Science.gov

Sample records for airfoil trailing edge

  1. 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.

  2. 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.

  3. Trailing edge flow conditions as a factor in airfoil design

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.; Maughmer, M. D.

    1984-01-01

    Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the role played by the non-dimensional trailing edge velocity in design is considered and the relationship between the specified value and the resulting airfoil geometry is explored. In addition, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution of the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler is presented and the modified scheme employed to generate a number of airfoils for consideration. The detailed viscous analysis of airfoils having finite trailing edge pressure gradients demonstrates a reduction in the strong inviscid-viscid interactions generally present near the trailing edge of an airfoil.

  4. LES tests on airfoil trailing edge serration

    NASA Astrophysics Data System (ADS)

    Zhu, Wei Jun; Shen, Wen Zhong

    2016-09-01

    In the present study, a large number of acoustic simulations are carried out for a low noise airfoil with different Trailing Edge Serrations (TES). The Ffowcs Williams-Hawkings (FWH) acoustic analogy is used for noise prediction at trailing edge. The acoustic solver is running on the platform of our in-house incompressible flow solver EllipSys3D. The flow solution is first obtained from the Large Eddy Simulation (LES), the acoustic part is then carried out based on the instantaneous hydrodynamic pressure and velocity field. To obtain the time history data of sound pressure, the flow quantities are integrated around the airfoil surface through the FWH approach. For all the simulations, the chord based Reynolds number is around 1.5x106. In the test matrix, the effects from angle of attack, the TE flap angle, the length/width of the TES are investigated. Even though the airfoil under investigation is already optimized for low noise emission, most numerical simulations and wind tunnel experiments show that the noise level is further decreased by adding the TES device.

  5. Investigation of wave phenomena on a blunt airfoil with straight and serrated trailing edges

    NASA Astrophysics Data System (ADS)

    Nies, Juliane M.; Gageik, Manuel A.; Klioutchnikov, Igor; Olivier, Herbert

    2015-07-01

    An investigation of pressure waves in compressible subsonic and transonic flow around a generic airfoil is performed in a modified shock tube. New comprehensive results are presented on pressure waves in compressible flow. For the first time, the influence of trailing edge serration will be examined in terms of the reduction in pressure wave amplitude. A generic airfoil is tested in two main configurations, one with blunt trailing edges and the other one with serrated trailing edges in a Mach number range from 0.6 to 0.8 and at chord Reynolds numbers of 1 × 106 < Re c < 5 ×106. The flow of the blunt trailing edge is characterized by a regular vortex street in the wake creating a regular pattern of upstream-moving pressure waves along the airfoil. The observed pressure waves lead to strong pressure fluctuations within the local flow field. A reduction in the trailing edge thickness leads to a proportional increase in the frequency of the vortex street in the wake as well as the frequency of the waves deduced from constant Strouhal number. By serrating the trailing edge, the formation of vortices in the wake is disturbed. Therefore, also the upstream-moving waves are influenced and reduced in their strength resulting in a steadier flow. An increasing length of the saw tooth enhances the three dimensionality of the structures in the wake and causes a strong decrease in the wave amplitude.

  6. Experimental Study of Airfoil Trailing Edge Noise: Instrumentation, Methodology and Initial Results. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Manley, M. B.

    1980-01-01

    The mechanisms of aerodynamic noise generation at the trailing edge of an airfoil is investigated. Instrumentation was designed, a miniature semiconductor strain-gauge pressure transducer and associated electronic amplifier circuitry were designed and tested and digital signal analysis techniques applied to gain insight into the relationship between the dynamic pressure close to the trailing edge and the sound in the acoustic far-field. Attempts are made to verify some trailing-edge noise generation characteristics as theoretically predicted by several contemporary acousticians. It is found that the noise detected in the far-field is comprised of the sum of many uncorrelated emissions radiating from the vicinity of the trailing edge. These emissions appear to be the result of acoustic energy radiation which has been converted by the trailing-edge noise mechanism from the dynamic fluid energy of independent streamwise 'strips' of the turbulent boundary layer flow.

  7. An experimental study of airfoil instability tonal noise with trailing edge serrations

    NASA Astrophysics Data System (ADS)

    Chong, Tze Pei; Joseph, Phillip F.

    2013-11-01

    This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien-Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack. Larger Δf, which is defined as (fn+1-fn). In other words, a larger margin of velocity increase is required in order to "shift" the fn and fn+1 across fs

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

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1971-01-01

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

  9. 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.

