shear layers is modeled in the potential flow by an outflow boundary condition on the airfoil surface and the wake centerline. Solutions show that high accuracy of the potential flow solution is required to justify the correction for viscous effects, especially near the wing trailing edge of a slotted configuration. Leading edge stall was studied. Turbulent boundary layer separation at 3% chord downstream of a laminar separation bubble is found on the nose of a NACA 63-009 airfoil at high Reynolds number indicating the occurence of turbulent leading edge stall. This phenomenon, being difficult to distinguish from laminar leading edge (short bubble) stall, occurs at such a small scale that it is observed only with high resolution results. The importance of the wing wake of a slotted configuration is examined.

  16. Unsteady two dimensional airloads acting on oscillating thin airfoils in subsonic ventilated wind tunnels

    NASA Technical Reports Server (NTRS)

    Fromme, J.; Golberg, M.

    1978-01-01

    The numerical calculation of unsteady two dimensional airloads which act upon thin airfoils in subsonic ventilated wind tunnels was studied. Neglecting certain quadrature errors, Bland's collocation method is rigorously proved to converge to the mathematically exact solution of Bland's integral equation, and a three way equivalence was established between collocation, Galerkin's method and least squares whenever the collocation points are chosen to be the nodes of the quadrature rule used for Galerkin's method. A computer program displayed convergence with respect to the number of pressure basis functions employed, and agreement with known special cases was demonstrated. Results are obtained for the combined effects of wind tunnel wall ventilation and wind tunnel depth to airfoil chord ratio, and for acoustic resonance between the airfoil and wind tunnel walls. A boundary condition is proposed for permeable walls through which mass flow rate is proportional to pressure jump.

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

    NASA Technical Reports Server (NTRS)

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

    1973-01-01

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

  18. Mild Dissonance Preferred Over Consonance in Single Chord Perception.

    PubMed

    Lahdelma, Imre; Eerola, Tuomas

    2016-05-01

    Previous research on harmony perception has mainly been concerned with horizontal aspects of harmony, turning less attention to how listeners perceive psychoacoustic qualities and emotions in single isolated chords. A recent study found mild dissonances to be more preferred than consonances in single chord perception, although the authors did not systematically vary register and consonance in their study; these omissions were explored here. An online empirical experiment was conducted where participants (N = 410) evaluated chords on the dimensions of Valence, Tension, Energy, Consonance, and Preference; 15 different chords were played with piano timbre across two octaves. The results suggest significant differences on all dimensions across chord types, and a strong correlation between perceived dissonance and tension. The register and inversions contributed to the evaluations significantly, nonmusicians distinguishing between triadic inversions similarly to musicians. The mildly dissonant minor ninth, major ninth, and minor seventh chords were rated highest for preference, regardless of musical sophistication. The role of theoretical explanations such as aggregate dyadic consonance, the inverted-U hypothesis, and psychoacoustic roughness, harmonicity, and sharpness will be discussed to account for the preference of mild dissonance over consonance in single chord perception. PMID:27433333

  19. Mild Dissonance Preferred Over Consonance in Single Chord Perception

    PubMed Central

    Eerola, Tuomas

    2016-01-01

    Previous research on harmony perception has mainly been concerned with horizontal aspects of harmony, turning less attention to how listeners perceive psychoacoustic qualities and emotions in single isolated chords. A recent study found mild dissonances to be more preferred than consonances in single chord perception, although the authors did not systematically vary register and consonance in their study; these omissions were explored here. An online empirical experiment was conducted where participants (N = 410) evaluated chords on the dimensions of Valence, Tension, Energy, Consonance, and Preference; 15 different chords were played with piano timbre across two octaves. The results suggest significant differences on all dimensions across chord types, and a strong correlation between perceived dissonance and tension. The register and inversions contributed to the evaluations significantly, nonmusicians distinguishing between triadic inversions similarly to musicians. The mildly dissonant minor ninth, major ninth, and minor seventh chords were rated highest for preference, regardless of musical sophistication. The role of theoretical explanations such as aggregate dyadic consonance, the inverted-U hypothesis, and psychoacoustic roughness, harmonicity, and sharpness will be discussed to account for the preference of mild dissonance over consonance in single chord perception. PMID:27433333

  20. Turbine airfoil film cooling

    NASA Astrophysics Data System (ADS)

    Hylton, Larry D.

    1986-10-01

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

  1. 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.

  2. A new airfoil design concept

    NASA Technical Reports Server (NTRS)

    Henne, P. A.; Gregg, R. D.

    1989-01-01

    The present airfoil design concept is based on utilizing unconventional geometry characteristics near the airfoil trailing edge which include a finite trailing edge thickness, strongly divergent trailing edge upper and lower surfaces, and high surface curvature on the lower surface at or near the lower surface trailing edge. This paper presents computational analyses of airfoils and a wing utilizing the concept, airfoil validation wind tunnel test results of several configurations, and wing-validation wind tunnel test results for a complete wing design. In addition to validating the concept, the airfoil and wing testing provided additional detailed data to better understand the aerodynamic advantage of such an unconventional trailing edge configuration. It is demonstrated that the concept represents a significant step in airfoil technology beyond that achieved with the Supercritical Airfoil. This concept provides the aerodynamicist an additional degree of design freedom and flexibility previously unrecognized.

  3. Airfoil Design and Rotorcraft Performance

    NASA Technical Reports Server (NTRS)

    Bousman, William G.

    2003-01-01

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

  4. 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.

  5. An experimental investigation of the effect of vortex generators on the aerodynamic characteristics of a NACA 0021 airfoil undergoing large amplitude pitch oscillations

    SciTech Connect

    Rueger, M.L.; Gregorek, G.M. . Dept. of Aeronautical and Astronautical Engineering)

    1991-04-01

    A NACA 0021 14-chord airfoil was subjected to large amplitude pitch oscillations in The Ohio State University Low Speed Wind Tunnel at a Reynolds number of 1.2 {times} 10{sup 6}. Surface pressures were measured with an electronically scanned pressure measurement system at sampling rates up to 50 Hz. Data were acquired for the clean airfoil and for the airfoil with vortex generators located at 0.1 and 0.3 chord distances aft of the leading edge. The vortex generators increase the maximum lift coefficient and the lift curve slope for both the static and dynamic tests. The magnitude and detail of the vortex generator effects were found to depend on the amplitude and frequency of the pitch oscillations. 18 refs., 76 figs., 21 tabs.

  6. Chordwise Pressure Distributions Over Several NACA 16-Series Airfoils at Transonic Mach Numbers up to 1.25

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.

    1959-01-01

    A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10(exp 6) to 2.8 x l0(exp 6) and in angle of attack from -10 to 12 deg. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.

  7. 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.

  8. 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.

  9. Airfoil profile drag increase due to acoustic excitation

    NASA Technical Reports Server (NTRS)

    Shearin, John G.; Jones, Michael G.

    1989-01-01

    A two-dimensional airfoil (NACA-0009) is subjected to high intensity pure-tone sound over a 1-5 kHz frequency range while immersed in a flow with 240 ft/sec velocity in a quiet flow facility with a Reynolds number of 3 million. Wake dynamic pressures are determined, and the momentum deficit is used to calculate a two-dimensional drag coefficient. Significant increases in drag are observed when the airfoil is subjected to high-intensity sound at critical frequencies. The increased drag is accompanied by movement of the natural transition location. When the transition is fixed by roughness at 10 percent chord, no further transition movement is observed in response to an acoustic Tollmien-Schlichting disturbance. However, a 4 percent increase in the sectional drag coefficient is noted. It is believed to be due to the sound exciting the flow near the airfoil surface (shear layer), thus causing the existing turbulence to become more intense, possess a higher mixing rate (momentum), and increase the skin friction.

  10. Airfoil profile drag increase due to acoustic excitation

    NASA Astrophysics Data System (ADS)

    Shearin, John G.; Jones, Michael G.

    1989-04-01

    A two-dimensional airfoil (NACA-0009) is subjected to high intensity pure-tone sound over a 1-5 kHz frequency range while immersed in a flow with 240 ft/sec velocity in a quiet flow facility with a Reynolds number of 3 million. Wake dynamic pressures are determined, and the momentum deficit is used to calculate a two-dimensional drag coefficient. Significant increases in drag are observed when the airfoil is subjected to high-intensity sound at critical frequencies. The increased drag is accompanied by movement of the natural transition location. When the transition is fixed by roughness at 10 percent chord, no further transition movement is observed in response to an acoustic Tollmien-Schlichting disturbance. However, a 4 percent increase in the sectional drag coefficient is noted. It is believed to be due to the sound exciting the flow near the airfoil surface (shear layer), thus causing the existing turbulence to become more intense, possess a higher mixing rate (momentum), and increase the skin friction.

  11. 10. DETAIL OF JUNCTION BETWEEN LOWER CHORD, VERTICAL LACED CHANNEL, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    10. DETAIL OF JUNCTION BETWEEN LOWER CHORD, VERTICAL LACED CHANNEL, FLOOR BEAM, EYE BAR, AND U-BOLT. WEST ABUTMENT. - River Road Bridge, Spanning Spring Creek in Spring Creek Township, Hallton, Elk County, PA

  12. 3. VIEW OF MAKERS PLATE ATTACHED TO UPPER CHORD MEMBER ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    3. VIEW OF MAKERS PLATE ATTACHED TO UPPER CHORD MEMBER WHICH STATES 'HUSTON AND CLEVELAND CONTRACTORS, COLUMBUS, OHIO, 1904.' - Main Street Parker Pony Truss Bridge, Main Street (Route 170) spanning Yellow Creek, Poland, Mahoning County, OH

  13. 31. LOWER CHORD / FLOOR STRUCTURE DETAIL OF THROUGH TRUSS. ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    31. LOWER CHORD / FLOOR STRUCTURE DETAIL OF THROUGH TRUSS. VIEW TO NORTH. - Abraham Lincoln Memorial Bridge, Spanning Missouri River on Highway 30 between Nebraska & Iowa, Blair, Washington County, NE

  14. 32. LOWER CHORD / FLOOR STRUCTURE DETAIL OF THROUGH TRUSS. ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    32. LOWER CHORD / FLOOR STRUCTURE DETAIL OF THROUGH TRUSS. VIEW TO NORTH. - Abraham Lincoln Memorial Bridge, Spanning Missouri River on Highway 30 between Nebraska & Iowa, Blair, Washington County, NE

  15. 13. VIEW OF TOP CHORD CONNECTION, SHOWING DIAGONAL AND VERTICAL ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. VIEW OF TOP CHORD CONNECTION, SHOWING DIAGONAL AND VERTICAL WEB MEMBERS AND TOP LATERAL BRACING, SOUTHEAST SPAN, LOOKING NORTHEAST - Linden Avenue Bridge, Spanning Purgatoire River on Linden Avenue, Trinidad, Las Animas County, CO

  16. 45. UPPER CHORD / END POST DETAIL OF WEST DECK ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    45. UPPER CHORD / END POST DETAIL OF WEST DECK TRUSS APPROACH SPAN, SHOWING FLOOR STRUCTURE. VIEW TO NORTHEAST. - MacArthur Bridge, Spanning Mississippi River on Highway 34 between IA & IL, Burlington, Des Moines County, IA

  17. 8. Pin connecting and eye bar nest, lower chord, down ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. Pin connecting and eye bar nest, lower chord, down river truss 132-0 Span 2 from Hot Metal Bridge. - Monongahela Connecting Railroad Company, Main Bridge, Spanning Monongahela River at mile post 3.1, Pittsburgh, Allegheny County, PA

  18. 7. Pin connections and eye bar nest, lower chord, up ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. Pin connections and eye bar nest, lower chord, up river truss, 321-4 Span 3. - Monongahela Connecting Railroad Company, Main Bridge, Spanning Monongahela River at mile post 3.1, Pittsburgh, Allegheny County, PA

  19. 6. Pin connection and eye bar nest, lower chord, up ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    6. Pin connection and eye bar nest, lower chord, up river truss, 321-4 Span 3. - Monongahela Connecting Railroad Company, Main Bridge, Spanning Monongahela River at mile post 3.1, Pittsburgh, Allegheny County, PA

  20. 61. VIEW OF THE HOISTING OF THE LOWER CHORD OF ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    61. VIEW OF THE HOISTING OF THE LOWER CHORD OF SHOOFLY BRIDGE LIFT SPAN, January 19, 1935 - Sacramento River Bridge, Spanning Sacramento River at California State Highway 275, Sacramento, Sacramento County, CA

  1. 8. Detail, lower chord, looking west from southeastern abutment ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. Detail, lower chord, looking west from southeastern abutment - Hayden Bridge, Spanning McKenzie River at Southern Pacific Railroad (moved from Springfield, Lake County, OR), Springfield, Lane County, OR

  2. 17. LOWER CHORD CONNECTION, FLOOR BEAM & STRINGER SYSTEMS, WALKWAY ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    17. LOWER CHORD CONNECTION, FLOOR BEAM & STRINGER SYSTEMS, WALKWAY CANTILEVER SUPPORTS; WALKWAY CANTILEVER SUPPORTS; LOOKING N - Traer Street Bridge, Spanning Shell Rock River at Traer Street, Greene, Butler County, IA

  3. 5. DETAIL VIEW OF UPPER CHORD MEMBER, SHOWING MAKER'S PLATE ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. DETAIL VIEW OF UPPER CHORD MEMBER, SHOWING MAKER'S PLATE STATING 'KING IRON BRIDGE & MFG. CO., K & F & Z KING PATENT, CLEVELAND, O.' - Smith Road Bowstring Arch Bridge, Spanning Sycamore Creek at Smith Road (TR 62), Lykens, Crawford County, OH

  4. 5. DETAIL VIEW OF LOWER CHORD CONNECTION FROM UNDERNEATH BRIDGE, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. DETAIL VIEW OF LOWER CHORD CONNECTION FROM UNDERNEATH BRIDGE, SHOWING DECK BEAM AND STIRRUP - John Bright No. 1 Iron Bridge, Spanning Poplar Creek at Havenport Road (TR 263), Carroll, Fairfield County, OH

  5. 17. DETAIL VIEW OF LOWER CHORD OF 1886 TRUSS, SHOWING ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    17. DETAIL VIEW OF LOWER CHORD OF 1886 TRUSS, SHOWING FLOOR BEAM (WITH CANTILEVERED SIDEWALK EXTENSION), STRINGERS AND LATERAL BRACING, LOOKING NORTHEAST - Sixth Street Viaduct, Spanning Burlington Northern Railroad & Valley Street, Burlington, Des Moines County, IA

  6. 13. Detail, connection point of end post, top chord, portal ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. Detail, connection point of end post, top chord, portal strut, and tension members at upstream side of west portal, view to northwest. - Dry Creek Bridge, Spanning Dry Creek at Cook Road, Ione, Amador County, CA

  7. 10. View of end portals, top chords, diagonals, verticals and ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    10. View of end portals, top chords, diagonals, verticals and strut connections. Looking from east span to east end of west span. - Boomershine Bridge, Spanning Twin Creek, Farmersville, Montgomery County, OH

  8. 5. UPPER CHORD UNDERSIDE VIEW SHOWING CHANNEL IRON AND LATTICE ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. UPPER CHORD UNDERSIDE VIEW SHOWING CHANNEL IRON AND LATTICE WORK CONSTRUCTION AND TENSION MEMBER CONNECTION. DECK VIEW LOOKING EAST. - Cunningham Lane Bridge, Spanning Pine River on Cunningham Lane near Highway 80, Rockbridge, Richland County, WI

  9. 12. NORTH WEB DETAIL, SHOWING LOWER CHORD CONNECTION OF WEBBED ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. NORTH WEB DETAIL, SHOWING LOWER CHORD CONNECTION OF WEBBED VERTICAL (PANEL POINT C). VIEW TO SOUTHWEST. - Lower Plymouth Rock Bridge, Spanning Upper Iowa River, Kendallville, Winneshiek County, IA

  10. 8. SOUTH VIEW, RIVETED CONNECTION POINT OF TOP CHORD, DIAGONAL ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. SOUTH VIEW, RIVETED CONNECTION POINT OF TOP CHORD, DIAGONAL TENSION MEMBER, INCLINED END POST AND VERTICAL MEMBER - Jordan Narrows Bridge, Crossing Jordan River at 9600 North, Lehi, Utah County, UT

  11. 10. DETAIL OF UNDERSIDE OF ARCH, SHOWING LOWER CHORDS, VERTICALS, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    10. DETAIL OF UNDERSIDE OF ARCH, SHOWING LOWER CHORDS, VERTICALS, LATERAL BRACES AND ABUTMENTS, VIEW TO SOUTHEAST. - Navajo Bridge, Spanning Colorado River at U.S. Highway 89 Alternate, Page, Coconino County, AZ

  12. 7. VIEW OF SOUTHERN PACIFIC RAILROAD BRIDGE, LOWER CHORD, LATERAL ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. VIEW OF SOUTHERN PACIFIC RAILROAD BRIDGE, LOWER CHORD, LATERAL BRACING AND FLOOR SYSTEM, LOOKING WEST. - Southern Pacific Railroad Bridge, Spanning Rio Grande at Southern Pacific Railroad, El Paso, El Paso County, TX

  13. 14. VIEW OF FLOOR BEAM CONNECTIONS WITH LOWER CHORD, BOTTOM ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    14. VIEW OF FLOOR BEAM CONNECTIONS WITH LOWER CHORD, BOTTOM LATERAL BRACING AND VERTICAL BRACE; FACING NORTHWEST. - Walker Bridge, Spanning Klamath River and connecting Highway 96 and Walker Road, Klamath River, Siskiyou County, CA

  14. 17. VIEW, LOOKING NORTH, SHOWING LOWER CHORD INTERMEDIATE JOINT, TYPICAL ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    17. VIEW, LOOKING NORTH, SHOWING LOWER CHORD INTERMEDIATE JOINT, TYPICAL PANELS 1, 2, 3, 4 AND 5 - Hegeman-Hill Street Bridge, Spanning Batten Kill, .65 mile West of Greenwich, Easton, Washington County, NY

  15. 8. View of substructure showing the lower chord of the ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. View of substructure showing the lower chord of the Howe truss, flared board-and-batten siding, and pier configuration - Drift Creek Bridge, Spanning Drift Creek on Drift Creek County Road, Lincoln City, Lincoln County, OR

  16. 15. DETAIL VIEW OF UPPER CHORD ON 1886 TRUSS, SHOWING ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    15. DETAIL VIEW OF UPPER CHORD ON 1886 TRUSS, SHOWING ENDPOST, PORTAL STRUT, LATERAL BRACING, HIP VERTICAL AND DIAGONAL, LOOKING NORTHEAST - Sixth Street Viaduct, Spanning Burlington Northern Railroad & Valley Street, Burlington, Des Moines County, IA

  17. 16. DETAIL VIEW OF UPPER CHORD ON 1886 TRUSS, SHOWING ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    16. DETAIL VIEW OF UPPER CHORD ON 1886 TRUSS, SHOWING TYPICAL VERTICAL, STRUT, LATERAL BRACING AND DIAGONAL, LOOKING SOUTHEAST - Sixth Street Viaduct, Spanning Burlington Northern Railroad & Valley Street, Burlington, Des Moines County, IA

  18. 8. UPPER INSIDE CHORD, VERTICAL, LATERAL STRUT, UPPER LATERAL & ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. UPPER INSIDE CHORD, VERTICAL, LATERAL STRUT, UPPER LATERAL & GUSSET PLATE, ONE DIAGONAL BRACE - Enterprise Parker Truss Bridge, Spanning Smoky Hill River on K-43 Highway, Enterprise, Dickinson County, KS

  19. 7. UPPER INSIDE CHORD, VERTICAL, LATERAL STRUT, UPPER LATERAL & ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. UPPER INSIDE CHORD, VERTICAL, LATERAL STRUT, UPPER LATERAL & GUSSET PLATE, TWO DIAGONAL BRACES - Enterprise Parker Truss Bridge, Spanning Smoky Hill River on K-43 Highway, Enterprise, Dickinson County, KS

  20. 32. VERTICAL / STRUT / UPPER CHORD DETAIL AT PINCONNECTED ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    32. VERTICAL / STRUT / UPPER CHORD DETAIL AT PIN-CONNECTED EXPANSION JOINT BETWEEN CANTILEVER ARM AND SUSPENDED SPAN. VIEW TO NORTHEAST. - MacArthur Bridge, Spanning Mississippi River on Highway 34 between IA & IL, Burlington, Des Moines County, IA

  1. 43. DETAIL OF PINNED UPPER CHORD CONNECTION BETWEEN ANCHOR ARM ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    43. DETAIL OF PINNED UPPER CHORD CONNECTION BETWEEN ANCHOR ARM AND SUSPENDED (PANEL 67). VIEW TO NORTH. - Blue Water Bridge, Spanning St. Clair River at I-69, I-94, & Canadian Route 402, Port Huron, St. Clair County, MI

  2. 11. VIEW OF PIN CONNECTION, SOUTH WEB, SHOWING TOP CHORD, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    11. VIEW OF PIN CONNECTION, SOUTH WEB, SHOWING TOP CHORD, LATTICE BRACING, HIP VERTICAL, EYEBARS, TOP LATERAL BRACING, AND STRUTS, LOOKING SOUTH - Four Mile Bridge, Spanning Elk River on County Road 42, Steamboat Springs, Routt County, CO

  3. Detail of lower chord connections Waterville Bridge, Spanning Swatara ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Detail of lower chord connections - Waterville Bridge, Spanning Swatara Creek at Appalachian Trail (moved from Little Pine Creek at State Route 44, Waterville, Lycoming County), Green Point, Lebanon County, PA

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

    NASA Technical Reports Server (NTRS)

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

    1956-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  6. Experimental and Computational Investigation of Lift-Enhancing Tabs on a Multi-Element Airfoil

    NASA Technical Reports Server (NTRS)

    Ashby, Dale L.