  10. Optimization of the poro-serrated trailing edges for airfoil broadband noise reduction.

    PubMed

    Chong, Tze Pei; Dubois, Elisa

    2016-08-01

    This paper reports an aeroacoustic investigation of a NACA0012 airfoil with a number of poro-serrated trailing edge devices that contain porous materials of various air flow resistances at the gaps between adjacent members of the serrated-sawtooth trailing edge. The main objective of this work is to determine whether multiple-mechanisms on the broadband noise reduction can co-exist on a poro-serrated trailing edge. When the sawtooth gaps are filled with porous material of low-flow resistivity, the vortex shedding tone at low-frequency could not be completely suppressed at high-velocity, but a reasonably good broadband noise reduction can be achieved at high-frequency. When the sawtooth gaps are filled with porous material of very high-flow resistivity, no vortex shedding tone is present, but the serration effect on the broadband noise reduction becomes less effective. An optimal choice of the flow resistivity for a poro-serrated configuration has been identified, where it can surpass the conventional serrated trailing edge of the same geometry by achieving a further 1.5 dB reduction in the broadband noise while completely suppressing the vortex shedding tone. A weakened turbulent boundary layer noise scattering at the poro-serrated trailing edge is reflected by the lower-turbulence intensity at the near wake centreline across the whole spanwise wavelength of the sawtooth.

  11. Optimization of the poro-serrated trailing edges for airfoil broadband noise reduction.

    PubMed

    Chong, Tze Pei; Dubois, Elisa

    2016-08-01

    This paper reports an aeroacoustic investigation of a NACA0012 airfoil with a number of poro-serrated trailing edge devices that contain porous materials of various air flow resistances at the gaps between adjacent members of the serrated-sawtooth trailing edge. The main objective of this work is to determine whether multiple-mechanisms on the broadband noise reduction can co-exist on a poro-serrated trailing edge. When the sawtooth gaps are filled with porous material of low-flow resistivity, the vortex shedding tone at low-frequency could not be completely suppressed at high-velocity, but a reasonably good broadband noise reduction can be achieved at high-frequency. When the sawtooth gaps are filled with porous material of very high-flow resistivity, no vortex shedding tone is present, but the serration effect on the broadband noise reduction becomes less effective. An optimal choice of the flow resistivity for a poro-serrated configuration has been identified, where it can surpass the conventional serrated trailing edge of the same geometry by achieving a further 1.5 dB reduction in the broadband noise while completely suppressing the vortex shedding tone. A weakened turbulent boundary layer noise scattering at the poro-serrated trailing edge is reflected by the lower-turbulence intensity at the near wake centreline across the whole spanwise wavelength of the sawtooth. PMID:27586762

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

    NASA Astrophysics Data System (ADS)

    Jaksich, Dylan; Shen, Jinwei

    2014-11-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  14. LES of High-Reynolds-Number Coanda Flow Separating from a Rounded Trailing Edge of a Circulation Control Airfoil

    NASA Technical Reports Server (NTRS)

    Nichino, Takafumi; Hahn, Seonghyeon; Shariff, Karim

    2010-01-01

    This slide presentation reviews the Large Eddy Simulation of a high reynolds number Coanda flow that is separated from a round trailing edge of a ciruclation control airfoil. The objectives of the study are: (1) To investigate detailed physics (flow structures and statistics) of the fully turbulent Coanda jet applied to a CC airfoil, by using LES (2) To compare LES and RANS results to figure out how to improve the performance of existing RANS models for this type of flow.

  15. Effect of airfoil (trailing-edge) thickness on the numerical solution of panel methods based on the Dirichlet boundary condition

    NASA Technical Reports Server (NTRS)

    Yon, Steven; Katz, Joseph; Plotkin, Allen

    1992-01-01

    The practical limit of airfoil thickness ratio for which acceptable engineering results are obtainable with the Dirichlet boundary-condition-based numerical methods is investigated. This is done by studying the effect of thickness on the calculated pressure distribution near the trailing edge and by comparing the aerodynamic coefficients with available exact solutions. The first objective of this study, owing to the wide use of such computational methods, is to demonstrate the numerical symptoms that occur when the body or wing thickness approaches zero and to increase the awareness of potential users of these methods. Additionally, an effort is made to obtain the practical limits of the trailing-edge thickness where such problems will appear in the flow solution, and to propose some possible cures for very thin airfoils or those with cusped trailing edges.

  16. Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections

    NASA Technical Reports Server (NTRS)

    Brooks, T. F.; Marcolini, M. A.; Pope, D. S.

    1984-01-01

    Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.

  17. 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.

  18. An experimental study of a turbulent boundary layer in the trailing edge region of a circulation-control airfoil

    NASA Technical Reports Server (NTRS)

    Heinemann, K.; Brown, Jeff

    1992-01-01

    This report discusses progress made on NASA Cooperative Agreement NCC2-545, 'An Experimental Study of a Turbulent Boundary Layer in the Trailing-Edge Region of a Circulation-Control Airfoil' during the period 9/1/91 through 9/30/92. The study features 2-component laser Doppler velocimeter (LDV) measurements in the trailing edge and wake regions of a generic 2-dimensional circulation-control model. The final experimental phase of the study will be carried out in the Ames High Reynolds Number Channel 2 (HRC2) transonic blow-down-facility. During the 13-month period covered by this report, work continued on the development of the near-wall laser Doppler velocimeter (LDV) described in previous reports.

  19. Circulation control on a rounded trailing-edge wind turbine airfoil using plasma actuators

    NASA Astrophysics Data System (ADS)

    Baleriola, S.; Leroy, A.; Loyer, S.; Devinant, P.; Aubrun, S.

    2016-09-01

    This experimental study focuses on the implementation via plasma actuators of a circulation control strategy on a wind turbine aerofoil with a rounded trailing-edge with the objective of reducing the aerodynamic load fluctuations on blades. Three sets of multi-DBD (Dielectric Barrier Discharge) actuators with different positions around the trailing-edge are studied. These actuators create a tangential jet that adheres to the blade model wall and diffuses along it. According to the jet direction, lift is increased or decreased. Load and pressure measurements as well as Particle Image Velocimetry (PIV) show respectively the actuation effectiveness in terms of load modification and flow topology alteration.

  20. An experimental study of a turbulent boundary layer in the trailing edge region of a curculation-control airfoil

    NASA Technical Reports Server (NTRS)

    Brown, Jeff

    1993-01-01

    This report discusses progress made on NASA Cooperative Agreement NCC2-545, 'An Experimental Study of a Turbulent Boundary Layer in the Trailing-Edge Region of a Circulation-Control Airfoil,' during the period 1 Oct. 1992 - 30 Jun. 1993. The study, being conducted by Jeff Brown of the Eloret Institute, in conjunction with the Experimental Fluid Dynamics Branch at NASA Ames (Dennis Johnson, technical monitor), features 2-component laser Doppler velocimeter (LDV) measurements in the trailing edge and wake regions of a generic circulation-control airfoil model. The final experimental phase of the study will be carried out in the Ames High Reynolds Number Channel II (HRC2) transonic blow-down facility. During the 9-month period covered by this report, important data were acquired using the near-wall laser Doppler velocimeter (LDV) whose development has been described in previous reports. These data point strongly to the viability of this new technique for measuring the full Reynolds Stress Tensor in 3D flows.

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

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  2. Investigation of the near and far wake of a bluff airfoil model with trailing edge modifications using time-resolved particle image velocimetry

    NASA Astrophysics Data System (ADS)

    Krentel, Daniel; Nitsche, Wolfgang

    2013-07-01

    Experimental investigations on the topology and the structure of the near and far wake of a quasi-2D blunt NACA0012 airfoil cut at 80 % of the original chord length c master have been performed by means of time-resolved particle image velocimetry. The experiments took place in a closed-loop water tunnel at a model-thickness H-based Reynolds number of Re H = 44,000. The periodic and alternating vortex formation process at the base of the bluff model with a dimensionless frequency of Sr h = 0.2 (relating to the trailing edge height h) was investigated in detail. Subsequently, four modifications of the trailing edge geometry (broken trailing edge, square-wave base, stepped afterbody and extension of the reference model by Δ c/c_master = 7.5 %) have been investigated in order to mitigate the periodic vortex formation and the alternating shedding process. In the far wake, a considerable decrease in momentum loss and resulting drag force in the range of 29 % has been achieved for this specific Reynolds number. Investigations of the time-resolved flow field proved that the periodic, alternating flow separation can be attenuated resulting in an optimized recirculation region and a low-loss wake. It can be inferred that passive flow control means like modifications of the rear end geometry of quasi-2D blunt models are a capable method to improve the flow field with respect to a minimization of momentum losses in the wake.

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

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

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

  4. Aircraft wing trailing-edge noise

    NASA Technical Reports Server (NTRS)

    Underwood, R. L.; Hodgson, T. H.

    1981-01-01

    The mechanism and sound pressure level of the trailing-edge noise for two-dimensional turbulent boundary layer flow was examined. Experiment is compared with current theory. A NACA 0012 airfoil of 0.61 m chord and 0.46 m span was immersed in the laminar flow of a low turbulence open jet. A 2.54 cm width roughness strip was placed at 15 percent chord from the leading edge on both sides of the airfoil as a boundary layer trip so that two separate but statistically equivalent turbulent boundary layers were formed. Tests were performed with several trailing-edge geometries with the upstream velocity U sub infinity ranging from a value of 30.9 m/s up to 73.4 m/s. Properties of the boundary layer for the airfoil and pressure fluctuations in the vicinity of the trailing-edge were examined. A scattered pressure field due to the presence of the trailing-edge was observed and is suggested as a possible sound producing mechanism for the trailing-edge noise.