    1996-01-01

    An experimental and computational investigation of the effect of lift-enhancing tabs on a two-element airfoil has been 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. An NACA 63(2)-215 ModB airfoil with a 30% 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. A combination of tabs located at the main element and flap trailing edges increased the airfoil lift coefficient by 11% relative to the highest lift coefficient achieved by any baseline configuration at an angle of attack of 0 deg, and C(sub 1max) was increased by 3%. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computed results predicted all of the trends observed in the experimental data quite well. In addition, 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 air-foil or flap trailing 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 effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

  7. Heat transfer and film cooling effectiveness in a linear airfoil cascade

    SciTech Connect

    Abuaf, N.; Bunker, R.; Lee, C.P.

    1997-04-01

    A warm (315 C) wind tunnel test facility equipped with a linear cascade of film cooled vane airfoils was used in the simultaneous determination of the local gas side heat transfer coefficients and the adiabatic film cooling effectiveness. The test rig can be operated in either a steady-state or a transient mode. The steady-state operation provides adiabatic film cooling effectiveness values while the transient mode generates data for the determination of the local heat transfer coefficients from the temperature-time variations and of the film effectiveness from the steady wall temperatures within the same aerothermal environment. The linear cascade consists of five airfoils. The 14% cascade inlet free-stream turbulence intensity is generated by a perforated plate, positioned upstream of the airfoil leading edge. For the first transient tests, five cylinders having roughly the same blockage as the initial 20% axial chord of the airfoils were used. The cylinder stagnation point heat transfer coefficients compare well with values calculated from correlations. Static pressure distributions measured over an instrumented airfoil agree with inviscid predictions. Heat transfer coefficients and adiabatic film cooling effectiveness results were obtained with a smooth airfoil having three separate film injection locations, two along the suction side, and the third one covering the leading edge showerhead region. Near the film injection locations, the heat transfer coefficients increase with the blowing film. At the termination of the film cooled airfoil tests, the film holes were plugged and heat transfer tests were conducted with non-film cooled airfoils. These results agree with boundary layer code predictions.

  8. 19. 80 foot pony truss view of upper chord ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    19. 80 foot pony truss - view of upper chord pin connection at the end post, typical of the five 80 foot trusses and similar to the 64 foot tress. There are two pair per pony truss for a total of 24. Shown are the vertical lace post, end post, top chord member, and a diagonal member. - Weidemeyer Bridge, Spanning Thomes Creek at Rawson Road, Corning, Tehama County, CA

  9. 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.

  10. Flying-hot-wire study of two-dimensional mean flow past an NACA 4412 airfoil at maximum lift

    NASA Technical Reports Server (NTRS)

    Coles, D.; Wadcock, A. J.

    1978-01-01

    Hot-wire measurements have been made in the boundary layer, the separated region, and the near wake for flow past an NACA 4412 airfoil at maximum lift. The Reynolds number based on chord was about 1,500,000. The main instrumentation was a hot-wire probe mounted on the end of a rotating arm. A digital computer was used to control synchronized sampling of hot-wire data at closely spaced points along the probe arc. Ensembles of data were obtained at several thousand locations in the flow field. The data include intermittency, two components of mean velocity, and twelve mean values for double, triple, and quadruple products of two velocity fluctuations. The data are available on punched cards in raw form and also after use of smoothing and interpolation routines to obtain values on a fine rectangular grid aligned with the airfoil chord. The data are displayed in the paper as contour plots.

  11. Tabulation of data from tests of an NPL 9510 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Jenkins, R. V.

    1983-01-01

    The tabulated data from tests of a six inch chord NPL 9510 airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The tests were performed over the following range of conditions: Mach numbers of 0.35 to 0.82, total temperature of 94 K to 300 K, total pressure of 1.20 to 5.81 atm, Reynolds number based on chord of 1.34 x 10 to the 6th to 48.23 x 10 to the 6th, and angle of attack of 0 deg to 6 deg. The NPL 9510 airfoil was observed to have decreasing drag coefficient up to the highest test Reynolds number.

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

    SciTech Connect

    Somers, D. M.

    2005-01-01

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

  13. The Influences of Progression Type and Distortion on the Perception of Terminal Power Chords

    ERIC Educational Resources Information Center

    Juchniewicz, Jay; Silverman, Michael J.

    2013-01-01

    The purpose of this study was to investigate the tonal perception and restoration of thirds within power chords with the instruments and sounds idiosyncratic to the Western rock/pop genre. Four separate chord sequences were performed on electric guitar in four versions; as full chord and power chord versions as well as under both clean-tone and…

  14. 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.

  15. Airfoil with nested cooling channels

    SciTech Connect

    Levengood, J.L.; Auxier, T.A.

    1988-06-28

    A turbine blade is described which consists of a root portion and wall means integral with the root portion defining an airfoil, the wall means including a pressure sidewall and a suction sidewall, joined together to define a forwardly located leading edge and rearwardly located trailing edge of the airfoil and spaced apart to define a spanwise and chordwise extending coolant cavity within the airfoil, and root portion including root passage means therethrough for receiving coolant fluid form outside the blade and for directing the fluid into the airfoil cavity.

  16. 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.

  17. Direct numerical simulations of tonal noise generated by laminar flow past airfoils

    NASA Astrophysics Data System (ADS)

    Sandberg, R. D.; Jones, L. E.; Sandham, N. D.; Joseph, P. F.

    2009-03-01

    A numerical investigation is presented of noise generated by flow past symmetric NACA airfoils with different thickness and at various angles of attack at M=0.4 and a Reynolds number based on chord of Re=50,000. Direct numerical simulations (DNS) are employed to directly compute both the near-field hydrodynamics and the far-field sound. The DNS data are then used to investigate whether the approach of determining tonal noise radiation based on the surface pressure difference, as done in the classical trailing-edge theory of Amiet, yields satisfactory results for finite thickness airfoils subject to mean loading effects. In addition, the accuracy of Amiet's surface pressure jump function is evaluated. Overall, the modified theory of Amiet appears to be suitable for finite thickness airfoils up to moderate incidence. However, when increasing the airfoil thickness to 12% chord, which corresponds to a trailing-edge angle of 16.8∘, an unexpected phase change between the incident and scattered pressure is found at the frequency of the forced instability waves. This phase change is attributed to the flow oscillating around the trailing edge at a separate wake frequency. For the largest incidence investigated, Amiet's response function does not predict the total surface pressure difference as accurately as for zero or small incidence at the vortex shedding frequency, resulting in a poor prediction of the directivity and amplitude of the acoustic pressure. Moreover, predicting the airfoil self-noise based on the surface pressure difference appears not to be generally applicable at higher angles of attack because the radiated sound is only partly due to classical trailing-edge noise mechanisms. In these cases, it appears as if volume sources in the flow cannot be neglected.

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

    NASA Astrophysics Data System (ADS)

    Manela, Avshalom

    2013-12-01

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

  19. 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.

  20. NREL airfoil families for HAWTs

    SciTech Connect

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

    1995-12-31

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

  1. 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.

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

    SciTech Connect

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

    1988-04-01

    The aerodynamic similarity between a small (4-in. chord) wind tunnel model and a full-scale wind turbine blade (24-ft tip section with a 36-in. chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-In. High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-Ft Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the three-dimensional effects of blade twist and downwash in the blade tip section. Good correlation was obtained between the results of the two different experimental methods for angles of attack from -4/degree/ to 36/degree/ and aileron deflections from 0/degree/ to 90/degree/. 4 refs., 11 figs., 1 tab.

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

  4. Active Control of Separation from the Slat Shoulder of a Supercritical Airfoil

    NASA Technical Reports Server (NTRS)

    Pack, LaTunia G.; Schaeffler, Norman W.; Yao, Chung-Sheng; Seifert, Avi

    2002-01-01

    Active flow control in the form of zero-mass-flux excitation was applied at the slat shoulder of a simplified high-lift airfoil to delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge slat and a 25% chord simply hinged trailing edge flap. The cruise configuration data was successfully reproduced, repeating previous experiments. The effects of flap and slat deflection angles on the performance of the airfoil integral parameters were quantified. Detailed flow features were measured as well, in an attempt to identify optimal actuator placement. The measurements included: steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization and Particle Image Velocimetry (PIV). High frequency periodic excitation was applied to delay the occurrence of slat stall and improve the maximum lift by 10 to 15%. Low frequency amplitude modulation was used to reduce the oscillatory momentum coefficient by roughly 50% with similar aerodynamic performance.

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

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  6. Ice-induced unsteady flowfield effects on airfoil performance

    NASA Astrophysics Data System (ADS)

    Gurbacki, Holly Marie

    Numerical prediction of iced-airfoil performance prior to and at maximum lift is often inaccurate due to large-scale flow unsteadiness. New computational models are being developed to improve predictions of complex separated flowfields; however, experimental data are required to improve and validate these algorithms. The objective of this investigation was to examine the unsteady flow behavior and the time-dependent performance of an iced airfoil to determine the flowfield characteristics with the most influence on airfoil performance, especially near stall. A NACA 0012 airfoil with two-dimensional and three-dimensional leading-edge simulated glaze ice shapes was tested in a wind tunnel at Reynolds numbers 1.8 x 106 and 1.0 x 106. Time-dependent surface pressure measurements were used to calculate root-mean-square lift and quarter-chord pitching-moment coefficients. Surface and flowfield visualization and wake hot-wire data were acquired. Spectral, correlation and phase-angle analyses were performed. The most significant unsteady flowfield effect on the iced-airfoil performance was a low-frequency flow phenomenon on the order of 10 Hz that resulted in Strouhal numbers of 0.0048--0.0101. The low-frequency oscillation produced large-scale pressure fluctuations nears eparation at high angles of attack and elevated lift and moment fluctuations as low as alpha = 4°. The low-frequency motion of surface pressure coefficients convected downstream at velocities 4%--34% of the freestream value and in one case, upstream at 0.18Uinfinity. The iced-airfoil flowfield exhibited a separation bubble of varying thickness and fluctuating reattachment, characteristics similar to those associated with the low-frequency shear-layer flapping and bubble growth and decay of other separated and reattached flows. Vortex structures observed in the shear layer were presumed to be the cause of large-scale pressure fluctuations upstream of reattachment at small angles of attack. Pressure

  7. Influence of structural flexibility on the wake vortex pattern of airfoils undergoing harmonic pitch oscillation

    NASA Astrophysics Data System (ADS)

    Monnier, B.; Naguib, A. M.; Koochesfahani, M. M.

    2015-04-01

    Reported herein is an investigation of the influence of the structural flexibility of sinusoidally pitching airfoils on the pattern of vorticity shed into the wake. For rigid airfoils, it is well known that, depending on the oscillation frequency and amplitude, this pattern takes the form of the classical or reverse von Kármán vortex street. The pattern may be characterized by the vortex circulation ( Γ o ), vortex-to-vortex streamwise and cross-stream spacing ( a and b, respectively), and vortex core radius ( R). In the present work, these four parameters are obtained from particle image velocimetry measurements in the wake of airfoils consisting of a rigid "head" and flexible "tail" at chord Reynolds number of 2010 for different tail flexibilities. The results show that flexible airfoils exhibit the switch from classical to reverse von Kármán vortex street (i.e., change in the sign of b) at a reduced frequency of oscillation lower than their rigid counterpart. At a given oscillation frequency, the Strouhal number at which this switch occurs is smallest for a given airfoil structural flexibility; which becomes stiffer with increasing frequency. Using Strouhal number based on the actual trailing edge oscillation amplitude, reasonable scaling is found of the dependence of not only b but also Γ o , a and R on the motion and structure parameters for all airfoils investigated. These results are complemented with analyses using a vortex array model, which together with the identified scaling of the wake vortex parameters, provide basis for the computation of the net thrust acting on the airfoil.

  8. Wing chord extension on D-558-2

    NASA Technical Reports Server (NTRS)

    1953-01-01

    had completed 21 contractor flights by Douglas pilots Eugene F. May and Bill Bridgeman in November 1950. In this jet-and-rocket-propelled craft, Scott Crossfield and Walter Jones began the NACA's investigation of pitch-up lasting from September 1951 well into the summer of 1953. They flew the Skyrocket with a variety of wing-fence, wing-slat, and leading-edge chord extension configurations, performing various maneuvers as well as straight-and-level flying at transonic speeds. While fences significantly aided recovery from pitch-up conditions, leading edge chord extensions did not, disproving wind-tunnel tests to the contrary. Slats (long, narrow auxiliary airfoils) in the fully open position eliminated pitch-up except in the speed range around Mach 0.8 to 0.85. In June 1954, Crossfield began an investigation of the effects of external stores (bomb shapes and fuel tanks) upon the aircraft's transonic behavior. McKay and Stanley Butchart completed the NACA's investigation of this issue, with McKay flying the final mission on August 28, 1956. Besides setting several records, the Skyrocket pilots had gathered important data and understanding about what would and would not work to provide stable, controlled flight of a swept-wing aircraft in the transonic and supersonic flight regimes. The data they gathered also helped to enable a better correlation of wind-tunnel test results with actual flight values, enhancing the abilities of designers to produce more capable aircraft for the armed services, especially those with swept wings. Moreover, data on such matters as stability and control from this and other early research airplanes aided in the design of the century series of fighter airplanes, all of which featured the movable horizontal stabilizers first employed on the X-1 and D-558 series.

  9. Lift enhancing tabs for airfoils

    NASA Technical Reports Server (NTRS)

    Ross, James C. (Inventor)

    1994-01-01

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

  10. High Reynolds Number tests of the NASA SC(2)-0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Mineck, Raymond E.; Lawing, Pierce L.

    1987-01-01

    A wind-tunnel investigation of the NASA SC(2)-0012 airfoil has been conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation supplements the two-dimensional airfoil studies of the Advanced Technology Airfoil Test Program. The Mach number was varied from 0.60 to 0.84. The stagnation temperature and pressure were varied to provide a Reynolds number range from 6 to 40 x 10 to the 6th power based on a 6.0-in. (15.24-cm) airfoil chord. No corrections for wind-tunnel wall interference have been made to the data. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions without any analysis.

  11. Convective heat transfer measurements from a NACA 0012 airfoil in flight and in the NASA Lewis Icing Research Tunnel

    NASA Technical Reports Server (NTRS)

    Poinsatte, Philip E.; Vanfossen, G. James; Dewitt, Kenneth J.

    1989-01-01

    Local heat transfer coefficients were measured on a smooth and roughened NACA 0012 airfoil. Heat transfer measurements on the 0.533 m chord airfoil were made both in flight on the NASA Lewis Twin Otter Icing Research Aircraft and in the NASA Lewis Icing Research Tunnel (IRT). Roughness was obtained by the attachment of uniform 2 mm diameter hemispheres to the airfoil surface in 4 distinct patterns. Flight data were taken for the smooth and roughened airfoil at various Reynolds numbers based on chord in the range 1.24 to 2.50 x 10(exp 6) and at various angles of attack up to 4 deg. During these flight tests, the free stream velocity turbulence intensity was found to be very low (less than 0.1 percent). Wind tunnel data were acquired in the Reynolds number range 1.20 to 4.25 x 10(exp 6) and at angles of attack from -4 to 8 deg. The turbulence intensity in the IRT was 0.5 to 0.7 percent with the cloud generating sprays off. A direct comparison was made between the results obtained in flight and in the IRT. The higher level of turbulence in the IRT vs. flight had little effect on the heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the high Reynolds numbers. Roughness generally increased the heat transfer.

  12. Turbine airfoil to shround attachment

    SciTech Connect

    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.

  13. Dynamic stall on a pitching and surging airfoil

    NASA Astrophysics Data System (ADS)

    Dunne, Reeve; McKeon, Beverley J.

    2015-08-01

    Vertical axis wind turbine blades undergo dynamic stall due to the large angle of attack variation they experience during a turbine rotation. The flow over a single blade was modeled using a sinusoidally pitching and surging airfoil in a non-rotating frame with a constant freestream flow at a mean chord Reynolds number of . Two-dimensional, time-resolved velocity fields were acquired using particle image velocimetry. Vorticity contours were used to visualize shear layer and vortex activity. A low-order model of dynamic stall was developed using dynamic mode decomposition, from which primary and secondary dynamic separation modes were identified. The interaction between these two modes was able to capture the physics of dynamic stall and as such can be extended to other turbine configurations and problems in unsteady aerodynamics. Results from the linear pitch/surge frame are extrapolated to the rotating VAWT frame to investigate the behavior of identified flow structures.