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

    NASA Technical Reports Server (NTRS)

    Allen, H Julian

    1938-01-01

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

  6. Airfoil

    SciTech Connect

    Ristau, Neil; Siden, Gunnar Leif

    2015-07-21

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

  7. Trailing Edge Blowing on a Two-Dimensional Six-Percent Thick Elliptical Circulation Control Airfoil Up to Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Alexander, Michael G.; Anders, Scott G.; Johnson, Stuart K.; Florance, Jennifer P.; Keller, Donald F.

    2005-01-01

    A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

  8. Leading-Edge "Pop-Up" Spoiler For Airfoil

    NASA Technical Reports Server (NTRS)

    Wilson, John C.; Lance, Michael B.

    1991-01-01

    New concept places spoiler in leading edge of airfoil, hinged along its trailing edge, so airflow helps to deploy it and force it against mechanical stop. Deployed "pop-up" spoiler quickly eliminates almost all aerodynamic lift of stabilator. Designed to be added to leading edge of existing stabilator, without major rework. Though initial application to be on helicopter stabilators, equally applicable to wings or winglike components.

  9. Trailing Edge Noise Prediction Based on a New Acoustic Formulation

    NASA Technical Reports Server (NTRS)

    Casper, J.; Farassat, F.

    2002-01-01

    A new analytic result in acoustics called 'Formulation 1B,' proposed by Farassat, is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Williams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experiment. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, using both analytical and experimental data on the airfoil surface. The results are compared to analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

  10. Broadband Trailing Edge Noise Predictions in the Time Domain. Revised

    NASA Technical Reports Server (NTRS)

    Casper, Jay; Farassat, Fereidoun

    2003-01-01

    A recently developed analytic result in acoustics, "Formulation 1B," is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Willliams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experimental results. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, by using both analytical and experimental data on the airfoil surface. The acoustic predictions are compared with analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

  11. Shape Optimization for Trailing Edge Noise Control

    NASA Astrophysics Data System (ADS)

    Marsden, Alison; Wang, Meng; Mohammadi, Bijan; Moin, Parviz

    2001-11-01

    Noise generated by turbulent boundary layers near the trailing edge of lifting surfaces continues to pose a challenge for many applications. In this study, we explore noise reduction strategies through shape optimization. A gradient based shape design method is formulated and implemented for use with large eddy simulation of the flow over an airfoil. The cost function gradient is calculated using the method of incomplete sensitivities (Mohammadi and Pironneau 2001 ph Applied shape Optimization for Fluids, Oxford Univ. Press). This method has the advantage that effects of geometry changes on the flow field can be neglected when computing the gradient of the cost function, making it far more cost effective than solving the full adjoint problem. Validation studies are presented for a model problem of the unsteady laminar flow over an acoustically compact airfoil. A section of the surface is allowed to deform and the cost function is derived based on aeroacoustic theroy. Rapid convergence of the trailing-edge shape and significant reduction of the noise due to vortex shedding and wake instability have been achieved. The addition of constraints and issues of extension to fully turbulent flows past an acoustically noncompact airfoil are also discussed.

  12. Simulation of Acoustic Scattering from a Trailing Edge

    NASA Technical Reports Server (NTRS)

    Singer, Bart A.; Brentner, Kenneth S.; Lockhard, David P.; Lilley, Geoffrey M.

    1999-01-01

    Three model problems were examined to assess the difficulties involved in using a hybrid scheme coupling flow computation with the the Ffowcs Williams and Hawkings equation to predict noise generated by vortices passing over a sharp edge. The results indicate that the Ffowcs Williams and Hawkings equation correctly propagates the acoustic signals when provided with accurate flow information on the integration surface. The most difficult of the model problems investigated inviscid flow over a two-dimensional thin NACA airfoil with a blunt-body vortex generator positioned at 98 percent chord. Vortices rolled up downstream of the blunt body. The shed vortices possessed similarities to large coherent eddies in boundary layers. They interacted and occasionally paired as they convected past the sharp trailing edge of the airfoil. The calculations showed acoustic waves emanating from the airfoil trailing edge. Acoustic directivity and Mach number scaling are shown.

  13. Large eddy simulation of trailing edge noise

    NASA Astrophysics Data System (ADS)

    Keller, Jacob; Nitzkorski, Zane; Mahesh, Krishnan

    2015-11-01

    Noise generation is an important engineering constraint to many marine vehicles. A significant portion of the noise comes from propellers and rotors, specifically due to flow interactions at the trailing edge. Large eddy simulation is used to investigate the noise produced by a turbulent 45 degree beveled trailing edge and a NACA 0012 airfoil. A porous surface Ffowcs-Williams and Hawkings acoustic analogy is combined with a dynamic endcapping method to compute the sound. This methodology allows for the impact of incident flow noise versus the total noise to be assessed. LES results for the 45 degree beveled trailing edge are compared to experiment at M = 0 . 1 and Rec = 1 . 9 e 6 . The effect of boundary layer thickness on sound production is investigated by computing using both the experimental boundary layer thickness and a thinner boundary layer. Direct numerical simulation results of the NACA 0012 are compared to available data at M = 0 . 4 and Rec = 5 . 0 e 4 for both the hydrodynamic field and the acoustic field. Sound intensities and directivities are investigated and compared. Finally, some of the physical mechanisms of far-field noise generation, common to the two configurations, are discussed. Supported by Office of Naval research.

  14. Effect of Trailing Edge Shape on the Unsteady Aerodynamics of Reverse Flow Dynamic Stall

    NASA Astrophysics Data System (ADS)

    Lind, Andrew; Jones, Anya

    2015-11-01

    This work considers dynamic stall in reverse flow, where flow travels over an oscillating airfoil from the geometric trailing edge towards the leading edge. An airfoil with a sharp geometric trailing edge causes early formation of a primary dynamic stall vortex since the sharp edge acts as the aerodynamic leading edge in reverse flow. The present work experimentally examines the potential merits of using an airfoil with a blunt geometric trailing edge to delay flow separation and dynamic stall vortex formation while undergoing oscillations in reverse flow. Time-resolved and phase-averaged flow fields and pressure distributions are compared for airfoils with different trailing edge shapes. Specifically, the evolution of unsteady flow features such as primary, secondary, and trailing edge vortices is examined. The influence of these flow features on the unsteady pressure distributions and integrated unsteady airloads provide insight on the torsional loading of rotor blades as they oscillate in reverse flow. The airfoil with a blunt trailing edge delays reverse flow dynamic stall, but this leads to greater downward-acting lift and pitching moment. These results are fundamental to alleviating vibrations of high-speed helicopters, where much of the rotor operates in reverse flow.

  15. The effect of multiple fixed slots and a trailing-edge flap on the lift and drag of a Clark Y airfoil

    NASA Technical Reports Server (NTRS)

    Weick, Fred E; Shortal, Joseph A

    1933-01-01

    Lift and drag tests were made on a Clark Y wing equipped with four fixed slots and a trailing-edge flap in the 5-foot vertical wind tunnel of the National Advisory Committee for Aeronautics. All possible combinations of the four slots were tested with the flap neutral and the most promising combinations were tested with the flap down 45 degrees. Considering both the maximum lift coefficient and the speed-range ratio with the flap neutral no appreciable improvement was found with the use of more than the single leading-edge slot. With the flap down 45 degrees a maximum lift coefficient of 2.60 was obtained but the particular slot combination used had a rather large minimum drag coefficient with the flap neutral. With the flap down 45 degrees the optimum combination, considering both the maximum lift coefficient and the speed-range ratio, was obtained with only the two rearmost slots in use. For this arrangement the maximum lift coefficient was 2.44.

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

    NASA Technical Reports Server (NTRS)

    Mclachlan, B. G.

    1987-01-01

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

  17. Static Extended Trailing Edge for Lift Enhancement: Experimental and Computational Studies

    NASA Technical Reports Server (NTRS)

    Liu, Tianshu; Montefort; Liou, William W.; Pantula, Srinivasa R.; Shams, Qamar A.

    2007-01-01

    A static extended trailing edge attached to a NACA0012 airfoil section is studied for achieving lift enhancement at a small drag penalty. It is indicated that the thin extended trailing edge can enhance the lift while the zero-lift drag is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge with those of Gurney flap and conventional flap. The extended trailing edge, as a simple mechanical device added on a wing without altering the basic configuration, has a good potential to improve the cruise flight efficiency.

  18. Survey of techniques for reduction of wind turbine blade trailing edge noise.

    SciTech Connect

    Barone, Matthew Franklin

    2011-08-01

    Aerodynamic noise from wind turbine rotors leads to constraints in both rotor design and turbine siting. The primary source of aerodynamic noise on wind turbine rotors is the interaction of turbulent boundary layers on the blades with the blade trailing edges. This report surveys concepts that have been proposed for trailing edge noise reduction, with emphasis on concepts that have been tested at either sub-scale or full-scale. These concepts include trailing edge serrations, low-noise airfoil designs, trailing edge brushes, and porous trailing edges. The demonstrated noise reductions of these concepts are cited, along with their impacts on aerodynamic performance. An assessment is made of future research opportunities in trailing edge noise reduction for wind turbine rotors.