  14. A direct numerical simulation investigation of the synthetic jet frequency effects on separation control of low-Re flow past an airfoil

    NASA Astrophysics Data System (ADS)

    Zhang, Wei; Samtaney, Ravi

    2015-05-01

    We present results of direct numerical simulations of a synthetic jet (SJ) based separation control of flow past a NACA-0018 (National Advisory Committee for Aeronautics) airfoil, at 10° angle of attack and Reynolds number 104 based on the airfoil chord length C and uniform inflow velocity U0. The actuator of the SJ is modeled as a spanwise slot on the airfoil leeward surface and is placed just upstream of the leading edge separation position of the uncontrolled flow. The momentum coefficient of the SJ is chosen at a small value 2.13 × 10-4 normalized by that of the inflow. Three forcing frequencies are chosen for the present investigation: the low frequency (LF) F+ = feC/U0 = 0.5, the medium frequency (MF) F+ = 1.0, and the high frequency (HF) F+ = 4.0. We quantify the effects of forcing frequency for each case on the separation control and related vortex dynamics patterns. The simulations are performed using an energy conservative fourth-order parallel code. Numerical results reveal that the geometric variation introduced by the actuator has negligible effects on the mean flow field and the leading edge separation pattern; thus, the separation control effects are attributed to the SJ. The aerodynamic performances of the airfoil, characterized by lift and lift-to-drag ratio, are improved for all controlled cases, with the F+ = 1.0 case being the optimal one. The flow in the shear layer close to the actuator is locked to the jet, while in the wake this lock-in is maintained for the MF case but suppressed by the increasing turbulent fluctuations in the LF and HF cases. The vortex evolution downstream of the actuator presents two modes depending on the frequency: the vortex fragmentation and merging mode in the LF case where the vortex formed due to the SJ breaks up into several vortices and the latter merge as convecting downstream; the discrete vortices mode in the HF case where discrete vortices form and convect downstream without any fragmentation and merging

  15. Noise models for superoperators in the chord representation

    SciTech Connect

    Aolita, Mario Leandro; Garcia-Mata, Ignacio; Saraceno, Marcos

    2004-12-01

    We study many-qubit generalizations of quantum noise channels that can be written as an incoherent sum of translations in phase space, for which the chord representation results specially useful. Physical descriptions in terms of the spectral properties of the superoperator and the action in phase space are provided. A very natural description of decoherence leading to a preferred basis is achieved with diffusion along a phase space line. The numerical advantages of using the chord representation are illustrated in the case of coarse-graining noise.

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  17. 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.

  18. Calculation of steady and unsteady airfoil flow fields via the Navier-Stokes equations

    NASA Technical Reports Server (NTRS)

    Shamroth, S. J.

    1985-01-01

    A compressible time-dependent procedure for the two-dimensional ensemble averaged Navier-Stokes equations has been applied to the isolated airfoil problem in steady and unsteady flows. The procedure solves the governing equations via the linearized block implicit technique. Turbulence is modeled either via a mixing length or turbulence energy approach. The equations are solved in general non-orthogonal form with no-slip boundary conditions applied at the airfoil surface. Results are presented for airfoils at constant incidence, an airfoil in ramp motion and an airfoil oscillating through a dynamic stall loop. In general, steady converged solutions are obtained within 70 time steps over the range of Mach numbers considered. Comparisons with measured data show good agreement between computation and measurement.

  19. Unsteady pressure measurements on a supercritical airfoil at high Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Hess, R. W.

    1989-01-01

    Steady and unsteady pressures were measured on a 14 percent supercritical airfoil at transonic Mach numbers at Reynolds numbers from 6,000,000 to 35,000,000. Instrumentation techniques were developed to measure unsteady pressures in a cryogenic tunnel at flight Reynolds numbers. Experimental steady data, corrected for wall effects show very good agreement with calculations from a full potential code with an interacted boundary layer. The steady and unsteady pressures both show a shock position that is dependent on Reynolds number. For a supercritical pressure distribution at a chord Reynolds number of 35,000,000 laminar flow was observed between the leading edge and the shock wave at 45 percent chord.

  20. Development of a Fowler flap system for a high performance general aviation airfoil

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

    A two-dimensional wind-tunnel evaluation of two Fowler flap configurations on the new GA(W)-1 airfoil was conducted. One configuration used a computer-designed 29-percent chord Fowler flap. The second configuration was modified to have increased Fowler action with a 30-percent chord flap. Force, pressure, and flow-visualization data were obtained at Reynolds numbers of 2.2 million to 2.9 million. Optimum slot geometry and performance were found to be close to computer predictions. A C sub L max of 3.8 was achieved. Optimum flap deflection, slot gap, and flap overlap are presented as functions of C sub L. Tests were made with the lower surface cusp filled in to show the performance penalties that result. Some data on the effects of adding vortex generators and hinged-plate spoilers were obtained.

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

    NASA Technical Reports Server (NTRS)

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

    1986-01-01

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

  2. Active Control of Separation From the Flap of a Supercritical Airfoil

    NASA Technical Reports Server (NTRS)

    Melton, La Tunia Pack; Yao, Chung-Sheng; Seifert, Avi

    2003-01-01

    Active flow control in the form of periodic zero-mass-flux excitation was applied at several regions on the leading edge and trailing edge flaps of a simplified high-lift system t o delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approx.= 10) and low frequency amplitude modulation (F(+)AM approx.= 1) of the high frequency excitation were used for control. Preliminary efforts to combine leading and trailing edge flap excitations are also reported.

  3. High-Lift System for a Supercritical Airfoil: Simplified by Active Flow Control

    NASA Technical Reports Server (NTRS)

    Melton, LaTunia Pack; Schaeffler, Norman W.; Lin, John C.

    2007-01-01

    Active flow control wind tunnel experiments were conducted in the NASA Langley Low-Turbulence Pressure Tunnel using a two-dimensional supercritical high-lift airfoil with a 15% chord hinged leading-edge flap and a 25% chord hinged trailing-edge flap. This paper focuses on the application of zero-net-mass-flux periodic excitation near the airfoil trailing edge flap shoulder at a Mach number of 0.1 and chord Reynolds numbers of 1.2 x 10(exp 6) to 9 x 10(exp 6) with leading- and trailing-edge flap deflections of 25 deg. and 30 deg., respectively. The purpose of the investigation was to increase the zero-net-mass-flux options for controlling trailing edge flap separation by using a larger model than used on the low Reynolds number version of this model and to investigate the effect of flow control at higher Reynolds numbers. Static and dynamic surface pressures and wake pressures were acquired to determine the effects of flow control on airfoil performance. Active flow control was applied both upstream of the trailing edge flap and immediately downstream of the trailing edge flap shoulder and the effects of Reynolds number, excitation frequency and amplitude are presented. The excitations around the trailing edge flap are then combined to control trailing edge flap separation. The combination of two closely spaced actuators around the trailing-edge flap knee was shown to increase the lift produced by an individual actuator. The phase sensitivity between two closely spaced actuators seen at low Reynolds number is confirmed at higher Reynolds numbers. The momentum input required to completely control flow separation on the configuration was larger than that available from the actuators used.

  4. 28. 100 foot through truss a typical lower chord ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    28. 100 foot through truss - a typical lower chord pin connection, located below each vertical lace post on the through trusses. Each truss has four of these for a total of eight. Shown is the floor beam below the pin connection, and the four inch conduit. - Weidemeyer Bridge, Spanning Thomes Creek at Rawson Road, Corning, Tehama County, CA

  5. 35. View is the underside of a lower chord pin ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    35. View is the underside of a lower chord pin connection showing the top strut, along with lateral and diagonal members. There are four of these per through truss for a total of eight. - Weidemeyer Bridge, Spanning Thomes Creek at Rawson Road, Corning, Tehama County, CA

  6. Chord-based image reconstruction from clinical projection data

    NASA Astrophysics Data System (ADS)

    King, Martin; Xia, Dan; Pan, Xiaochuan; Vannier, Michael; Köhler, Thomas; La Riviére, Patrick; Sidky, Emil; Giger, Maryellen

    2008-03-01

    Chord-based algorithms can eliminate cone-beam artifacts in images reconstructed from a clinical computed tomography (CT) scanner. The feasibility of using chord-based reconstruction algorithms was evaluated with three clinical CT projection data sets. The first projection data set was acquired using a clinical 64-channel CT scanner (Philips Brilliance 64) that consisted of an axial scan from a quality assurance phantom. Images were reconstructed using (1) a full-scan FDK algorithm, (2) a short-scan FDK algorithm, and (3) the chord-based backprojection filtration algorithm (BPF) using full-scan data. The BPF algorithm was capable of reproducing the morphology of the phantom quite well, but exhibited significantly less noise than the two FDK reconstructions as well as the reconstruction obtained from the clinical scanner. The second and third data sets were obtained from scans of a head phantom and a patient's thorax. For both of these data sets, the BPF reconstructions were comparable to the short-scan FDK reconstructions in terms of image quality, although sharper features were indistinct in the BPF reconstructions. This research demonstrates the feasibility of chord-based algorithms for reconstructing images from clinical CT projection data sets and provides a framework for implementing and testing algorithmic innovations.

  7. 17. DETAIL OF LOWER CHORDS AND BOTTOM LATERAL BRACING OF ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    17. DETAIL OF LOWER CHORDS AND BOTTOM LATERAL BRACING OF WEST DECK TRUSS AND PIER NO. 2, FROM WEST RIVERBANK. VIEW TO EAST. - MacArthur Bridge, Spanning Mississippi River on Highway 34 between IA & IL, Burlington, Des Moines County, IA

  8. 15. Detail showing lower chord pinconnected to vertical member, showing ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    15. Detail showing lower chord pin-connected to vertical member, showing floor beam riveted to extension of vertical member below pin-connection, and showing brackets supporting cantilevered sidewalk. View to southwest. - Selby Avenue Bridge, Spanning Short Line Railways track at Selby Avenue between Hamline & Snelling Avenues, Saint Paul, Ramsey County, MN

  9. 27. DETAIL VIEW OF LOWER CHORD OF 1903 GIRDERS, SHOWING ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    27. DETAIL VIEW OF LOWER CHORD OF 1903 GIRDERS, SHOWING PIER CAP, CANTILEVERED SIDEWALK BRACKETS, FLOOR BEAMS, LATERAL BRACING AND UNDERSIDE OF CORRUGATED STEEL DECK. LOOKING SOUTHEAST - Sixth Street Viaduct, Spanning Burlington Northern Railroad & Valley Street, Burlington, Des Moines County, IA

  10. 14. Detail, connection point of upper chord and end post, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    14. Detail, connection point of upper chord and end post, showing aforementioned members, vertical and diagonal tension members, lateral, latticed portal strut and decorative strut bracing. Note also decorative fluted Classical urn atop end post. View to west of upstream side of northwest portal. - Red Bank Creek Bridge, Spanning Red Bank Creek at Rawson Road, Red Bluff, Tehama County, CA

  11. 13. DETAIL OF CONNECTION BETWEEN TOP CHORD AND POST IN ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. DETAIL OF CONNECTION BETWEEN TOP CHORD AND POST IN WEST TRUSS, SHOWING CHANNELS AND REINFORCED CAST-IRON LACING, I-BEAMS FASTENED TOGETHER WITH RIVETTED PLATES, AND ASSEMBLY OF DIAGONAL EYE BEAM AND BOLT; VIEW FROM EAST SIDE. - Mitchell's Mill Bridge, Spanning Winter's Run on Carrs Mill Road, west of Bel Air, Bel Air, Harford County, MD

  12. Analyzing Sound Waves Produced by Musical Notes & Chords.

    ERIC Educational Resources Information Center

    Cassidy, Michael

    This project description is designed to show how graphing calculators and calculator-based laboratories (CBL) can be used to explore topics in the physics of sound. The activities address topics such as sound waves, musical notes, and chords. Teaching notes, calculator instructions, and blackline masters are included. (MM)

  13. Interior detail, building 810, view to north showing curved chord ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Interior detail, building 810, view to north showing curved chord sections of roof trusses, 90mm lens plus electronic flash fill lighting. - Travis Air Force Base, B-36 Hangar, Between Woodskill Avenue & Ellis, adjacent to Taxiway V & W, Fairfield, Solano County, CA

  14. 31. DECK / VERTICAL / UPPER CHORD DETAIL OF THROUGH ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    31. DECK / VERTICAL / UPPER CHORD DETAIL OF THROUGH TRUSS AT PIN-CONNECTED EXPANSION JOINT BETWEEN CANTILEVER ARM AND SUSPENDED SPAN. VIEW TO NORTHEAST. - MacArthur Bridge, Spanning Mississippi River on Highway 34 between IA & IL, Burlington, Des Moines County, IA

  15. 6. CONNECTION OF END POST AND TOP CHORD (TWO CHANNELS ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    6. CONNECTION OF END POST AND TOP CHORD (TWO CHANNELS CONNECTED WITH COVER PLATES AND LACING BARS) WITH DIAGONALS (LOOP WELDED DOUBLE RECTILINEAR EYE BARS). - Daphna Creek Pratt Truss Bridge, State Route 1414, Holly Hill Street, Spanning Dahna Creek, Broadway, Rockingham County, VA

  16. Aerodynamic sound of flow past an airfoil

    NASA Technical Reports Server (NTRS)

    Wang, Meng

    1995-01-01

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

  17. 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.

  18. 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.

  19. Large-eddy simulation of flow around an airfoil on a structured mesh

    NASA Technical Reports Server (NTRS)

    Kaltenbach, Hans-Jakob; Choi, Haecheon

    1995-01-01

    The diversity of flow characteristics encountered in a flow over an airfoil near maximum lift taxes the presently available statistical turbulence models. This work describes our first attempt to apply the technique of large-eddy simulation to a flow of aeronautical interest. The challenge for this simulation comes from the high Reynolds number of the flow as well as the variety of flow regimes encountered, including a thin laminar boundary layer at the nose, transition, boundary layer growth under adverse pressure gradient, incipient separation near the trailing edge, and merging of two shear layers at the trailing edge. The flow configuration chosen is a NACA 4412 airfoil near maximum lift. The corresponding angle of attack was determined independently by Wadcock (1987) and Hastings & Williams (1984, 1987) to be close to 12 deg. The simulation matches the chord Reynolds number U(sub infinity)c/v = 1.64 x 10(exp 6) of Wadcock's experiment.

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

    NASA Astrophysics Data System (ADS)

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

    2016-06-01

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

  1. Determination of the pressure drag of airfoils by integration of surface pressures

    NASA Technical Reports Server (NTRS)

    Phillips, William H.

    1990-01-01

    A study was conducted of the causes of pressure drag of subsonic airfoils. In a previous paper by the author, the pressure drag is obtained by calculating the total drag from the momentum defect in the boundary layer at the trailing edge and subtracting the friction drag obtained from integration of surface friction along the chord. Herein, the pressure drag is obtained by integrating the streamwise components of surface pressure around the airfoil. Studies were made to verify the accuracy of the integration procedure. The values of pressure drag were much smaller than those obtained by the previous method. This lack of agreement is attributed to the difficulty of calculating boundary layer conditions in the vicinity of the trailing edge and to the extreme sensitivity of the circulation and lift to the trailing edge conditions. The results of these studies are compared with those of previous investigations.

  2. Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

    Wind-tunnel tests were conducted in the Langley low-turbulence pressure tunnel to evaluate the effects on performance of modifying a 17-percent-thick low-speed airfoil. The airfoil contour was altered to reduce the pitching-moment coefficient by increasing the forward loading and to increase the climb lift-drag ratio by decreasing the aft upper surface pressure gradient. The tests were conducted over a Mach number range from 0.07 to 0.32, a chord Reynolds number range 1.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle-of-attack range from about -10 deg to 20 deg.

  3. The effect of acoustic forcing on trailing edge separation and near wake development of an airfoil

    NASA Technical Reports Server (NTRS)

    Huang, L. S.; Bryant, T. D.; Maestrello, L.

    1988-01-01

    An experimental study was conducted to investigate the effect of acoustic forcing on flow fields near the trailing edge of a symmetric airfoil at zero angle of attack. At low chord Reynolds numbers, the boundary layers separate from the surfaces in the rear part of the airfoil and create recirculation regions near the trailing edge. It is shown that with the introduction of acoustic forcing through a slot in the vicinity of the separation point, periodic large-scale structures are generated in the trailing edge region. Significant reduction of trailing edge separation is achieved. It is also found that the most effective forcing frequency to control trailing edge separation is the wake vortex shedding frequency. As a result of forcing, applied only on the upper surface, the upper boundary layer is accelerated and the flow over the lower surface decelerated. Consequently, an asymmetric wake is formed. The results presented indicate that the development of the near wake varies with forcing conditions.

  4. Evaluation of Icing Scaling on Swept NACA 0012 Airfoil Models

    NASA Technical Reports Server (NTRS)

    Tsao, Jen-Ching; Lee, Sam

    2012-01-01

    Icing scaling tests in the NASA Glenn Icing Research Tunnel (IRT) were performed on swept wing models using existing recommended scaling methods that were originally developed for straight wing. Some needed modifications on the stagnation-point local collection efficiency (i.e., beta(sub 0) calculation and the corresponding convective heat transfer coefficient for swept NACA 0012 airfoil models have been studied and reported in 2009, and the correlations will be used in the current study. The reference tests used a 91.4-cm chord, 152.4-cm span, adjustable sweep airfoil model of NACA 0012 profile at velocities of 100 and 150 knot and MVD of 44 and 93 mm. Scale-to-reference model size ratio was 1:2.4. All tests were conducted at 0deg angle of attack (AoA) and 45deg sweep angle. Ice shape comparison results were presented for stagnation-point freezing fractions in the range of 0.4 to 1.0. Preliminary results showed that good scaling was achieved for the conditions test by using the modified scaling methods developed for swept wing icing.

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

    NASA Technical Reports Server (NTRS)

    Johnston, Patrick J.

    1959-01-01

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

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

    1986-01-01

    Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols 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 clasical usnteady 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.

  11. 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.

  12. Effects of mean flow convection, quadrupole sources and vortex shedding on airfoil overall sound pressure level

    NASA Astrophysics Data System (ADS)

    Wolf, William R.; Azevedo, João L. F.; Lele, Sanjiva K.

    2013-12-01

    This paper presents a further analysis of results of airfoil self-noise prediction obtained in the previous work using large eddy simulation and acoustic analogy. The physical mechanisms responsible for airfoil noise generation in the aerodynamic flows analyzed are a combination of turbulent and laminar boundary layers, as well as vortex shedding (VS) originated due to trailing edge bluntness. The primary interest here consists of evaluating the effects of mean flow convection, quadrupole sources and vortex shedding tonal noise on the overall sound pressure level (OASPL) of a NACA0012 airfoil at low and moderate freestream Mach numbers. The overall sound pressure level is the measured quantity which eventually would be the main concern in terms of noise generation for aircraft and wind energy companies, and regulating agencies. The Reynolds number based on the airfoil chord is fixed at Rec=408,000 for all flow configurations studied. The results demonstrate that, for moderate Mach numbers, mean flow effects and quadrupole sources considerably increase OASPL and, therefore, should be taken into account in the acoustic prediction. For a low Mach number flow with vortex shedding, it is observed that OASPL is higher when laminar boundary layer separation is the VS driving mechanism compared to trailing edge bluntness.

  13. Assessment of Control Volume Estimation of Thrust for a Sinusoidally Pitching Airfoil at Low Reynolds Number

    NASA Astrophysics Data System (ADS)

    Hammer, Patrick; Naguib, Ahmed; Koochesfahani, Manoochehr

    2015-11-01

    The proper estimation of thrust is very important for understanding the aerodynamics of oscillating airfoils at low chord Reynolds number Re. Although direct force measurement is possible, force values at low Re are often small, and separation of the test-model's inertia forces from the data may not be straightforward. A common alternative is a control-volume (CV) approach, where terms in the integral momentum equation are computed from measured wake velocity profiles. Although it is acceptable to use only the mean streamwise-velocity profile in estimating the streamwise force on stationary airfoils, recent work has highlighted the importance of terms relating the velocity fluctuation and pressure distribution in the wake for unsteady airfoils. The goal of the present work is to capitalize on 2D computational data for a harmonically pitching airfoil at Re in the range 2,000-22,000, where all terms in the momentum-integral equation are accessible, to evaluate the importance of the various terms in the equation and assess the accuracy of the assumptions that are typically made in experiments due to the difficulty in measuring certain terms (such as the wake pressure distribution) by comparing the CV results with the actual computed thrust. This work was supported by AFOSR grant number FA9550-10-1-0342.

  14. 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)

  15. 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.

  16. Langley airfoil-research program

    NASA Technical Reports Server (NTRS)

    Bobbitt, P. J.