  19. On the generation of side-edge flap noise. [part span trailing edge flaps

    NASA Technical Reports Server (NTRS)

    Howe, M. S.

    1981-01-01

    A theory is proposed for estimating the noise generated at the side edges of part span trailing edge flaps in terms of pressure fluctuations measured just in-board of the side edge of the upper surface of the flap. Asymptotic formulae are developed in the opposite extremes of Lorentz contracted acoustic wavelength large/small compared with the chord of the flap. Interpolation between these limiting results enables the field shape and its dependence on subsonic forward flight speed to be predicted over the whole frequency range. It is shown that the mean width of the side edge gap between the flap and the undeflected portion of the airfoil has a significant influence on the intensity of the radiated sound. It is estimated that the noise generated at a single side edge of a full scale part span flap can exceed that produced along the whole of the trailing edge of the flap by 3 dB or more.

  20. Computational investigation of miniature trailing edge effectors

    NASA Astrophysics Data System (ADS)

    Lee, Hak-Tae

    Miniature trailing edge effectors (MiTEs) are small flaps (typically 1% to 5% chord) actuated with deflection angles of up to 90 degrees. The small size, combined with little required power and good control authority, enables the device to be used for high bandwidth control as well as conventional attitude control. However, some of the aerodynamic characteristics of these devices are complex and poorly understood. This research investigated the aerodynamics of MiTEs using incompressible Navier-Stokes flow solvers, INS2D and INS3D. To understand the flow structure and establish a parametric database, two dimensional steady-state computations were performed for MiTEs with various geometries and flow conditions. Time accurate computations were used to resolve the unsteady characteristics including transient response and vortex shedding phenomena. The frequency response was studied to fully identify the dynamics of MiTEs. Three dimensional computations show the change in control effectiveness with respect to the spanwise length of MiTEs as well as the spanwise lift distribution induced by these devices. Based on the CFD results, an approximate vortex panel model was developed for design purposes that reproduces the key characteristics of MiTEs. Two application areas for MiTEs were explored. Flutter suppression was demonstrated by combining a finite element structural model with the vortex panel model. The application of MiTEs to augment maximum lift and improve the post stall behavior of an airfoil was also investigated.

  1. Asymptotic theory of two-dimensional trailing-edge flows

    NASA Technical Reports Server (NTRS)

    Melnik, R. E.; Chow, R.

    1975-01-01

    Problems of laminar and turbulent viscous interaction near trailing edges of streamlined bodies are considered. Asymptotic expansions of the Navier-Stokes equations in the limit of large Reynolds numbers are used to describe the local solution near the trailing edge of cusped or nearly cusped airfoils at small angles of attack in compressible flow. A complicated inverse iterative procedure, involving finite-difference solutions of the triple-deck equations coupled with asymptotic solutions of the boundary values, is used to accurately solve the viscous interaction problem. Results are given for the correction to the boundary-layer solution for drag of a finite flat plate at zero angle of attack and for the viscous correction to the lift of an airfoil at incidence. A rational asymptotic theory is developed for treating turbulent interactions near trailing edges and is shown to lead to a multilayer structure of turbulent boundary layers. The flow over most of the boundary layer is described by a Lighthill model of inviscid rotational flow. The main features of the model are discussed and a sample solution for the skin friction is obtained and compared with the data of Schubauer and Klebanoff for a turbulent flow in a moderately large adverse pressure gradient.

  2. Aerodynamic Characteristic of the Active Compliant Trailing Edge Concept

    NASA Astrophysics Data System (ADS)

    Nie, Rui; Qiu, Jinhao; Ji, Hongli; Li, Dawei

    2016-06-01

    This paper introduces a novel Morphing Wing structure known as the Active Compliant Trailing Edge (ACTE). ACTE structures are designed using the concept of “distributed compliance” and wing skins of ACTE are fabricated from high-strength fiberglass composites laminates. Through the relative sliding between upper and lower wing skins which are connected by a linear guide pairs, the wing is able to achieve a large continuous deformation. In order to present an investigation about aerodynamics and noise characteristics of ACTE, a series of 2D airfoil analyses are established. The aerodynamic characteristics between ACTE and conventional deflection airfoil are analyzed and compared, and the impacts of different ACTE structure design parameters on aerodynamic characteristics are discussed. The airfoils mentioned above include two types (NACA0012 and NACA64A005.92). The computing results demonstrate that: compared with the conventional plane flap airfoil, the morphing wing using ACTE structures has the capability to improve aerodynamic characteristic and flow separation characteristic. In order to study the noise level of ACTE, flow field analysis using LES model is done to provide noise source data, and then the FW-H method is used to get the far field noise levels. The simulation results show that: compared with the conventional flap/aileron airfoil, the ACTE configuration is better to suppress the flow separation and lower the overall sound pressure level.

  3. Miniature Trailing Edge Effector for Aerodynamic Control

    NASA Technical Reports Server (NTRS)

    Lee, Hak-Tae (Inventor); Bieniawski, Stefan R. (Inventor); Kroo, Ilan M. (Inventor)

    2008-01-01

    Improved miniature trailing edge effectors for aerodynamic control are provided. Three types of devices having aerodynamic housings integrated to the trailing edge of an aerodynamic shape are presented, which vary in details of how the control surface can move. A bucket type device has a control surface which is the back part of a C-shaped member having two arms connected by the back section. The C-shaped section is attached to a housing at the ends of the arms, and is rotatable about an axis parallel to the wing trailing edge to provide up, down and neutral states. A flip-up type device has a control surface which rotates about an axis parallel to the wing trailing edge to provide up, down, neutral and brake states. A rotating type device has a control surface which rotates about an axis parallel to the chord line to provide up, down and neutral states.

  4. Laminar-flow airfoil

    NASA Technical Reports Server (NTRS)

    Somers, Dan M. (Inventor)

    2005-01-01

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

  5. Near Field Trailing Edge Tone Noise Computation

    NASA Technical Reports Server (NTRS)

    Loh, Ching Y.

    2002-01-01

    Blunt trailing edges in a flow often generate tone noise due to wall-jet shear layer and vortex shedding. In this paper, the space-time conservation element (CE/SE) method is employed to numerically study the near-field noise of blunt trailing edges. Two typical cases, namely, flow past a circular cylinder (aeolian noise problem) and flow past a flat plate of finite thickness are considered. The computed frequencies compare well with experimental data. For the aeolian noise problem, comparisons with the results of other numerical approaches are also presented.

  6. Trailing edge noise prediction from measured surface pressures

    NASA Technical Reports Server (NTRS)

    Brooks, T. F.; Hodgson, T. H.

    1981-01-01

    Trailing edge (TE) noise is investigated for the case of a two-dimensional airfoil embedded in a uniform low Mach number flow, and the usefulness of several TE noise theories is examined by applying them to the measured data. The TE noise spectra and directivity are quantitatively determined for the case of a high Reynolds number and a fully turbulent boundary layer. Parameters include angle of attack, flow velocity and TE bluntness. Evanescent wave theories by Chase (1975) and Chandiramani (1974) are compared to the present results and show good agreement. Agreement of the near field pressure scatter phenomenon analysis with measurements implies that the basic assumptions used in the analysis are correct, i.e., the turbulent boundary layer (TBL) flow passes the trailing edge into the wake region. No hydrodynamic wake shedding activity is confirmed for the two-sided TBL flow, and a method incorporating the principles of the coherent output power method is used to determine the sound field. The near field edge scatter model is found to not only establish optimum sizing of edge treatment for noise control, but also to separate and identify the scattered field from the incident hydrodynamic field.

  7. Helicopter rotor trailing edge noise. [noise prediction

    NASA Technical Reports Server (NTRS)

    Schlinker, R. H.; Amier, R. K.

    1981-01-01

    A two dimensional section of a helicopter main rotor blade was tested in an acoustic wind tunnel at close to full-scale Reynolds numbers to obtain boundary layer data and acoustic data for use in developing an acoustic scaling law and testing a first principles trailing edge noise theory. Results were extended to the rotating frame coordinate system to develop a helicopter rotor trailing edge noise prediction. Comparisons of the calculated noise levels with helicopter flyover spectra demonstrate that trailing edge noise contributes significantly to the total helicopter noise spectrum at high frequencies. This noise mechanism is expected to control the minimum rotor noise. In the case of noise radiation from a local blade segment, the acoustic directivity pattern is predicted by the first principles trailing edge noise theory. Acoustic spectra are predicted by a scaling law which includes Mach number, boundary layer thickness and observer position. Spectrum shape and sound pressure level are also predicted by the first principles theory but the analysis does not predict the Strouhal value identifying the spectrum peak.

  8. Critical review of the trailing edge condition in steady and unsteady flow. Blade flutter in compressors and fans: Numerical simulation of the aerodynamic loading

    NASA Technical Reports Server (NTRS)

    Radwan, S. F.; Rockwell, D. O.; Johnson, S. H.