    1979-01-01

    An overview of past, present, and future airfoil research activities at the Langley Research Center is given. The immediate past and future occupy most of the discussion; however, past accomplishments and milestones going back to the early NACA years are dealt with in a broad-brush way to give a better perspective of current developments and programs. In addition to the historical perspective, a short description of the facilities which are now being used in the airfoil program is given. This is followed by a discussion of airfoil developments, advances in airfoil design and analysis tools (mostly those that have taken place over the past 5 or 6 years), and tunnel-wall-interference predictive methods and measurements. Future research requirements are treated.

  17. 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.

  18. Aerodynamic, aeroacoustic, and aeroelastic investigations of airfoil-vortex interaction using large-eddy simulation

    NASA Astrophysics Data System (ADS)

    Ilie, Marcel

    In helicopters, vortices (generated at the tip of the rotor blades) interact with the next advancing blades during certain flight and manoeuvring conditions, generating undesirable levels of acoustic noise and vibration. These Blade-Vortex Interactions (BVIs), which may cause the most disturbing acoustic noise, normally occur in descent or high-speed forward flight. Acoustic noise characterization (and potential reduction) is one the areas generating intensive research interest to the rotorcraft industry. Since experimental investigations of BVI are extremely costly, some insights into the BVI or AVI (2-D Airfoil-Vortex Interaction) can be gained using Computational Fluid Dynamics (CFD) numerical simulations. Numerical simulation of BVI or AVI has been of interest to CFD for many years. There are still difficulties concerning an accurate numerical prediction of BVI. One of the main issues is the inherent dissipation of CFD turbulence models, which severely affects the preservation of the vortex characteristics. Moreover this is not an issue only for aerodynamic and aeroacoustic analysis but also for aeroelastic investigations as well, especially when the strong (two-way) aeroelastic coupling is of interest. The present investigation concentrates mainly on AVI simulations. The simulations are performed for Mach number, Ma = 0.3, resulting in a Reynolds number, Re = 1.3 x 106, which is based on the chord, c, of the airfoil (NACA0012). Extensive literature search has indicated that the present work represents the first comprehensive investigation of AVI using the LES numerical approach, in the rotorcraft research community. The major factor affecting the aerodynamic coefficients and aeroacoustic field as a result of airfoil-vortex interaction is observed to be the unsteady pressure generated at the location of the interaction. The present numerical results show that the aerodynamic coefficients (lift, moment, and drag) and aeroacoustic field are strongly dependent on

  19. Subsonic natural-laminar-flow airfoils

    NASA Technical Reports Server (NTRS)

    Somers, Dan M.

    1992-01-01

    An account is given of the development history of natural laminar-flow (NLF) airfoil profiles under guidance of an experimentally well-verified theoretical method for the design of airfoils suited to virtually all subcritical applications. This method, the Eppler Airfoil Design and Analysis Program, contains a conformal-mapping method for airfoils having prescribed velocity-distribution characteristics, as well as a panel method for the analysis of potential flow about given airfoils and a boundary-layer method. Several of the NLF airfoils thus obtained are discussed.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

  1. Chords and harmonies in mixed optical and acoustical stimuli

    NASA Astrophysics Data System (ADS)

    Hahlweg, Cornelius; Dannenberg, Florian; Dörfler, Joachim; Weber, Bernhard; Weyer, Cornelia; Gercke-Hahn, Harald; Freimuth, Steffen; Heucke, Sören; Gutzmann, Holger Ludwig

    2014-09-01

    The paper is a follow up of the work presented in last year's Optics and Music session on the perception of coherence between low frequency power modulated light and periodical acoustic stimuli. The composition of chords and harmonies from power modulated light sources and their effect as stand-alone stimulus and in conjunction with the equivalent acoustic signal is discussed. Of special interest here is the modulation near perceptible flicker frequency. The substitution of acoustical chord components by their optical counterpart and vice versa is investigated. Further, concepts of a training application for trombone players and other instrumentalists are presented: since the mean slide of the trombone does not have fixed positions, the note must be found and two players might influence each other. The possibility of helping them to synchronize by optical stimuli derived from their playing is investigated. Beside possible applications in emotional reinforcing multimedia oriented entertainment and training support for musicians, again implications for occupational medicine are discussed.

  2. Chord-wise Tip Actuation on Flexible Flapping Plates

    NASA Astrophysics Data System (ADS)

    Martin, Nathan; Gharib, Morteza

    2015-11-01

    The aerodynamic characteristics of low aspect ratio flapping plates are strongly influenced by the interaction between tip and edge vortices. This has led to the development of tip actuation mechanisms which bend the tip towards the root of the plate in the span-wise direction during oscillation to investigate its impact. In our current work, a tip actuation mechanism to bend a flat plate's two free corners towards one another in the chord-wise direction is developed using a shape memory alloy. The aerodynamic forces and resulting flow field are investigated from dynamically altering the tip chord-wise curvature while flapping. The frequency of oscillation, stroke angle, flexibility, and tip actuation timing are independently varied to determine their individual effects. These results will further the fundamental understanding of flapping wing aerodynamics. This material is based upon work supported by the National Science Foundation Graduate Research Fellowship under Grant No. DGE 1144469.

  3. Closeup view showing portion of continuous bottom chord of truss ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    Close-up view showing portion of continuous bottom chord of truss with other web members and posts of the truss connected thereto at a joint by the use of a large steel pin. Note: The timber ties supporting the track (not shown but above) span transversely from truss to truss which are on 16' -0 centers. - Bridgeport Swing Span Bridge, Spanning Tennessee River, Bridgeport, Jackson County, AL

  4. 20. 80 foot pony truss an upper chord pin ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    20. 80 foot pony truss - an upper chord pin connection at a vertical post other than at the end post. Common to the five 80 foot trusses and similar to the 64 foot truss, there are two pairs per 80 foot truss and one pair on the 64 foot truss for a total of 22. - Weidemeyer Bridge, Spanning Thomes Creek at Rawson Road, Corning, Tehama County, CA

  5. High Reynolds number tests of a Douglas DLBA 032 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, Charles B.; Dress, David A.; Hill, Acquilla S.; Wilcox, Peter A.; Bui, Minh H.

    1986-01-01

    A wind-tunnel investigation of a Douglas advanced-technology airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). The temperature was varied from 227 K (409 R) to 100 K (180 R) at pressures ranging from about 159 kPa (1.57 atm) to about 514 kPa (5.07 atm). Mach number was varied from 0.50 to 0.78. These variables provided a Reynolds number range (based on airfoil chord) from 6.0 to 30.0 x 10 to the 6th power. This investigation was specifically designed to: (1) test a Douglas airfoil from moderately low to flight-equivalent Reynolds numbers, and (2) evaluate sidewall-boundary-layer effects on transonic airfoil performance characteristics by a systematic variation of Mach number, Reynolds number, and sidewall-boundary-layer removal. Data are included which demonstrate the effects of fixing transition, Mach number, Reynolds number, and sidewall-boundary-layer removal on the aerodynamic characteristics of the airfoil. Also included are remarks on model design and model structural integrity.

  6. Effect of initial acceleration on the development of the flow field of an airfoil pitching at constant rate

    NASA Technical Reports Server (NTRS)

    Koochesfahani, M. M.; Smiljanovski, V.; Brown, T. A.

    1992-01-01

    We present results from a series of experiments where an airfoil is pitched at constant rate from 0 to 60 degrees angle of attack. It is well documented that the dynamic stall behavior of such an airfoil strongly depends on the nondimensional pitch rate K = dot-alpha C/(2U(sub infinity)), where C is the chord, dot-alpha the constant pitch rate, and U(sub infinity) the free stream speed. In reality, the actual motion of the airfoil deviates from the ideal ramp due to the finite acceleration and deceleration periods imposed by the damping of drive system and response characteristics of the airfoil. It is possible that the pitch rate alone may not suffice in describing the flow and that the details of the motion trajectory before achieving a desired constant pitch rate may also affect the processes involved in the dynamic stall phenomenon. The effects of acceleration and deceleration periods are investigated by systematically varing the acceleration magnitude and its duration through the initial acceleration phase to constant pitch rate. The magnitude and duration of deceleration needed to bring the airfoil motion to rest is similarly controlled.

  7. Absolute and relative pitch: Global versus local processing of chords

    PubMed Central

    Ziv, Naomi; Radin, Shulamit

    2014-01-01

    Absolute pitch (AP) is the ability to identify or produce notes without any reference note. An ongoing debate exists regarding the benefits or disadvantages of AP in processing music. One of the main issues in this context is whether the categorical perception of pitch in AP possessors may interfere in processing tasks requiring relative pitch (RP). Previous studies, focusing mainly on melodic and interval perception, have obtained inconsistent results. The aim of the present study was to examine the effect of AP and RP separately, using isolated chords. Seventy-three musicians were categorized into four groups of high and low AP and RP, and were tested on two tasks: identifying chord types (Task 1), and identifying a single note within a chord (Task 2). A main effect of RP on Task 1 and an interaction between AP and RP in reaction times were found. On Task 2 main effects of AP and RP, and an interaction were found, with highest performance in participants with both high AP and RP. Results suggest that AP and RP should be regarded as two different abilities, and that AP may slow down reaction times for tasks requiring global processing. PMID:24855499

  8. Absolute and relative pitch: Global versus local processing of chords.

    PubMed

    Ziv, Naomi; Radin, Shulamit

    2014-01-01

    Absolute pitch (AP) is the ability to identify or produce notes without any reference note. An ongoing debate exists regarding the benefits or disadvantages of AP in processing music. One of the main issues in this context is whether the categorical perception of pitch in AP possessors may interfere in processing tasks requiring relative pitch (RP). Previous studies, focusing mainly on melodic and interval perception, have obtained inconsistent results. The aim of the present study was to examine the effect of AP and RP separately, using isolated chords. Seventy-three musicians were categorized into four groups of high and low AP and RP, and were tested on two tasks: identifying chord types (Task 1), and identifying a single note within a chord (Task 2). A main effect of RP on Task 1 and an interaction between AP and RP in reaction times were found. On Task 2 main effects of AP and RP, and an interaction were found, with highest performance in participants with both high AP and RP. Results suggest that AP and RP should be regarded as two different abilities, and that AP may slow down reaction times for tasks requiring global processing. PMID:24855499

  9. 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.

  10. Preliminary large-eddy simulations of flow around a NACA 4412 airfoil using unstructured grids

    NASA Technical Reports Server (NTRS)

    Jansen, Kenneth

    1995-01-01

    Large-eddy simulation (LES) has matured to the point where application to complex flows is desirable. The extension to higher Reynolds numbers leads to an impractical number of grid points with existing structured-grid methods. Furthermore, most real world flows are rather difficult to represent geometrically with structured grids. Unstructured-grid methods offer a release from both of these constraints. However, just as it took many years for structured-grid methods to be well understood and reliable tools for LES, unstructured-grid methods must be carefully studied before we can expect them to attain their full potential. In the past two years, important building blocks have been put into place making possible a careful study of LES on unstructured grids. The first building block was an efficient mesh generator which allowed the placement of points according to smooth variation of physical length scales. This variation of length scales is in all three directions independently, which allows a large reduction in points when compared to structured-grid methods, which can only vary length scales in one direction at a time. The second building block was the development of a dynamic model appropriate for unstructured grids. The principle obstacle was the development of an unstructured-grid filtering operator. In the past year, some of the new filters developed by Jansen have been implemented into a highly parallelized finite element code based on the Galerkin/least-squares finite element method. We have chosen the NACA 4412 airfoil at maximum lift as the first simulation for a variety of reasons. First, it is a problem of significant interest since it would be the first LES of an aircraft component. Second, this flow has been the subject of three experimental studies. The third reason for considering this flow is the variety of flow features which provide an important test of the dynamic model. Only the dynamic model can be expected to perform satisfactorily in this

  11. Inverse Design of a Thick Supercritical Airfoil

    NASA Astrophysics Data System (ADS)

    Pambagjo, Tjoetjoek Eko; Nakahashi, Kazuhiro; Obayashi, Shigeru

    In this paper, a study on designing a thick supercritical airfoil by utilizing Takanashi’s inverse design method is discussed. One of the problems to design a thick supercritical airfoil by Takanashi’s method is that an oscillation of the geometry may occur during the iteration process. To reduce the oscillation, an airfoil parameterization method is utilized as the smoothing procedure. A guideline to determine the target pressure distribution to realize the thick airfoil is also discussed.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  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. 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.

  18. 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

  19. Experimental study of full-scale iced-airfoil aerodynamic performance using sub-scale simulations

    NASA Astrophysics Data System (ADS)

    Busch, Greg T.

    Determining the aerodynamic effects of ice accretion on aircraft surfaces is an important step in aircraft design and certification. The goal of this work was to develop a complete sub-scale wind tunnel simulation methodology based on knowledge of the detailed iced-airfoil flowfield that allows the accurate measurement of aerodynamic penalties associated with the accretion of ice on an airfoil and to validate this methodology using full-scale iced-airfoil performance data obtained at near-flight Reynolds numbers. In earlier work, several classifications of ice shape were developed based on key aerodynamic features in the iced-airfoil flowfield: ice roughness, streamwise ice, horn ice, and tall and short spanwise-ridge ice. Castings of each of these classifications were acquired on a full-scale NACA 23012 airfoil model and the aero-dynamic performance of each was measured at a Reynolds number of 12.0 x 106 and a Mach number = 0.20. In the current study, sub-scale simple-geometry and 2-D smooth simulations of each of these castings were constructed based on knowledge of iced-airfoil flowfields. The effects of each simulation on the aerodynamic performance of an 18-inch chord NACA 23012 airfoil model was measured in the University of Illinois 3 x 4 ft. wind tunnel at a Reynolds number of 1.8 x 106 and a Mach number of 0.18 and compared with that measured for the corresponding full-scale casting at high Reynolds number. Geometrically-scaled simulations of the horn-ice and tall spanwise-ridge ice castings modeled C l,maxto within 2% and Cd,min to within 15%. Good qualitative agreement in the Cp distributions suggests that important geometric features such as horn and ridge height, surface location, and angle with respect to the airfoil chordline were appropriately modeled. Geometrically-scaled simulations of the ice roughness, streamwise ice, and short-ridge ice tended to have conservative C l,max and Cd. The aerodynamic performance of simulations of these types of

  20. High Reynolds number tests of a NASA SC(3)-0712(B) airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, W. G., Jr.; Hill, A. S.; Eichmann, O.

    1985-01-01

    A wind tunnel investigation of a NASA 12-percent-thick, advanced-technology supercritical airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT). This investigation represents another in the series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Tests program. Test temperature was varied from 220 K to 96 K at pressures ranging from 1.2 to 4.3 atm. Mach number was varied from 0.60 to 0.80. These variables provided a Reynolds number range from 4,400,000 to 40,000,000 based on a 15.24-cm (6.0-in.) airfoil chord. This investigation was designed to test a NASA advanced-technology airfoil from low to flight-equivalent Reynolds numbers, provide experience in cryogenic wind tunnel model design and testing techniques, and demonstrate the suitability of the 0.3-m TCT as an airfoil test facility. The aerodynamic results are presented as integrated force and moment coefficients and pressure distributions. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

  1. 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

  2. An exploratory investigation of the effect of a plastic coating on the profile drag of a practical-metal-construction sailplane airfoil

    NASA Technical Reports Server (NTRS)

    Somers, D. M.; Foster, J. M.

    1979-01-01

    The Langley low-turbulence pressure tunnel to determine the effect of a plastic coating on the profile drag of a practical-metal-construction sailplane airfoil was investigated. The model was tested with three surface configurations: (1) filled, painted, and sanded smooth; (2) rough bare metal; and (3) plastic-coated. The results are compared with data for the design airfoil (Wortmann FX 67-K-170/17) from another low-turbulence wind tunnel. The investigation was conducted at Reynolds numbers based on airfoil chord of 1.1 x 10 to the 6th power, 2.2 x 10 to the 6th power, and 3.3 x 10 to the 6th power at a Mach number of 0.10.

  3. Tests of a NACA 65(sub 1)-213 airfoil in the NASA Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Plentovich, E. B.; Ladson, C. L.; Hill, A. S.

    1984-01-01

    A wind-tunnel investigation was conducted to study the two dimensional aerodynamic characteristics of the NACA 65 sub 1-213 airfoil over a wide range of Reynolds numbers. Test temperature ranged from ambient to about 100K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from 0.22 to 0.80 and Reynolds number (based on airfoil chord) from 3 million to 40 million. Data are included which demonstrate the effects of fixed transition, Mach number, and Reynolds number on the aerodynamic characteristics of the airfoil. A sample of data showing the effects of angle of attack on the pressure distribution is also given. The data are presented in an uncorrected form with no analysis.

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

    SciTech Connect

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

    1995-01-01

    The objective of this wind-tunnel test was to verify the predictions of the Eppler Airfoil Design and Analysis Code for a very thick airfoil having a high maximum lift coefficient (c{sub 1,max} designed to be largely insensitive to leading edge roughness effects. The 24-percent-thick S814 airfoil was designed with these characteristics to accommodate aerodynamic and structural considerations for the root region of a wind-turbine blade. In addition, the airfoil`s maximum lift-to-drag ratio was designed to occur it a high lift coefficient. To accomplish the objective, a two-dimensional wind-tunnel test of the S814 thick root airfog was conducted in January 1994 in the low-turbulence wind tunnel of the Delft University of Technology Low Speed Laboratory. Data were obtained for transition-free and transition-fixed conditions at Reynolds numbers of 0.7, 1.0, 1.5, 2.0, and 3.0 {times} 10{sup 6}. For the design Reynolds numbers of 1.5 {times} l0{sup 6}, the transition-free c{sub 1,max} is 1.3 which satisfies the design specification. However, this value is significantly lower than the predicted c{sub 1,max} of almost l.6. With transition-fixed at the is 1.2. The difference in c{sub 1,max} between the transition-free and transition-fixed conditions demonstrates the airfoil`s minimal sensitivity to roughness effects. The S814 root airfoil was designed to complement existing NREL low c{sub 1,max} tip-region airfoils for rotor blades 10 to 15 meters in length.

  5. 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.

  6. Hook nozzle arrangement for supporting airfoil vanes

    DOEpatents

    Shaffer, James E.; Norton, Paul F.

    1996-01-01

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

  7. An airfoil design method for viscous flows

    NASA Technical Reports Server (NTRS)

    Malone, J. B.; Narramore, J. C.; Sankar, L. N.

    1990-01-01

    An airfoil design procedure is described that has been incorporated into an existing two-dimensional Navier-Stokes airfoil analysis method. The resulting design method, an iterative procedure based on a residual-correction algorithm, permits the automated design of airfoil sections with prescribed surface pressure distributions. This paper describes the inverse design method and the technique used to specify target pressure distributions. An example airfoil design problem is described to demonstrate application of the inverse design procedure. It shows that this inverse design method develops useful airfoil configurations with a reasonable expenditure of computer resources.

  8. Airfoil shape for a turbine nozzle

    SciTech Connect

    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.

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

    NASA Astrophysics Data System (ADS)

    Karges, Adam T.

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

  10. Newborn infants' auditory system is sensitive to Western music chord categories

    PubMed Central

    Virtala, Paula; Huotilainen, Minna; Partanen, Eino; Fellman, Vineta; Tervaniemi, Mari

    2013-01-01

    Neural encoding of abstract rules in the audition of newborn infants has been recently demonstrated in several studies using event-related potentials (ERPs). In the present study the neural encoding of Western music chords was investigated in newborn infants. Using ERPs, we examined whether the categorizations of major vs. minor and consonance vs. dissonance are present at the level of the change-related mismatch response (MMR). Using an oddball paradigm, root minor, dissonant and inverted major chords were presented in a context of consonant root major chords. The chords were transposed to several different frequency levels, so that the deviant chords did not include a physically deviant frequency that could result in an MMR without categorization. The results show that the newborn infants were sensitive to both dissonant and minor chords but not to inverted major chords in the context of consonant root major chords. While the dissonant chords elicited a large positive MMR, the minor chords elicited a negative MMR. This indicates that the two categories were processed differently. The results suggest newborn infants are sensitive to Western music categorizations, which is consistent with the authors' previous studies in adults and school-aged children. PMID:23966962

  11. Preliminary Investigation of Cyclic De-Icing of an Airfoil Using an External Electric Heater

    NASA Technical Reports Server (NTRS)

    Lewis, James P.; Bowden, Dean T.

    1952-01-01

    An investigation was conducted in the NACA Lewis icing research tunnel to determine the characteristics and requirements of cyclic deicing of a 65,2-216 airfoil by use of an external electric heater. The present investigation was limited to an airspeed of 175 miles per hour. Data are presented to show the effects of variations in heat-on and heat-off periods, ambient air temperature, liquid-water content, angle of attack, and. heating distribution on the requirements for cyclic deicing. The external heat flow at various icing and heating conditions is also presented. A continuously heated parting strip at the airfoil leading edge was found necessary for quick, complete, and consistent ice removal. The cyclic power requirements were found to be primarily a function of the datum temperature and heat-on time, with the other operating and meteorological variables having a second-order effect. Short heat-on periods and high power densities resulted in the most efficient ice removal, the minimum energy input, and the minimum runback ice formations. The optimum chordwise heating distribution pattern was found to consist of a uniform distribution of cycled power density in the impingement region. Downstream of the impingement region the power density decreased to the limits of heating which, for the conditions investigated, extended from 5.7 percent chord on the upper surface of the airfoil to 8.9 percent chord on the lower surface. Ice removal did not take place at a heater surface temperature of 32 F; surface temperatures of approximately 50 to 100 F were required to effect removal. Better de-icing performance and greater energy savings would be possible with a heater having a higher thermal efficiency.

  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. Pneumatic Spoiler Controls Airfoil Lift

    NASA Technical Reports Server (NTRS)

    Hunter, D.; Krauss, T.

    1991-01-01

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

  14. Turbulence model evaluation for the prediction of flows over a supercritical airfoil with deflected aileron at high Reynolds number

    NASA Technical Reports Server (NTRS)

    Londenberg, W. K.

    1993-01-01

    Navier-Stokes solutions about a supercritical airfoil with aileron deflection have been computed using the CFL3D code coupled with the Baldwin-Lomax, Johnson-King, Baldwin-Barth, and Spalart-Allmaras turbulence models. Computations were made at a Mach number of 0.716 and chord Reynolds numbers of 5, 15, and 25 million. The airfoil was analyzed with both 0 deg and 2 deg (TED) aileron deflections. Comparisons over a range of angles-of-attack showed that solutions obtained using the Baldwin-Barth turbulence model presented the best agreement with experimental pressures and sectional lift coefficients. However, Reynolds number trends in sectional lift coefficient and in aileron effectiveness were not predicted consistently.

  15. Post-stall wind tunnel data for NACA 44XX series airfoil sections

    SciTech Connect

    Ostowari, C.; Naik, D.

    1985-01-01

    Wind turbine blades operate over a wide angle of attach range. Unlike aircraft, a wind turbine's angle of attach range extends deep into stall where the three-dimensional performance characteristics of airfoils are not generally known. Peak power predictions upon which wind turbine components are sized depend on a good understanding of a blade's post-stall characteristics. The purpose of this wind tunnel study is to characterize the performance characteristics of a blade in stall as a function of its aspect ratio, airfoil thickness, and Reynolds number. This report documents results of the wind tunnel investigation of constant chord blades having four aspect ratios, with NACA 44XX series airfoil sections, at angles of attack ranging from -10/sup 0/ to 110/sup 0/. Tests were conducted at Reynolds number ranging from 0.25 x 106 to 1.0 x 106. The thickness ratios studied were 0.18, 0.15, and 0.12, and 0.09. The aspect ratios were 6, 9, 12 and infinity. Results of force and pitching moment measurements over the angle of attack range for all combinations of Reynolds numbers, thickness, and aspect ratios, and the effects of boundary layer tripping are presented.

  16. Evaluation of tunnel sidewall boundary-layer-control systems for high-lift airfoil testing

    NASA Technical Reports Server (NTRS)

    Paschal, K.; Goodman, W.; Mcghee, R.; Walker, B.; Wilcox, Peter A.

    1991-01-01

    An experimental study was conducted in the NASA Langley Low-Turbulence Pressure Tunnel to evaluate a suction sidewall boundary-layer-control (BLC) technique used in testing 2D high-lift airfoils. Sidewall BLC is required to maintain spanwise two-dimensionality of the flow over the airfoil at large angles of attack. A supercritical-type high-lift air-foil, equipped with a double-slotted flap and a leading-edge slat, was used for the study which was conducted at a Mach number of 0.2 and Reynolds numbers based on chord of 9 and 16 million. The sidewall BLC technique, which features distributed suction through porous endplates connected to a venting system, was able to control sidewall boundary-layer separation and maintain two-dimensional flow over the high-lift configuration for both Reynolds numbers tested. Discussions on porous endplate optimization and effects of suction on section lift are presented. Results obtained with the suction system were also compared with previous data obtained with a tangential blowing BLC system for the same high-lift configuration.

  17. Interaction of an Artificially Thickened Boundary Layer with a Vertically Mounted Pitching Airfoil

    NASA Astrophysics Data System (ADS)

    Hohman, Tristen; Smits, Alexander; Martinelli, Luigi

    2011-11-01

    Wind energy represents a large portion of the growing market in alternative energy technologies and the current landscape has been dominated by the more prevalent horizontal axis wind turbine. However, there are several advantages to the vertical axis wind turbine (VAWT) or Darrieus type design and yet there is much to be understood about how the atmospheric boundary layer (ABL) affects their performance. In this study the ABL was simulated in a wind tunnel through the use of elliptical shaped vortex generators, a castellated wall, and floor roughness elements as described in the method of Counihan (1967) and then verified its validity by hot wire measurement of the mean velocity profile as well as the turbulence intensity. The motion of an blade element around a vertical axis is approximated through the use of a pitching airfoil. The wake of the airfoil is investigated through hot wire anemometry in both uniform flow and in the simulated boundary layer both at Re = 1 . 37 ×105 based on the chord of the airfoil. Sponsored by Hopewell Wind Power (Hong Kong) Limited.

  18. Experimental Investigation of Dynamic Stall on a NACA0012 Airfoil Undergoing Sinusoidal Pitching

    NASA Astrophysics Data System (ADS)

    Bohl, Douglas; Green, Melissa

    2015-11-01

    In this work, the flow field around a NACA0012 Airfoil undergoing large amplitude sinusoidal pitching is investigated using Particle Image Velocimetry (PIV). The airfoil is pitched symmetrically about the quarter chord point with a peak angle of 20 deg, at reduced frequencies of k =0.2-0.6 and Rec = 12000. Sixteen different Fields of View are phase averaged and combined to quantify the flow field from 0.75c upstream of the leading edge to 1c downstream of the trailing edge. This provides spatially and temporally resolved data sets that include the downstream evolution of the flow fields. The velocity and vorticity fields, both around the airfoil and downstream of the trailing edge, will be investigated as a function of the reduced frequency to better understand the dynamics (i.e. formation, separation and development) of the leading edge vortex and the resulting downstream flow evolution. This work was supported by the Office of Naval Research under ONR Award No. N00014-14-1-0418.

  19. Active Flow Control at Low Reynolds Numbers on a NACA 0015 Airfoil

    NASA Technical Reports Server (NTRS)

    Melton, LaTunia Pack; Hannon, Judith; Yao, Chung-Sheng; Harris, Jerome

    2008-01-01

    Results from a low Reynolds number wind tunnel experiment on a NACA 0015 airfoil with a 30% chord trailing edge flap tested at deflection angles of 0, 20, and 40 are presented and discussed. Zero net mass flux periodic excitation was applied at the ap shoulder to control flow separation for flap deflections larger than 0. The primary objective of the experiment was to compare force and moment data obtained from integrating surface pressures to data obtained from a 5-component strain-gage balance in preparation for additional three-dimensional testing of the model. To achieve this objective, active flow control is applied at an angle of attack of 6 where published results indicate that oscillatory momentum coefficients exceeding 1% are required to delay separation. Periodic excitation with an oscillatory momentum coefficient of 1.5% and a reduced frequency of 0.71 caused a significant delay of separation on the airfoil with a flap deflection of 20. Higher momentum coefficients at the same reduced frequency were required to achieve a similar level of flow attachment on the airfoil with a flap deflection of 40. There was a favorable comparison between the balance and integrated pressure force and moment results.

  20. Flow around a NACA0018 airfoil with a cavity and its dynamical response to acoustic forcing

    NASA Astrophysics Data System (ADS)

    Olsman, W. F. J.; Willems, J. F. H.; Hirschberg, A.; Colonius, T.; Trieling, R. R.

    2011-08-01

    Trapping of vortices in a cavity has been explored in recent years as a drag reduction measure for thick airfoils. If, however, trapping fails, then oscillation of the cavity flow may couple with elastic vibration modes of the airfoil. To examine this scenario, the effect of small amplitude vertical motion on the oscillation of the shear layer above the cavity is studied by acoustic forcing simulating a vertical translation of a modified NACA0018 profile. At low Reynolds numbers based on the chord ( O(104)), natural instability modes of this shear layer are observed for Strouhal numbers based on the cavity width of order unity. Acoustic forcing sufficiently close to the natural instability frequency induces a strong non-linear response due to lock-in of the shear layer. At higher Reynolds numbers (above 105) for Strouhal number 0.6 or lower, no natural instabilities of the shear layer and only a linear response to forcing were observed. The dynamical pressure difference across the airfoil is then dominated by added mass effects, as was confirmed by numerical simulations.

  1. Turbulent separated flow in the vicinity of a single-slotted airfoil flap

    NASA Technical Reports Server (NTRS)

    Adair, Desmond; Horne, W. Clifton

    1988-01-01

    Detailed measurements of pressure and velocity characteristics are presented and analyzed for flow over and downstream of a NACA 4412 airfoil equipped with a NACA 4415 single-slotted flap at high angle of attack and close to maximum lift. The flow remained attached over the main element while a large region of recirculating flow occurred over the aft 61 percent of the flap. The airfoil configuration was tested at a Mach number of 0.09 and chord based Reynolds number of 1.8 x 10 to the 6th in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. Measurements of mean and fluctuating velocities were obtained in the region of recirculation and high turbulence intensity using three-dimensional laser velocimetry. In regions where the flow had a preferred direction and relatively low turbulence intensity hot-wire anemometry was used. Emphasis was placed on obtaining flow characteristics in the confluent boundary layer, the region of recirculating flow and in the downstream wake. Surface pressure measurements were made on the main airfoil, flap, wind tunnel roof, and wind tunnel floor. In addition to the presentation of pressure and velocity characteristics, the near wall results inside the separated region are analyzed as are the relative importance of terms in the momentum and turbulence kinetic energy equations in the confluent separated boundary layer and the recirculating region of the near wake.

  2. A Wind Tunnel Study of Icing Effects on a Business Jet Airfoil

    NASA Technical Reports Server (NTRS)

    Addy, Harold E., Jr.; Broeren, Andy P.; Zoeckler, Joesph G.; Lee, Sam

    2003-01-01

    Aerodynamic wind tunnel tests were conducted to study the effects of various ice accretions on the aerodynamic performance of a 36-inch chord, two-dimensional business jet airfoil. Eight different ice shape configurations were tested. Four were castings made from molds of ice shapes accreted in an icing wind tunnel. Two were made using computationally smoothed tracings of two of the ice shapes accreted in the icing tunnel. These smoothed profiles were then extended in the spanwise direction to form a two-dimensional ice shape. The final two configurations were formed by applying grit to the smoothed ice shapes. The ice shapes resulted in as much as 48% reduction in maximum lift coefficient from that of the clean airfoil. Large increases in drag and changes in pitching moment were also observed. The castings and their corresponding smoothed counterparts yielded similar results. Little change in performance was observed with the addition of grit to the smoothed ice shapes. Changes in the Reynolds number (from 3 x 10(exp 6) to 10.5 x 10(exp 6) and Mach number (from 0.12 to 0.28) did not significantly affect the iced-airfoil performance coefficients.

  3. Effect of Chord Size on Weight and Cooling Characteristics of Air-Cooled Turbine Blades

    NASA Technical Reports Server (NTRS)

    Esgar, Jack B; Schum, Eugene F; Curren, Arthur N

    1958-01-01

    An analysis has been made to determine the effect of chord size on the weight and cooling characteristics of shell-supported, air-cooled gas-turbine blades. In uncooled turbines with solid blades, the general practice has been to design turbines with high aspect ratio (small blade chord) to achieve substantial turbine weight reduction. With air-cooled blades, this study shows that turbine blade weight is affected to a much smaller degree by the size of the blade chord.

  4. Multi-chord fiber-coupled interferometry of supersonic plasma jets andcomparisons with synthetic data

    SciTech Connect

    Merritt, Elizabeth C.; Lynn, Alan G.; Gilmore, Mark A.; Thoma, Carsten; Loverich, John; Hsu, Scott C.

    2012-05-03

    A multi-chord fiber-coupled interferometer [Merritt et al., Rev. Sci. Instrum. 83, 033506 (2012)] is being used to make time-resolved density measurements of supersonic argon plasma jets on the Plasma Liner Experiment [Hsu et al., Bull. Amer. Phys. Soc. 56, 307 (2011)]. The long coherence length of the laser (> 10 m) allows signal and reference path lengths to be mismatched by many meters without signal degradation, making for a greatly simplified optical layout. Measured interferometry phase shifts are consistent with a partially ionized plasma in which an initially positive phase shift becomes negative when the ionization fraction drops below a certain threshold. In this case, both free electrons and bound electrons in ions and neutral atoms contribute to the index of refraction. This paper illustrates how the interferometry data, aided by numerical modeling, are used to derive total jet density, jet propagation velocity ({approx} 15-50 km/s), jet length ({approx} 20-100 cm), and 3D expansion.

  5. Musical chords and emotion: major and minor triads are processed for emotion.

    PubMed

    Bakker, David Radford; Martin, Frances Heritage

    2015-03-01

    Musical chords are arguably the smallest building blocks of music that retain emotional information. Major chords are generally perceived as positive- and minor chords as negative-sounding, but there has been debate concerning how early these emotional connotations may be processed. To investigate this, emotional facial stimuli and musical chord stimuli were simultaneously presented to participants, and facilitation of processing was measured via event-related potential (ERP) amplitudes. Decreased amplitudes of the P1 and N2 ERP components have been found to index the facilitation of early processing. If simultaneously presented musical chords and facial stimuli are perceived at early stages as belonging to the same emotional category, then early processing should be facilitated for these congruent pairs, and ERP amplitudes should therefore be decreased as compared to the incongruent pairs. ERPs were recorded from 30 musically naive participants as they viewed happy, sad, and neutral faces presented simultaneously with a major or minor chord. When faces and chords were presented that contained congruent emotional information (happy-major or sad-minor), processing was facilitated, as indexed by decreased N2 ERP amplitudes. This suggests that musical chords do possess emotional connotations that can be processed as early as 200 ms in naive listeners. The early stages of processing that are involved suggest that major and minor chords have deeply connected emotional meanings, rather than superficially attributed ones, indicating that minor triads possess negative emotional connotations and major triads possess positive emotional connotations. PMID:24957406

  6. 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

  7. Heat transfer measurements from a NACA 0012 airfoil in flight and in the NASA Lewis icing research tunnel. M.S. Thesis Final Report

    NASA Technical Reports Server (NTRS)

    Poinsatte, Philip E.

    1990-01-01

    Local heat transfer coefficients from a smooth and roughened NACA 0012 airfoil were measured using a steady state heat flux method. Heat transfer measurements on the specially constructed 0.533 meter chord airfoil were made both in flight on the NASA Lewis Twin Otter Research Aircraft and in the NASA Lewis Icing Research Tunnel (IRT). Roughness was obtained by the attachment of small, 2 mm diameter, hemispheres of uniform size to the airfoil surface in four distinct patterns. The flight data was taken for the smooth and roughened airfoil at various Reynolds numbers based on chord in the range of 1.24x10(exp 6) to 2.50x10(exp 6) and at various angles of attack up to 4 degrees. During these flight tests the free stream velocity turbulence intensity was found to be very low (less than 0.1 percent). The wind tunnel data was taken in the Reynolds number range of 1.20x10(exp 6) to 4.52x10(exp 6) and at angles of attack from -4 degrees to +8 degrees. The turbulence intensity in the IRT was 0.5 to 0.7 percent with the cloud making spray off. Results for both the flight and tunnel tests are presented as Frossling number based on chord versus position on the airfoil surface for various roughnesses and angle of attack. A table of power law curve fits of Nusselt number as a function of Reynolds number is also provided. The higher level of turbulence in the IRT versus flight had little effect on heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the higher Reynolds numbers. Turning on the cloud making spray air in the IRT did not alter the heat transfer. Roughness generally increased the heat transfer by locally disturbing the boundary layer flow. Finally, the present data was not only compared with previous airfoil data where applicable, but also with leading edge cylinder and flat plate heat transfer values which are often used to estimate airfoil heat transfer in computer codes.

  8. Effect of High-Fidelity Ice Accretion Simulations on the Performance of a Full-Scale Airfoil Model

    NASA Technical Reports Server (NTRS)

    Broeren, Andy P.; Bragg, Michael B.; Addy, Harold E., Jr.; Lee, Sam; Moens, Frederic; Guffond, Didier

    2010-01-01

    The simulation of ice accretion on a wing or other surface is often required for aerodynamic evaluation, particularly at small scale or low-Reynolds number. While there are commonly accepted practices for ice simulation, there are no established and validated guidelines. The purpose of this article is to report the results of an experimental study establishing a high-fidelity, full-scale, iced-airfoil aerodynamic performance database. This research was conducted as a part of a larger program with the goal of developing subscale aerodynamic simulation methods for iced airfoils. Airfoil performance testing was carried out at the ONERA F1 pressurized wind tunnel using a 72-in. (1828.8-mm) chord NACA 23012 airfoil over a Reynolds number range of 4.5x10(exp 6) to 16.0 10(exp 6) and a Mach number range of 0.10 to 0.28. The high-fidelity, ice-casting simulations had a significant impact on the aerodynamic performance. A spanwise-ridge ice shape resulted in a maximum lift coefficient of 0.56 compared to the clean value of 1.85 at Re = 15.9x10(exp 6) and M = 0.20. Two roughness and streamwise shapes yielded maximum lift values in the range of 1.09 to 1.28, which was a relatively small variation compared to the differences in the ice geometry. The stalling characteristics of the two roughness and one streamwise ice simulation maintained the abrupt leading-edge stall type of the clean NACA 23012 airfoil, despite the significant decrease in maximum lift. Changes in Reynolds and Mach number over the large range tested had little effect on the iced-airfoil performance.

  9. Measurement of local convective heat transfer coefficients from a smooth and roughened NACA-0012 airfoil: Flight test data

    NASA Technical Reports Server (NTRS)

    Newton, James E.; Vanfossen, G. James; Poinsatte, Phillip E.; Dewitt, Kenneth J.

    1988-01-01

    Wind tunnels typically have higher free stream turbulence levels than are found in flight. Turbulence intensity was measured to be 0.5 percent in the NASA Lewis Icing Research Tunnel (IRT) with the cloud making sprays off and around 2 percent with cloud making equipment on. Turbulence intensity for flight conditions was found to be too low to make meaningful measurements for smooth air. This difference between free stream and wing tunnel conditions has raised questions as to the validity of results obtained in the IRT. One objective of these tests was to determine the effect of free stream turbulence on convective heat transfer for the NASA Lewis LEWICE ice growth prediction code. These tests provide in-flight heat transfer data for a NASA-0012 airfoil with a 533 cm chord. Future tests will measure heat transfer data from the same airfoil in the Lewis Icing Research Tunnel. Roughness was obtained by the attachment of small, 2 mm diameter hemispheres of uniform size to the airfoil in three different patterns. Heat transfer measurements were recorded in flight on the NASA Lewis Twin Otter Icing Research Aircraft. Measurements were taken for the smooth and roughened surfaces at various aircraft speeds and angles of attack up to four degrees. Results are presented as Frossling number versus position on the airfoil for various roughnesses and angles of attack.

  10. Measurement of local convective heat transfer coefficients from a smooth and roughened NACA-0012 airfoil - Flight test data

    NASA Technical Reports Server (NTRS)

    Van Fossen, G. James; De Witt, Kenneth J.; Newton, James E.; Poinsatte, Phillip E.

    1988-01-01

    Wind tunnels typically have higher free stream turbulence levels than are found in flight. Turbulence intensity was measured to be 0.5 percent in the NASA Lewis Icing Research Tunnel (IRT) with the cloud making sprays off and around 2 percent with cloud making equipment on. Turbulence intensity for flight conditions was found to be too low to make meaningful measurements for smooth air. This difference between free stream and wind tunnel conditions has raised questions as to the validity of results obtained in the IRT. One objective of these tests was to determine the effect of free stream turbulence on convective heat transfer for the NASA Lewis LEWICE ice growth prediction code. These tests provide in-flight heat transfer data for a NASA-0012 airfoil with a 533 cm chord. Future tests will measure heat transfer data from the same airfoil in the Lewis Icing Research Tunnel. Roughness was obtained by the attachment of small, 2 mm diameter hemispheres of uniform size to the airfoil in three different patterns. Heat transfer measurements were recorded in flight on the NASA Lewis Twin Otter Icing Research Aircraft. Measurements were taken for the smooth and roughened surfaces at various aircraft speeds and angles of attack up to four degrees. Results are presented as Frossling number versus position on the airfoil for various roughnesses and angles of attack.

  11. 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.

  12. Dynamic Chord-wise Tip Curvature on Flexible Flapping Plates

    NASA Astrophysics Data System (ADS)

    Martin, Nathan; Gharib, Morteza

    2014-11-01

    The aerodynamic characteristics of long rectangular flapping plates are strongly influenced by the interaction between tip and edge vortices. This has led to the development of many tip actuation mechanisms to independently bend or rotate the tip towards the root of the plate in the span-wise direction. In our current work, the influence of dynamically altering the chord-wise curvature of the tip on the generation of aerodynamic forces is investigated; for this case, the two free corners of the flat plate bend towards each other. The parameters of actuation timing, maximum curvature, Reynolds number, flexibility, and tip speed are independently varied to determine their influence. These results will further the fundamental understanding of unsteady aerodynamics. This material is based upon work supported by the National Science Foundation Graduate Research Fellowship and the Gordon and Betty Moore foundation.

  13. High-Lift Separated Flow About Airfoils

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1982-01-01

    TRANSEP Calculates flow field about low-speed single-element airfoil at high-angle-of-attack and high-lift conditions with massive boundary-layer separation. TRANSEP includes effects of weak viscous interactions and can be used for subsonic/transonic airfoil design and analysis. The approach used in TRANSEP is based on direct-inverse method and its ability to use either displacement surface or pressure as airfoil boundary condition.

  14. 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.

  15. Airfoil seal system for gas turbine engine

    DOEpatents

    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.

  16. 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.

  17. Unsteady aerodynamics of conventional and supercritical airfoils

    NASA Technical Reports Server (NTRS)

    Davis, S. S.; Malcolm, G. N.

    1980-01-01

    The unsteady aerodynamics of a conventional and a supercritical airfoil are compared by examining measured chordwise unsteady pressure time-histories from four selected flow conditions. Although an oscillating supercritical airfoil excites more harmonics, the strength of the airfoil's shock wave is the more important parameter governing the complexity of the unsteady flow. Whether they are conventional or supercritical, airfoils that support weak shock waves induce unsteady loads that are qualitatively predictable with classical theories; flows with strong shock waves are sensitive to details of the shock-wave and boundary-layer interaction and cannot be adequately predicted.

  18. High Lift, Low Pitching Moment Airfoils

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W. (Inventor)

    1988-01-01

    Two families of airfoil sections which can be used for helicopter/rotorcraft rotor blades or aircraft propellers of a particular shape are prepared. An airfoil of either family is one which could be produced by the combination of a camber line and a thickness distribution or a thickness distribution which is scaled from these. An airfoil of either family has a unique and improved aerodynamic performance. The airfoils of either family are intended for use as inboard sections of a helicopter rotor blade or an aircraft propeller.

  19. Generalized multi-point inverse airfoil design

    NASA Technical Reports Server (NTRS)

    Selig, Michael S.; Maughmer, Mark D.

    1991-01-01

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

  20. 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.

  1. The further development of circulation control airfoils

    NASA Technical Reports Server (NTRS)

    Wood, N. J.

    1987-01-01

    The performance trends of circulation control airfoils are reviewed and observations are made as to where improvements in performance and expansion of the flight envelope may be feasible. A new analytically defined family of airfoils is suggested, all of which maintain the fore and aft symmetry required for stopped rotor application. It is important to recognize that any improvements in section capabilities may not be totally applicable to the present vehicle operation. It remains for the designers of the rotor system to reappraise the three dimensional operating environment in view of the different airfoil operating characteristics and for the airfoil definitions to be flexible while maintaining satisfactory levels of performance.

  2. 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.

  3. The Tonal Function of a Task-Irrelevant Chord Modulates Speed of Visual Processing

    ERIC Educational Resources Information Center

    Escoffier, N.; Tillmann, B.

    2008-01-01

    Harmonic priming studies have provided evidence that musical expectations influence sung phoneme monitoring, with facilitated processing for phonemes sung on tonally related (expected) chords in comparison to less-related (less-expected) chords [Bigand, Tillmann, Poulin, D'Adamo, and Madurell (2001). "The effect of harmonic context on phoneme…

  4. Measurement of Plasma Ion Temperature and Flow Velocity from Chord-Averaged Emission Line Profile

    NASA Astrophysics Data System (ADS)

    Wei, Xu

    2011-06-01

    The distinction between Doppler broadening and Doppler shift has been analysed, the differences between Gaussian fitting and the distribution of chord-integral line shape have also been discussed. Local ion temperature and flow velocity have been derived from the chord-averaged emission line profile by a chosen-point Gaussian fitting technique.

  5. 4. Monadnock Mill No. 1 compression post/top chord detail. A ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    4. Monadnock Mill No. 1 compression post/top chord detail. A mortice from an earlier framing member is visible as is a peg at the top chord where the finish millwork trim has been removed from the post. - Monadnock Mills, Mill No. 1, 13-17 Water Street, Claremont, Sullivan County, NH

  6. 10. Detail, lower chord connection, Bridge Number 324.99, view to ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    10. Detail, lower chord connection, Bridge Number 324.99, view to northeast, 210mm lens. Visible are the two lines of eyebars that form the lower chord, with a laced vertical intermediate post, paired diagonals at right, and a single counter at left. - Southern Pacific Railroad Shasta Route, Bridge No. 324.99, Milepost 324.99, Shasta Springs, Siskiyou County, CA

  7. Timbre influences chord discrimination in black-capped chickadees (Poecile atricapillus) but not humans (Homo sapiens).

    PubMed

    Hoeschele, Marisa; Cook, Robert G; Guillette, Lauren M; Hahn, Allison H; Sturdy, Christopher B

    2014-11-01

    Timbre is an important attribute of sound both in music and nature. Previously, using an operant conditioning paradigm, we found that black-capped chickadees and humans show similar response patterns in discriminating triadic chords of the same timbre and transferred this discrimination to a novel key center (novel absolute pitch). The current study examined how varying the timbre of the chords influenced discrimination. Using a similar operant conditioning procedure, we trained humans (Experiment 1) and chickadees (Experiments 2 and 3) to discriminate a major chord from 6 other chord types that had semitone deviations from the major chord. The pattern of errors of the 2 species replicated our previous findings. We then tested participants with novel timbres. We found that humans readily transferred their discrimination to novel timbres, suggesting they were attending to triadic pitch relations. The chickadees failed to transfer to novel timbres, suggesting they were using a different strategy to perform the original chord discrimination. We conducted an acoustic analysis examining frequency ranges that are biologically relevant to chickadees. We found that the relative intensity within each chord of the frequencies used in black-capped chickadee song significantly correlated with chickadees' percent response during probe testing. In Experiment 3, we trained a new set of chickadees by including either expanded pitch or timbre training before testing. Although chickadees showed some transfer to novel chords following this expanded training, we found that neither type of expanded training helped the chickadees when probe tested with novel stimuli. PMID:25150964

  8. Amygdala activity can be modulated by unexpected chord functions during music listening.

    PubMed

    Koelsch, Stefan; Fritz, Thomas; Schlaug, Gottfried

    2008-12-01

    Numerous earlier studies have investigated the cognitive processing of musical syntax with regular and irregular chord sequences. However, irregular sequences may also be perceived as unexpected, and therefore have a different emotional valence than regular sequences. We provide behavioral data showing that irregular chord functions presented in chord sequence paradigms are perceived as less pleasant than regular sequences. A reanalysis of functional MRI data showed increased blood oxygen level-dependent signal changes bilaterally in the amygdala in response to music-syntactically irregular (compared with regular) chord functions. The combined data indicate that music-syntactically irregular events elicit brain activity related to emotional processes, and that, in addition to intensely pleasurable music or highly unpleasant music, single chord functions can also modulate amygdala activity. PMID:19050462

  9. 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.

  10. 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.

  11. 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…

  12. 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.

  13. 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.

  14. Wind powered generator with cyclic airfoil latching

    SciTech Connect

    Bair, P.

    1981-12-01

    A wind powered generator rotatable about a vertical axis is described. A plurality of vertically disposed airfoils are provided, the airfoils being rotatable about a vertical axis parallel to the axis of the generator. The airfoils are selectively latched to be disposed perpendicularly of the wind direction during one phase of their revolution about the generator axis and are selectively unlatched to be permitted to rotate into a position generally parallel to the wind direction during other phases of their revolution. The latching and unlatching of the airfoils is determined by the wind direction and is effected by electronic means which determine the point of latching and unlatching as a function of the wind direction measured by a wind vane. The airfoils may comprise sails composed of a flexible material stretched into a predetermined shape on a frame.

  15. Transonic flow past an airfoil with condensation

    NASA Technical Reports Server (NTRS)

    Schmidt, B.

    1978-01-01

    In connection with investigations conducted to determine the influence of water vapor on experiments in wind tunnels, the question arose as to what changes due to vapor condensation might be expected in airfoil measurements. Density measurements on circular-arc airfoils aided by an interferometer in choked tunnels with parallel walls show that increasing humidity produces increasing changes in the flow field. The flow becomes nonstationary at high humidity. At the airfoil, however, the influence of the condensation is only felt, inasmuch as the shock bounding the local supersonic region moves upstream with increasing humidity while its intensity decreases. The density distribution upstream of the shock remains unchanged. Even if the flow becomes nonstationary in the vicinity of the airfoil, no changes occur at the airfoil.

  16. Viscous Transonic Airfoil Workshop compendium of results

    NASA Technical Reports Server (NTRS)

    Holst, Terry L.

    1987-01-01

    Results from the Viscous Transonic Airfoil Workshop held at the AIAA 25th Aerospace Sciences Meeting at Reno, NV in January 1987, are compared with each other and with experimental data. Test cases used in this workshop include attached and separated transonic flows for three different airfoils: the NACA 0012 airfoil, the RAE 2822 airfoil, and the Jones airfoil. A total of 23 sets of numerical results from 15 different author groups are included. The numerical methods used vary widely and include: 16 Navier-Stokes methods, 2 Euler/boundary-layer methods, and 5 full-potential/boundary-layer methods. The results indicate a high degree of sophistication among the numerical methods with generally good agreement between the various computed and experimental results for attached or moderately-separated cases. The agreement for cases with larger separation is only fair and suggests additional work is required in this area.

  17. Turbulent Potential Model Predictions of High Re Flow Around the S809 Airfoil

    NASA Astrophysics Data System (ADS)

    Develder, Nathaniel

    2015-11-01

    Utility scale wind turbines operate at a range of chord-based Reynolds numbers often between 106 and 107. Reynolds Averaged Navier-Stokes (RANS) models offer computational efficiency at high Reynolds numbers. As a model that avoids the eddy-viscosity hypothesis, the Turbulent Potential Model, a time-varying RANS model, is identified as an appropriate balance between computational resource usage and physical fidelity. Development of the Turbulent Potential Model is discussed. Comparisons are made between Turbulent Potential Model results and Moser's Direct Numerical Simulation Reτ =590 channel flow. S809 airfoil simulations at α = 0 .02° , α = 4 .03° , α = 10 .03° , and α = 20 .11° are compared to results from the k - ωSST , Spalart-Allmaras, and v2 - f models, as well as wind tunnel results from Ohio State University.

  18. Study of the far wake vortex field generated by a rectangular airfoil in a water tank

    NASA Technical Reports Server (NTRS)

    Lezius, D. K.

    1973-01-01

    Underwater towing experiments were carried out with a rectangular airfoil of aspect ratio 5.3 at 4 and 8 deg angles of attack and at chord-based Reynolds numbers between 2 x 100,000 and 7.5 x 100,000. Quantitative measurements by means of the hydrogen bubble technique indicated lower peak swirl velocities in the range of 100 to 1000 lenghts downstream than have been measured in wind tunnel of flight tests. The maximum circumferential velocity decayed whereas the turbulent eddy viscosity increased. This behavior and other known rates of vortex decay are explained in terms of an analytical solution for the vortex problem with time varying eddy viscosity. It is shown that this case corresponds to nonequilibrium turbulent vortex flow.

  19. Distinct neural responses to chord violations: a multiple source analysis study.

    PubMed

    Garza Villarreal, Eduardo A; Brattico, Elvira; Leino, Sakari; Ostergaard, Leif; Vuust, Peter

    2011-05-10

    The human brain is constantly predicting the auditory environment by representing sequential similarities and extracting temporal regularities. It has been proposed that simple auditory regularities are extracted at lower stations of the auditory cortex and more complex ones at other brain regions, such as the prefrontal cortex. Deviations from auditory regularities elicit a family of early negative electric potentials distributed over the frontal regions of the scalp. In this study, we wished to disentangle the brain processes associated with sequential vs. hierarchical auditory regularities in a musical context by studying the event-related potentials (ERPs), the behavioral responses to violations of these regularities, and the localization of the underlying ERP generators using two different source analysis algorithms. To this aim, participants listened to musical cadences constituted by seven chords, each containing either harmonically congruous chords, harmonically incongruous chords, or harmonically congruous but mistuned chords. EEG was recorded and multiple source analysis was performed. Incongruous chords violating the rules of harmony elicited a bilateral ERAN, whereas mistuned chords within chord sequences elicited a right-lateralized MMN. We found that the dominant cortical sources for the ERAN were localized around Broca's area and its right homolog, whereas the MMN generators were localized around the primary auditory cortex. These findings suggest a predominant role of the auditory cortices in detecting sequential scale regularities and the posterior prefrontal cortex in parsing hierarchical regularities in music. PMID:21382359

  20. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  2. The development of a facility for full-scale testing of airfoil performance in simulated rain

    NASA Technical Reports Server (NTRS)

    Taylor, John T.; Moore, Cadd T., III; Campbell, Bryan A.; Melson, W. EDWARD., Jr.

    1988-01-01

    NASA Langley's Aircraft Landing Dynamics Facility has been adapted in order to test the performance of airfoils in a simulated rain environment, at rainfall rates of 2, 10, 30, and 40 inches/hour, and thereby derive the scaling laws associated with simulated rain in wind tunnel testing. A full-scale prototype of the rain-generation system has been constructed and tested for suitable rain intensity, uniformity, effects of crosswinds on uniformity, and drop size range. The results of a wind tunnel test aimed at ascertaining the minimum length of the simulated rain field required to yield an airfoil performance change due to the rain environment are presented.

  3. 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.

  4. 14. Detail, end post/top chord connection point, west end of ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    14. Detail, end post/top chord connection point, west end of upstream truss, view to east, 210mm lens. The inclined end post is visible at right, with the top chord at left; the vertical member is a hanger. The latticed portal strut is partially visible at upper right, while paired diagonals approach the connection point from lower left, and a latticed top lateral member is visible above the top chord. - Southern Pacific Railroad Shasta Route, Bridge No. 310.58, Milepost 310.58, Sims, Shasta County, CA

  5. Application for calculation of mean aerodynamic chord of arbitrary wing planform

    NASA Astrophysics Data System (ADS)

    Vogeltanz, Tomáš

    2016-06-01

    This paper presents an application for the calculation of the mean aerodynamic chord (MAC) of an arbitrary wing planform. The MAC is most often used in the aerodynamic and stability analysis. The calculation uses a method where the MAC is defined by an array of chords. The MAC of each chord is computed as separated trapezoidal wing and then, the final MAC is calculated. This approach may find the accurate solution of complicated wing planforms, including elliptical, in the regime of low subsonic speeds. The first section describes the MAC and essential equations used for the computation. Finally, the application and two various examples of calculation are illustrated and discussed.

  6. Spline-Based Smoothing of Airfoil Curvatures

    NASA Technical Reports Server (NTRS)

    Li, W.; Krist, S.

    2008-01-01

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

  7. A multi-point inverse airfoil design method based on conformal mapping

    NASA Technical Reports Server (NTRS)

    Selig, Michael S.; Maughmer, Mark D.

    1991-01-01

    An exact method of multipoint inverse airfoil design for incompressible flow is presented. Multipoint design is handled by dividing the airfoil into a number of desired segments. For each segment the velocity distribution is prescribed together with an angle of attack at which the prescribed velocity distribution is to be achieved. In this manner, multipoint design objectives can be taken into account in the initial specification of the velocity distribution. In order for the multipoint inverse airfoil design problem to be well posed, three integral constraints and several conditions arise which must be satisfied. Further restrictions are imposed if the airfoil is to have a specified pitching moment, thickness ratio, or other constraints. The system of equations is solved partly as a linear system and partly through multidimensional Newton iteration. Since the velocity distribution is prescribed about the circle in the angular coordinate, specification of the velocity in terms of arc length is handled through the multidimensional Newton iteration as well. The current formulation sets the stage for a more general multipoint inverse airfoil design method in which it will be possible to specify the velocity distribution, some boundary-layer development, or the surface geometry along a segment.

  8. Active Control of Separation From the Flap of a Supercritical Airfoil

    NASA Technical Reports Server (NTRS)

    Melton, LaTunia Pack; Yao, Chung-Sheng; Seifert, Avi

    2006-01-01

    Zero-mass-flux periodic excitation was applied at several regions on a simplified high-lift system to delay the occurrence of flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approximately equal to 10) and low frequency amplitude modulation (F(+) sub AM approximately equal to 1) of the high frequency excitation were used for control. It was noted that the same performance gains were obtained with amplitude modulation and required only 30% of the momentum input required by pure sine excitation.

  9. [Event-related brain activity changes to consonant and dissonant chords in musicans and non-musicans].

    PubMed

    Maslennikova, A; Varlamov, A; Strelets, V

    2013-01-01

    The present study assessed brain activity changes related to perception of consonant and dissonant chords by musicians and non-musicians. Perception of dissonant chords in non-musicians was accompanied by increase of lower theta activity over right anterior regions, while consonant chords induced greater theta activity over left anterior regions; this pattern of asymmetrical activation was not observed in musicians. ERP analysis revealed that musicians had greater amplitude of early components (P100, N200) than non-musicians irrespective of chord type. The obtained results reflect more efficient musical harmony processing and, possibly, less emotional perception of chords in musicians. PMID:25438583

  10. 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.

  11. Unsteady Airloads on Airfoils in Reverse Flow

    NASA Astrophysics Data System (ADS)

    Lind, Andrew; Jones, Anya

    2014-11-01

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

  12. 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.

  13. Optimal Control of Airfoil Flow Separation using Fluidic Excitation

    NASA Astrophysics Data System (ADS)

    Shahrabi, Arireza F.

    as well as F+ were evaluated and discussed. The computational model predictions showed good agreement with the experimental data. It was observed that different angles of attack and flap angles have different requirements for the minimum value of the momentum coefficient, Cμ, in order for the SJA to be effective for control of separation. It was also found that the variation of F + noticeably affects the lift and drag forces acting on the airfoil. The optimum values of parameters during open loop control simulations have been applied in order to introduce the optimal open loop control outcome. An innovative approach has been implemented to formulate optimal frequencies and momentum ratios of vortex shedding which depends on angle of attack and static pressure of the separation zone in the upper chord. Optimal open loop results have been compared with the optimal closed loop results. Cumulative case studies in the matter of angle of attacks, flap angles, Re, Cμ and F+ provide a convincing collection of evidence to the following conclusion. An improvement of a direct closed loop control was demonstrated, and an analytical formula describing the properties of a separated flow and vortex shedding was proposed. Best AFC solutions are offered by providing optimal frequencies and momentum ratios at a variety of flow conditions.

  14. 9. DETAIL VIEW OF BOTTOM CHORD/FLOOR BEAM/IBAR PIN CONNECTION. WELDED ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    9. DETAIL VIEW OF BOTTOM CHORD/FLOOR BEAM/I-BAR PIN CONNECTION. WELDED PLATE AT PIN CONNECTION IS 20TH CENTURY REVISION. - Bucks County Bridge No. 313, Spanning Delaware Canal at Letchworth Avenue, Yardley, Bucks County, PA

  15. TOP CHORD, CENTRAL BRACING DETAIL. AveryBartholomew Patent Railroad Iron ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    TOP CHORD, CENTRAL BRACING DETAIL. - Avery-Bartholomew Patent Railroad Iron Bridge, Town park south of Route 222, west of Owasco Inlet (moved from Elm Street Extension spanning Fall Creek, Nubia, NY), Groton, Tompkins County, NY

  16. Airfoil Lift with Changing Angle of Attack

    NASA Technical Reports Server (NTRS)

    Reid, Elliott G

    1927-01-01

    Tests have been made in the atmospheric wind tunnel of the National Advisory Committee for Aeronautics to determine the effects of pitching oscillations upon the lift of an airfoil. It has been found that the lift of an airfoil, while pitching, is usually less than that which would exist at the same angle of attack in the stationary condition, although exceptions may occur when the lift is small or if the angle of attack is being rapidly reduced. It is also shown that the behavior of a pitching airfoil may be qualitatively explained on the basis of accepted aerodynamic theory.

  17. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  19. Experimental Test Results of Energy Efficient Transport (EET) High-Lift Airfoil in Langley Low-Turbulence Pressure Tunnel

    NASA Technical Reports Server (NTRS)

    Morgan, Harry L., Jr.

    2002-01-01

    This report describes the results of an experimental study conducted in the Langley Low-Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of the Langley Energy Efficient Transport (EET) High-Lift Airfoil. The high-lift airfoil was a supercritical-type airfoil with a thickness-to- chord ratio of 0.12 and was equipped with a leading-edge slat and a double-slotted trailing-edge flap. The leading-edge slat could be deflected -30 deg, -40 deg, -50 deg, and -60 deg, and the trailing-edge flaps could be deflected to 15 deg, 30 deg, 45 deg, and 60 deg. The gaps and overlaps for the slat and flaps were fixed at each deflection resulting in 16 different configurations. All 16 configurations were tested through a Reynolds number range of 2.5 to 18 million at a Mach number of 0.20. Selected configurations were also tested through a Mach number range of 0.10 to 0.35. The plotted and tabulated force, moment, and pressure data are available on the CD-ROM supplement L-18221.

  20. Laminar separation and turbulent reattachment for the Eppler 387 airfoil in its transitional Reynolds number range

    NASA Astrophysics Data System (ADS)

    McArthur, John; Spedding, Geoff

    2008-11-01

    Previous studies have shown that the Eppler 387 airfoil at chord-based Re = 60k experiences a large drag increase at moderate lift coefficients and angles of incidence, and a subsequent drag decrease at higher lift coefficients and angles of incidence. This drag increase has been blamed on the formation of a laminar separation bubble, where laminar separation occurs early on the wing surface, and a turbulent shear layer reattaches to the surface. However, evidence for this was only found at higher Re where this drag increase is negligible. Also, the flow field for this case has never been measured. The present study makes these flow field measurements using a custom PIV setup, and also makes force measurements using a custom force balance. The flow field measurements measure the recirculating flow in the bubble, and average profiles are constructed downstream of reattachment. Relating the flow fields to the forces explains how laminar separation and turbulent reattachment affect the forces produced by this airfoil.

  1. A numerical study of the effects of wind tunnel wall proximity on an airfoil model

    NASA Technical Reports Server (NTRS)

    Potsdam, Mark; Roberts, Leonard

    1990-01-01

    A procedure was developed for modeling wind tunnel flows using computational fluid dynamics. Using this method, a numerical study was undertaken to explore the effects of solid wind tunnel wall proximity and Reynolds number on a two-dimensional airfoil model at low speed. Wind tunnel walls are located at varying wind tunnel height to airfoil chord ratios and the results are compared with freestream flow in the absence of wind tunnel walls. Discrepancies between the constrained and unconstrained flows can be attributed to the presence of the walls. Results are for a Mach Number of 0.25 at angles of attack through stall. A typical wind tunnel Reynolds number of 1,200,000 and full-scale flight Reynolds number of 6,000,000 were investigated. At this low Mach number, wind tunnel wall corrections to Mach number and angle of attack are supported. Reynolds number effects are seen to be a consideration in wind tunnel testing and wall interference correction methods. An unstructured grid Navier-Stokes code is used with a Baldwin-Lomax turbulence model. The numerical method is described since unstructured flow solvers present several difficulties and fundamental differences from structured grid codes, especially in the area of turbulence modeling and grid generation.

  2. Large-eddy simulations of a turbulent Coanda jet on a circulation control airfoil

    NASA Astrophysics Data System (ADS)

    Nishino, Takafumi; Hahn, Seonghyeon; Shariff, Karim

    2010-12-01

    Large-eddy simulations are performed of a turbulent Coanda jet separating from a rounded trailing edge of a simplified circulation control airfoil model. The freestream Reynolds number based on the airfoil chord is 0.49×106, the jet Reynolds number based on the jet slot height is 4470, and the ratio of the peak jet velocity to the freestream velocity is 3.96. Three different grid resolutions are used to show that their effect is very small on the mean surface pressure distribution, which agrees very well with experiments, as well as on the mean velocity profiles over the Coanda surface. It is observed that the Coanda jet becomes fully turbulent just downstream of the jet exit, accompanied by asymmetric alternating vortex shedding behind a thin (but blunt) jet blade splitting the jet and the external flow. A number of "backward-tilted" hairpin vortices (i.e., the head of each hairpin being located upstream of the legs) are observed around the outer edge of the jet over the Coanda surface. These hairpins create strong upwash between the legs and weak downwash around them, contributing to turbulent mixing of the high-momentum jet below the hairpins and the low-momentum external flow above them. The probability density distribution of velocity fluctuations is shown to be highly asymmetric in this region, consistent with the observation that the hairpin vortices create strong upwash and weak downwash. Turbulent structures inside the jet, its spreading rate, and self-similarity are also discussed.

  3. Prediction of ice accretion on a swept NACA 0012 airfoil and comparisons to flight test results

    NASA Technical Reports Server (NTRS)

    Reehorst, Andrew L.

    1992-01-01

    In the winter of 1989-90, an icing research flight project was conducted to obtain swept wing ice accretion data. Utilizing the NASA Lewis Research Center's DHC-6 DeHavilland Twin Otter aircraft, research flights were made into known icing conditions in Northeastern Ohio. The icing cloud environment and aircraft flight data were measured and recorded by an onboard data acquisition system. Upon entry into the icing environment, a 24 inch span, 15 inch chord NACA 0012 airfoil was extended from the aircraft and set to the desired sweep angle. After the growth of a well defined ice shape, the airfoil was retracted into the aircraft cabin for ice shape documentation. The ice accretions were recorded by ice tracings and photographs. Ice accretions were mostly of the glaze type and exhibited scalloping. The ice was accreted at sweep angles of 0, 30, and 45 degrees. A 3-D ice accretion prediction code was used to predict ice profiles for five selected flight test runs, which include sweep angle of zero, 30, and 45 degrees. The code's roughness input parameter was adjusted for best agreement. A simple procedure was added to the code to account for 3-D ice scalloping effects. The predicted ice profiles are compared to their respective flight test counterparts. This is the first attempt to predict ice profiles on swept wings with significant scalloped ice formations.

  4. Preliminary Results of Cyclical De-Icing of a Gas-Heated Airfoil

    NASA Technical Reports Server (NTRS)

    Gray, V. H.; Bowden, D. T.; VonGlahn, U.

    1952-01-01

    An NACA 65(sub 1)-212 airfoil of 8-foot chord was provided with a gas-heated leading edge for investigations of cyclical de-icing. De-icing was accomplished with intermittent heating of airfoil segments that supplied hot gas to chordwise passages in a double-skin construction. Ice removal was facilitated by a spanwise leading-edge parting strip which was continuously heated from the gas-supply duct. Preliminary results demonstrate that satisfactory cyclical ice removal occurs with ratios of cycle time to heat-on period (cycle ratio) from 10 to 26. For minimum runback, efficient ice removal, and minimum total heat input, short heat-on periods of about 15 seconds with heat-off periods of 260 seconds gave the best results. In the range of conditions investigated, the prime variables in the determination of the required heat input for cyclical ice removal were the air temperature and the cycle ratio; heat-off period, liquid water content, airspeed, and angle of attack had only secondary effects on heat input rate.

  5. Boundary-Layer Transition on the N.A.C.A. 0012 and 23012 Airfoils in the 8-Foot High-Speed Wind Tunnel, Special Report

    NASA Technical Reports Server (NTRS)

    Becker, John V.

    1940-01-01

    Determinations of boundary-layer transition on the NACA 0012 and 2301 airfoils were made in the 8-foot high-speed wind tunnel over a range of Reynolds Numbers from 1,600,000 to 16,800,000. The results are of particular significance as compared with flight tests and tests in wind tunnels of appreciable turbulence because of the extremely low turbulence in the high-speed tunnel. A comparison of the results obtained on NACA 0012 airfoils of 2-foot and 5-foot chord at the same Reynolds Number permitted an evaluation of the effect of compressibility on transition. The local skin friction along the surface of the NACA 0012 airfoil was measured at a Reynolds Number of 10,000,000. For all the lift coefficient at which tests were made, transition occurred in the region of estimated laminar separation at the low Reynolds Numbers and approach the point of minimum static pressure as a forward limit at the high Reynolds Numbers. The effect of compressibility on transition was slight. None of the usual parameters describing the local conditions in the boundary layer near the transition point served as an index for locating the transition point. As a consequence of the lower turbulence in the 8-foot high-speed tunnel, the transition points occurred consistently farther back along the chord than those measured in the NACA full-scale tunnel. An empirical relation for estimating the location of the transition point for conventional airfoils on the basis of static-pressure distribution and Reynolds Number is presented.

  6. A comparison of two- and three-dimensional S809 airfoil properties for rough and smooth HAWT (horizontal-axis wind turbine) rotor operation

    SciTech Connect

    Musial, W.D.; Butterfield, C.P.; Jenks, M.D.

    1990-02-01

    At the Solar Energy Research Institute (SERI), we carried out tests to measure the effects of leading-edge roughness on an S809 airfoil using a 10-m, three-bladed, horizontal-axis wind turbine (HAWT). The rotor employed a constant-chord (.457 m) blade geometry with zero twist. Blade structural loads were measured with strain gages mounted at 9 spanwise locations. Airfoil pressure measurements were taken at the 80% spanwise station using 32 pressure taps distributed around the airfoil surface. Detailed inflow measurements were taken using nine R.M. Young Model 8002 propvane anemometers on a vertical plane array (VPA) located 10 m upwind of the test turbine in the prevailing wind direction. The major objective of this test was to determine the sensitivity of the S809 airfoil to roughness on a rotating wind turbine blade. We examined this effect by comparing several parameters. We compared power curves to show the sensitivity of whole rotor performance to roughness. We used pressure measurements to generate pressure distributions at the 80% span which operates at a Reynolds number (Re) of 800,000. We then integrated these distributions to determine the effect of roughness on the section's lift and pressure-drag coefficients. We also used the shapes of these distributions to understand how roughness affects the aerodynamic forces on the airfoil. We also compared rough and smooth wind tunnel data to the rotating blade data to study the effects of blade rotation on the aerodynamic behavior of the airfoil below, near, and beyond stall. 13 refs., 11 figs.

  7. A comparison of two-and three-dimensional S809 airfoil properties for rough and smooth HAWT (Horizontal-Axis Wind Tunnel) rotor operation

    NASA Astrophysics Data System (ADS)

    Musial, W. D.; Butterfield, C. P.; Jenks, M. D.

    1990-02-01

    At the Solar Energy Research Institute (SERI), we carried out tests to measure the effects of leading-edge roughness on an S809 airfoil using a 10-m, three-bladed, horizontal-axis wind turbine (HAWT). The rotor employed a constant-chord (.457 m) blade geometry with zero twist. Blade structural loads were measured with strain gages mounted at 9 spanwise locations. Airfoil pressure measurements were taken at the 80 percent spanwise station using 32 pressure taps distributed around the airfoil surface. Detailed inflow measurements were taken using nine R.M. Young Model 8002 propvane anemometers on a vertical plane array (VPA) located 10 m upwind of the test turbine in the prevailing wind direction. The major objective of this test was to determine the sensitivity of the S809 airfoil to roughness on a rotating wind turbine blade. We examined this effect by comparing several parameters. We compared power curves to show the sensitivity of whole rotor performance to roughness. We used pressure measurements to generate pressure distributions at the 80 percent span which operates at a Reynolds number (Re) of 800,000. We then integrated these distributions to determine the effect of roughness on the section's lift and pressure-drag coefficients.

  8. Pressure distributions from high Reynolds number tests of a Boeing BAC 1 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

    NASA Technical Reports Server (NTRS)

    Johnson, W. G., Jr.; Hill, A. S.

    1985-01-01

    A wind-tunnel investigation designed to test a Boeing advanced-technology airfoil from low to flight-equivalent Reynolds numbers has been completed in the Langley 0.3-Meter Transonic Cryogenic Tunnel. This investigation represents the first in a series of NASA/U.S. industry two-dimensional airfoil studies to be completed in the Advanced Technology Airfoil Test program. Test temperature was varied from ambient to about 100 K at pressures ranging from about 1.2 to 6.0 atm. Mach number was varied from about 0.40 to 0.80. These variables provided a Reynolds number (based on airfoil chord) range from 4.4 X 10 to the 6th power to 50.0 X 10 to the 6th power. All the test objectives were met. The pressure data are presented without analysis in plotted and tabulated formats for use in conjunction with the aerodynamic coefficient data published as NASA TM-81922. At the time of the test, these pressure data were considered proprietary and have only recently been made available by Boeing for general release. Data are included which demonstrate the effects of fixed transition. Also included are remarks on the model design, the model structural integrity, and the overall test experience.

  9. The calculation of flow over iced airfoils

    NASA Technical Reports Server (NTRS)

    Cebeci, Tuncer

    1988-01-01

    Progress toward the development of a method for predicting the flowfield of an iced airfoil is described and shown to offer the prospect of a priori calculations of the effects of ice accretion and roughness on airfoil performance. The approach is based on interaction of inviscid flow solutions obtained by a panel method and improved upon by a finite-difference boundary-layer method which, operating in an inverse mode, incorporates viscous effects including those associated with separated flows. Results are presented for smooth, rough and iced airfoils as a function of angle of attack. Those for smooth and rough airfoils confirm the accuracy of the method and its applicability to surfaces with roughness similar to that associated with insect deposition and some forms of ice. Two procedures have been developed to deal with large ice accretion and their performance is examined and shown to be appropriate to the engineering requirements.

  10. Low speed airfoil design and analysis

    NASA Technical Reports Server (NTRS)

    Eppler, R.; Somers, D. M.

    1979-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. Effects of grit roughness and pitch oscillations on the S801 airfoil

    SciTech Connect

    Ramsay, R.R.; Hoffman, M.J.; Gregorek, G.M.

    1996-01-01

    Horizontal axis wind turbine rotors experience unsteady aerodynamics due to wind shear when the rotor is yawed, when rotor blades pass through the support tower wake, and when the wind is gusting. An understanding of this unsteady behavior is necessary to assist in the calculation of rotor performance and loads. The rotors also experience performance degradation due to surface roughness. These surface irregularities are due to the accumulation of insect debris, ice, and the aging process. Wind tunnel studies that examine both the steady and unsteady behavior of airfoils can help define pertinent flow phenomena, and the resultant data can be used to validate analytical computer codes. A S801 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3x5 subsonic wind tunnel (3x5) under steady flow and stationary model conditions, as well as with the model undergoing pitch oscillations. To study the possible extent of performance loss due to surface roughness, a standard grit pattern (LEGR) was used to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers used for steady state conditions were 0.75, 1, 1.25, and 1.5 million, while the angle of attack ranged from -20{degrees} to +40{degrees}. With the model undergoing pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.4 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions were used, {plus_minus} 5.5 {degrees}and {plus_minus} 10{degrees}, at mean angles of attack of 8{degrees} 14{degrees} and 20{degrees} For purposes herein, any reference to unsteady conditions means that the airfoil model was in pitch oscillation about the quarter chord.

  15. A rapidly settled closed-loop control for airfoil aerodynamics based on plasma actuation

    NASA Astrophysics Data System (ADS)

    Wu, Z.; Wong, C. W.; Wang, L.; Lu, Z.; Zhu, Y.; Zhou, Y.

    2015-08-01

    This paper presents an experimental investigation on the response of the slope seeking with extended Kalman filter (EKF) deployed in a closed-loop system for airfoil aerodynamics control. A novel dielectric barrier discharge (DBD) plasma actuator was used to manipulate the flow around the NACA 0015 airfoil. Experiments were performed under different freestream velocities U ∞, covering the chord Reynolds number Re from 4.4 × 104 to 7.7 × 104. Firstly, the advantages of applying this DBD plasma actuator (hereafter called sawtooth plasma actuator) on the airfoil were examined in an open-loop system at Re = 7.7 × 104. The sawtooth plasma actuator led to a delay in the stall angle α stall by 5° and an increase in the maximum lift coefficient by about 9 %. On the other hand, at the same input power, the traditional DBD plasma actuator managed a delay in α stall by only 3° and an increase in by about 3 %. Secondly, the convergence time t c of the lift force F L at Re from 4.4 × 104 to 7.7 × 104 was investigated for two closed-loop systems. It has been demonstrated that the t c was about 70 % less under the slope seeking with EKF than that under the conventional slope seeking with high-pass (HP) and low-pass (LP) filters at Re = 7.7 × 104. The reduction in t c was also observed at a different Re. Finally, the slope seeking with EKF showed excellent robustness over a moderate Re range; that is, the voltage amplitude determined by the control algorithm promptly responded to a change in Re, much faster than that of the conventional slope seeking with HP and LP filters.

  16. Post stall airfoil data for wind turbines: wind tunnel test results

    SciTech Connect

    Ostowari, C.; Naik, D.

    1984-07-01

    Wind turbine blades operate over a wide angle of attack range. Unlike aircraft, a wind turbine's angle of attack range extends deep into stall where the three dimensional performance characteristics of airfoils are not generally known. Peak power predictions upon which wind turbine components are sized depend on a good understanding of a blade's post stall characteristics. The purpose of this wind tunnel study is to characterize the performance characteristics of a blade in stall as a function of its aspect ratio, airfoil thickness and Reynolds number. This report documents results of the wind tunnel investigation of constant chord blades having four aspect ratios, with NACA 44XX series airfoil sections, at angles of attack ranging from -10 to 110/sup 0/. Tests were conducted at Reynolds number ranging from one-quarter million to one million. The thickness ratios studied were 0.18, 0.15, 0.12 and 0.09. The aspect ratios were 6, 9, 12 and infinity. Results of force and pitching moment measurements, over the angle of attack range, for all combinations of Reynolds numbers, thickness and aspect ratios, and the effects of boundary layer tripping, have been presented. Both initial and secondary stall are presented. The maximum drag coefficient is found to occur at an angle of attack of 90/sup 0/. The pitching moment is unstable beyond stall. The lift and post-stall drag coefficients decrease with decreasing aspect ratio. The lift coefficient decreases with decreasing thickness ratio, while the drag coefficient increases. The boundary layer tripping is observed to decrease the lift curve slope and stalling angle of attack. The drag coefficient (with tripping) is significantly affected only at low aspect ratio.

  17. Airfoil for a gas turbine

    DOEpatents

    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.

  18. 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.

  19. Transonic airfoil design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

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

  20. TAIR: A transonic airfoil analysis computer code

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

  1. Wing chord extension on D-558-2

    NASA Technical Reports Server (NTRS)

    1953-01-01

    This 1953 NACA High-Speed Flight Research Station photograph shows the wing chord extensions added to D-558-2 #3. The added width on the outer wing panels was one attempt made to solve the tendence of the D-558-2 aircraft to 'pitch-up' during flight. Other modifications tested included fixed and free-floating slats, one- or two-wing fences, and a slat and wing fence. None of these additions proved truly effective, and the chord extension actually aggravated the pitch-up. The research program nevertheless added greatly to the store of knowledge about how to deal with pitch-up (as well as how not to deal with it, itself valuable information). The best solution was to place the horizontal stabilizer low on the aft fuselage, where it would be below the downwash from the wing and the wing wake. The Douglas D-558-2 'Skyrockets' were among the early transonic research airplanes like the X-1, X-4, X-5, and X-92A. Three of the single-seat, swept-wing aircraft flew from 1948 to 1956 in a joint program involving the National Advisory Committee for Aeronautics (NACA), with its flight research done at the NACA's Muroc Flight Test Unit in Calif., redesignated in 1949 the High-Speed Flight Research Station (HSFRS). Also partners in the flight research were the Navy-Marine Corps and the Douglas Aircraft Co. The HSFRS became the High-Speed Flight Station in 1954 and is now known as the NASA Dryden Flight Research Center. The Skyrocket made aviation history when it became the first airplane to fly twice the speed of sound. The 2 in the aircraft's designation referred to the fact that the Skyrocket was the phase-two version of what had originally been conceived as a three-phase program, with the phase-one aircraft having straight wings. The third phase, which never came to fruition, would have involved constructing a mock-up of a combat-type aircraft embodying the results from the testing of the phase one and two aircraft. Douglas pilot John F. Martin made the first flight at Muroc

  2. Propulsion by active and passive airfoil oscillation

    NASA Astrophysics Data System (ADS)

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

    2013-11-01

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

  3. Leading-edge singularities in thin-airfoil theory

    NASA Technical Reports Server (NTRS)

    Jones, R. T.

    1976-01-01

    If the thin airfoil theory is applied to an airfoil having a rounded leading edge, a certain error will arise in the determination of the pressure distribution around the nose. It is shown that the evaluation of the drag of such a blunt nosed airfoil by the thin airfoil theory requires the addition of a leading edge force, analogous to the leading edge thrust of the lifting airfoil. The method of calculation is illustrated by application to: (1) The Joukowski airfoil in subsonic flow; and (2) the thin elliptic cone in supersonic flow. A general formula for the edge force is provided which is applicable to a variety of wing forms.

  4. Experience Drives Synchronization: The phase and Amplitude Dynamics of Neural Oscillations to Musical Chords Are Differentially Modulated by Musical Expertise

    PubMed Central

    Pallesen, Karen Johanne; Bailey, Christopher J.; Brattico, Elvira; Gjedde, Albert; Palva, J. Matias; Palva, Satu

    2015-01-01

    Musical expertise is associated with structural and functional changes in the brain that underlie facilitated auditory perception. We investigated whether the phase locking (PL) and amplitude modulations (AM) of neuronal oscillations in response to musical chords are correlated with musical expertise and whether they reflect the prototypicality of chords in Western tonal music. To this aim, we recorded magnetoencephalography (MEG) while musicians and non-musicians were presented with common prototypical major and minor chords, and with uncommon, non-prototypical dissonant and mistuned chords, while watching a silenced movie. We then analyzed the PL and AM of ongoing oscillations in the theta (4–8 Hz) alpha (8–14 Hz), beta- (14–30 Hz) and gamma- (30–80 Hz) bands to these chords. We found that musical expertise was associated with strengthened PL of ongoing oscillations to chords over a wide frequency range during the first 300 ms from stimulus onset, as opposed to increased alpha-band AM to chords over temporal MEG channels. In musicians, the gamma-band PL was strongest to non-prototypical compared to other chords, while in non-musicians PL was strongest to minor chords. In both musicians and non-musicians the long-latency (> 200 ms) gamma-band PL was also sensitive to chord identity, and particularly to the amplitude modulations (beats) of the dissonant chord. These findings suggest that musical expertise modulates oscillation PL to musical chords and that the strength of these modulations is dependent on chord prototypicality. PMID:26291324

  5. Experience Drives Synchronization: The phase and Amplitude Dynamics of Neural Oscillations to Musical Chords Are Differentially Modulated by Musical Expertise.

    PubMed

    Pallesen, Karen Johanne; Bailey, Christopher J; Brattico, Elvira; Gjedde, Albert; Palva, J Matias; Palva, Satu

    2015-01-01

    Musical expertise is associated with structural and functional changes in the brain that underlie facilitated auditory perception. We investigated whether the phase locking (PL) and amplitude modulations (AM) of neuronal oscillations in response to musical chords are correlated with musical expertise and whether they reflect the prototypicality of chords in Western tonal music. To this aim, we recorded magnetoencephalography (MEG) while musicians and non-musicians were presented with common prototypical major and minor chords, and with uncommon, non-prototypical dissonant and mistuned chords, while watching a silenced movie. We then analyzed the PL and AM of ongoing oscillations in the theta (4-8 Hz) alpha (8-14 Hz), beta- (14-30 Hz) and gamma- (30-80 Hz) bands to these chords. We found that musical expertise was associated with strengthened PL of ongoing oscillations to chords over a wide frequency range during the first 300 ms from stimulus onset, as opposed to increased alpha-band AM to chords over temporal MEG channels. In musicians, the gamma-band PL was strongest to non-prototypical compared to other chords, while in non-musicians PL was strongest to minor chords. In both musicians and non-musicians the long-latency (> 200 ms) gamma-band PL was also sensitive to chord identity, and particularly to the amplitude modulations (beats) of the dissonant chord. These findings suggest that musical expertise modulates oscillation PL to musical chords and that the strength of these modulations is dependent on chord prototypicality. PMID:26291324

  6. Investigation of Effectiveness of a Wing Equipped with a 50-percent-chord Sliding Flap, a 30-percent-chord Slotted Flap, and a 30-percent-chord Slat in Deflecting Propeller Slipstreams Downward for Vertical Take-off

    NASA Technical Reports Server (NTRS)

    Kuhn, Richard E

    1957-01-01

    Results are presented of an investigation of the effectiveness of a wing equipped with a 50-percent-chord sliding flap and a 30-percent-chord slotted flap in deflecting a propeller slipstream downward for vertical take-off. Tests were conducted at zero forward speed in a large room and included the effects of flap deflection, proximity to the ground, a leading-edge slat, and end plates. A turning angle of about 70 degrees and a resultant force of about 100 percent of the thrust were achieved near the ground. Out of the ground-effect region, the turning angle was also about 70 degrees but the resultant force was reduced to about 86 percent of the thrust.

  7. 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.

  8. High-Lift, Low-Pitching-Moment Airfoils

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W.

    1987-01-01

    Two families of airfoil shapes improve rotor performance. Improvements enhance performances of helicopters and other rotorcraft but also applicable to aircraft propellers. Airfoil shapes best suited for inboard segment of rotor blade.

  9. Design of a subsonic airfoil with upstream blowing

    NASA Astrophysics Data System (ADS)

    Il'Inskii, N. B.; Mardanov, R. F.

    2007-10-01

    The problem is solved of designing a symmetric airfoil with upstream blowing opposite to subsonic irrotational steady flow of an inviscid incompressible fluid. The solution relies on Sedov’s idea of a stagnation region developing in the neighborhood of the stagnation point. An iterative solution process is developed, and examples of airfoils are constructed. The numerical results are analyzed, and conclusions are drawn about the effect of blowing parameters on the airfoil geometry and the resultant force acting on the airfoil.

  10. 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.

  11. Experimental Investigation of Water Droplet Impingement on Airfoils, Finite Wings, and an S-duct Engine Inlet

    NASA Technical Reports Server (NTRS)

    Papadakis, Michael; Hung, Kuohsing E.; Vu, Giao T.; Yeong, Hsiung Wei; Bidwell, Colin S.; Breer, Martin D.; Bencic, Timothy J.

    2002-01-01

    Validation of trajectory computer codes, for icing analysis, requires experimental water droplet impingement data for a wide range of aircraft geometries as well as flow and icing conditions. This report presents improved experimental and data reduction methods for obtaining water droplet impingement data and provides a comprehensive water droplet impingement database for a range of test geometries including an MS(1)-0317 airfoil, a GLC-305 airfoil, an NACA 65(sub 2)-415 airfoil, a commercial transport tail section, a 36-inch chord natural laminar flow NLF(1)-0414 airfoil, a 48-inch NLF(1)-0414 section with a 25 percent chord simple flap, a state-of-the-art three-element high lift system, a NACA 64A008 finite span swept business jet tail, a full-scale business jet horizontal tail section, a 25 percent-scale business jet empennage, and an S-duct turboprop engine inlet. The experimental results were obtained at the NASA Glenn Icing Research Tunnel (IRT) for spray clouds with median volumetric diameter (MVD) of 11, 11.5, 21, 92, and 94 microns and for a range of angles of attack. The majority of the impingement experiments were conducted at an air speed of 175 mph corresponding to a Reynolds number of approximately 1.6 million per foot. The maximum difference of repeated tests from the average ranged from 0.24 to 12 percent for most of the experimental results presented. This represents a significant improvement in test repeatability compared to previous experimental studies. The increase in test repeatability was attributed to improvements made to the experimental and data reduction methods. Computations performed with the LEWICE-2D and LEWICE-3D computer codes for all test configurations are presented in this report. For the test cases involving median volumetric diameters of 11 and 21 microns, the correlation between the analytical and experimental impingement efficiency distributions was good. For the median volumetric diameters of 92 and 94-micron cases, however

  12. Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane

    NASA Technical Reports Server (NTRS)

    Mason, Homer P.

    1953-01-01

    Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

  13. Understanding the Role of Chord-wise Flexibility in Flapping Wing Flight

    NASA Astrophysics Data System (ADS)

    Gaston, Zachary; Dong, Haibo; Wan, Hui; Ol, Michael

    2011-11-01

    Aerodynamic performance of flapping hinged plates is numerically studied to explore the effects of chord-wise flexibility in flapping wing flight. The plate with chord-wise flexibility is modeled as a two-link mechanism with a torsional spring hinge in between. The upper-link of the plate is controlled by prescribed motion and the rest of body is subjected to passive deflection due to fluid-body interaction. The effect of forced to natural frequency ratio is studied first for a flapping hinged-plate, on which prescribed hovering motion is actively applied. The effects of torsional stiffness and chord-wise flexibility are further explored for pitching and plunging plates, observing the flow phenomena and lift production as a result of this change. Comparisons between rigid plates, free-to-pivot hinged plates, and the torsional spring hinged plates are made, identifying a more optimal model for promoting lift production in flapping plates.

  14. New chording text entry methods combining physical and virtual buttons on a mobile phone.

    PubMed

    Wu, Fong-Gong; Huang, Yu-Chun; Wu, Meng-Long

    2014-07-01

    Traditional mobile phones depend on MultiTap, virtual or physical QWERTY keyboard for text entry, and they had some respective drawbacks include low input performance, occupying too large an area, high error rates, lack of feedbacks, etc. Therefore, some researches utilized the characteristics of the chording keyboard to improve input performance. Yet, as the learning efficiency of the chording keyboard is too low, users are not highly willing to learn. In view of that, this study combines the physical and virtual keys, and develops two chording input methods, MagArea and MemoryTap. After three days of learning, the fourteen experiment participants show effectively reduce error rates on MagArea, and they enhance their input speed on MemoryTap. In addition, excellent learning efficiency is found in the two methods, will be more motivated and willing to employ. PMID:24263072

  15. 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.

  16. S814 and S815 Airfoils: October 1991--July 1992

    SciTech Connect

    Somers, D. M.

    2004-12-01

    Two thick laminar-flow airfoils for the root portion of a horizontal-axis wind turbine blade, the S814 and S815, have been designed and analyzed theoretically. For both airfoils, the primary objectives of high maximum lift, insensitive to roughness, and low profile drag have been achieved. The constraints on pitching moment and airfoil thicknesses have been satisfied.

  17. The Effect of Conditional Probability of Chord Progression on Brain Response: An MEG Study

    PubMed Central

    Kim, Seung-Goo; Kim, June Sic; Chung, Chun Kee

    2011-01-01

    Background Recent electrophysiological and neuroimaging studies have explored how and where musical syntax in Western music is processed in the human brain. An inappropriate chord progression elicits an event-related potential (ERP) component called an early right anterior negativity (ERAN) or simply an early anterior negativity (EAN) in an early stage of processing the musical syntax. Though the possible underlying mechanism of the EAN is assumed to be probabilistic learning, the effect of the probability of chord progressions on the EAN response has not been previously explored explicitly. Methodology/Principal Findings In the present study, the empirical conditional probabilities in a Western music corpus were employed as an approximation of the frequencies in previous exposure of participants. Three types of chord progression were presented to musicians and non-musicians in order to examine the correlation between the probability of chord progression and the neuromagnetic response using magnetoencephalography (MEG). Chord progressions were found to elicit early responses in a negatively correlating fashion with the conditional probability. Observed EANm (as a magnetic counterpart of the EAN component) responses were consistent with the previously reported EAN responses in terms of latency and location. The effect of conditional probability interacted with the effect of musical training. In addition, the neural response also correlated with the behavioral measures in the non-musicians. Conclusions/Significance Our study is the first to reveal the correlation between the probability of chord progression and the corresponding neuromagnetic response. The current results suggest that the physiological response is a reflection of the probabilistic representations of the musical syntax. Moreover, the results indicate that the probabilistic representation is related to the musical training as well as the sensitivity of an individual. PMID:21364895

  18. Cloud-Hosted Real-time Data Services for the Geosciences (CHORDS)

    NASA Astrophysics Data System (ADS)

    Daniels, M. D.; Graves, S. J.; Kerkez, B.; Chandrasekar, V.; Vernon, F.; Martin, C. L.; Maskey, M.; Keiser, K.; Dye, M. J.

    2015-12-01

    The Cloud-Hosted Real-time Data Services for the Geosciences (CHORDS) project, funded as part of NSF's EarthCube initiative, addresses the ever-increasing importance of real-time scientific data, particularly in mission critical scenarios, where informed decisions must be made rapidly. Advances in the distribution of real-time data are leading many new transient phenomena in space-time to be observed, however, real-time decision-making is infeasible in many cases as these streaming data are either completely inaccessible or only available to proprietary in-house tools or displays. This lack of accessibility prohibits advanced algorithm and workflow development that could be initiated or enhanced by these data streams. Small research teams do not have resources to develop tools for the broad dissemination of their valuable real-time data and could benefit from an easy to use, scalable, cloud-based solution to facilitate access. CHORDS proposes to make a very diverse suite of real-time data available to the broader geosciences community in order to allow innovative new science in these areas to thrive. This presentation will highlight recently developed CHORDS portal tools and processing systems aimed at addressing some of the gaps in handling real-time data, particularly in the provisioning of data from the "long-tail" scientific community through a simple interface deployed in the cloud. The CHORDS system will connect these real-time streams via standard services from the Open Geospatial Consortium (OGC) and does so in a way that is simple and transparent to the data provider. Broad use of the CHORDS framework will expand the role of real-time data within the geosciences, and enhance the potential of streaming data sources to enable adaptive experimentation and real-time hypothesis testing. Adherence to community data and metadata standards will promote the integration of CHORDS real-time data with existing standards-compliant analysis, visualization and modeling

  19. A universal prediction of stall onset for airfoils at a wide range of Reynolds number flows

    NASA Astrophysics Data System (ADS)

    Morris, Wallace J., II

    The inception of leading-edge stall on two-dimensional, smooth, thin airfoils at various Reynolds number flows in the range O(103) to O(107) is investigated by an asymptotic approach and numerical simulations. The theory demonstrates that a subsonic flow about a thin airfoil can be described in terms of an outer region, around most of the airfoil 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 coordinates and a modified Reynolds number ReM, both based on the nose radius of curvature, are used to account for the nonlinear behavior and extreme velocity changes in the nose region, where stagnation and high suction occur. It results in a reduced-order model problem of a uniform, compressible, viscous flow past a semi-infinite canonic parabola. The inner far-field is governed by a circulation parameter A that is related to the airfoil's angle of attack, nose radius of curvature, thickness ratio, camber, and flow Mach number. The model parabola problem is solved numerically for various ReM and A using two methods. The first technique uses the steady Reynolds-Averaged Navier-Stokes (RANS) equations with the Spalart-Allmaras turbulence model for simulating moderate to high ReM flows. The second method applies direct numerical simulation (DNS) of the unsteady and incompressible Navier-Stokes equations for low to moderate ReM flows. In both methods, the critical value As is determined when a large separation zone first appears in the nose flow and the minimum pressure coefficient suddenly drops. The change of As with ReM is determined and these values indicate the onset of stall on the airfoil. The DNS results show that As decreases with ReM for ReM < ˜250, in agreement with Marginal Separation Theory (MST). However, calculations display the appearance of unsteady waves above a limiting value ReMcrit ˜250, where A s reaches a minimum of ˜1.55. For Re

  20. Dynamic simulation of chorded mitral valve in a left ventricle using an immersed boundary method

    NASA Astrophysics Data System (ADS)

    Luo, Xiaoyu; Yin, Min; Liang, Chunlei; Wang, Tiejun; Watton, Paul

    2007-11-01

    We use an immersed boundary model to investigate the dynamic behaviour of a chorded mitral prosthesis placed within a left ventricle under physiological flow conditions. In vivo magnetic resonance images of the left ventricle are used to create a numerical ventricle model. The motion of the ventricle model is prescribed during a cardiac cycle. Fluid-structure interaction simulations are carried out to test the performance of the mitral valve in a more realistic physiological environment. These simulations enable us to assess the effect of the ventricle motion, especially its flow vortex structure, on the function of the chorded mitral valve.