    1982-01-01

    Existing interpretations of the trailing edge condition, addressing both theoretical and experimental works in steady, as well as unsteady flows are critically reviewed. The work of Kutta and Joukowski on the trailing edge condition in steady flow is reviewed. It is shown that for most practical airfoils and blades (as in the case of most turbomachine blades), this condition is violated due to rounded trailing edges and high frequency effects, the flow dynamics in the trailing edge region being dominated by viscous forces; therefore, any meaningful modelling must include viscous effects. The question of to what extent the trailing edge condition affects acoustic radiation from the edge is raised; it is found that violation of the trailing edge condition leads to significant sound diffraction at the tailing edge, which is related to the problem of noise generation. Finally, various trailing edge conditions in unsteady flow are discussed, with emphasis on high reduced frequencies.

  9. Equations and charts for the rapid estimation of hinge-moment and effectiveness parameters for trailing-edge controls having leading and trailing edges swept ahead of the Mach lines

    NASA Technical Reports Server (NTRS)

    Goin, Kennith L

    1951-01-01

    Existing conical-flow solutions have been used to calculate the hinge-moments and effectiveness parameters of trailing-edge controls having leading and trailing edges swept ahead of the Mach lines and having streamwise root and tip chords. Equations and detailed charts are presented for the rapid estimation of these parameters. Also included is an approximate method by which these parameters may be corrected for airfoil-section thickness.

  10. Symmetric airfoil geometry effects on leading edge noise.

    PubMed

    Gill, James; Zhang, X; Joseph, P

    2013-10-01

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

  11. Symmetric airfoil geometry effects on leading edge noise.

    PubMed

    Gill, James; Zhang, X; Joseph, P

    2013-10-01

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

  12. Mating system shifts on the trailing edge

    PubMed Central

    Levin, Donald A.

    2012-01-01

    Background The trailing edges of species ranges are becoming a subject of increasing interest as the environment changes due to global warming. Trailing edge populations are likely to face extinction because of a decline in numbers and an inability to evolve new adaptations with sufficient speed. Discussions of character change in the trailing edge have focused on physiological, exomorphic and phenological traits. The mating pattern within populations has not been part of the discourse, in spite of the fact that the mating pattern may affect the ability of populations to respond to environmental change and to maintain their sizes. In this paper, the case is made that a substantial increase in self-fertilization rates may occur via plastic responses to stress. Scope and Conclusions Small populations on the trailing edge are especially vulnerable to environmental change because of inadequate levels of cross-fertilization. Evidence is presented that a deficiency of cross-seed production is due to inadequate pollinator services and a paucity of self-incompatibility alleles within populations. Evidence also is presented that if plants are self-compatible, self-fertilization may compensate in part for this deficiency through a stress-induced increase in levels of self-compatibility and stress-induced alterations in floral morphology that elevate self-pollination. Whereas increased self-fertility may afford populations the time to adapt to their changing environments, it can be concluded that increased selfing is not a panacea for the ills of environmental change, because it will lead to substantial reductions in genetic diversity, which may render adaptation unlikely. PMID:21980190

  13. Experimental testing of spanwise morphing trailing edge concept

    NASA Astrophysics Data System (ADS)

    Pankonien, Alexander; Inman, Daniel J.

    2013-04-01

    Aircraft wings with smooth, hinge-less morphing ailerons exhibit increased chordwise aerodynamic efficiency over conventional hinged ailerons. Ideally, the wing would also use these morphing ailerons to smoothly vary its airfoil shape between spanwise stations to optimize the lift distribution and further increase aerodynamic efficiency. However, the mechanical complexity or added weight of achieving such a design has traditionally exceeded the potential aerodynamic gains. By expanding upon the previously developed cascading bimorph concept, this work uses embedded Macro-Fiber Composites and a flexure box mechanism, created using multi-material 3D printing, to achieve the Spanwise Morphing Trailing Edge (SMTE) concept. The morphing actuators are spaced spanwise along the wing with an elastomer spanning the gaps between them, which allows for optimization of the spanwise lift distribution while maintaining the continuity and efficiency of the morphing trailing edge. The concept is implemented in a representative section of a UAV wing with a 305 mm chord. A novel honeycomb skin is created from an elastomeric material using a 3D printer. The actuation capabilities of the concept are evaluated with and without spanning material on a test stand, free of aerodynamic loads. In addition, the actuation restrictions of the spanning elastomer, necessary in adapting the morphing concept from 2D to 3D, are characterized. Initial aerodynamic results from the 1'×1' wind-tunnel also show the effects of aerodynamic loading on the actuation range of the SMTE concept for uniform morphing.

  14. Flow topology and acoustic emissions of trailing edge serrations at incidence

    NASA Astrophysics Data System (ADS)

    Arce León, Carlos; Ragni, Daniele; Pröbsting, Stefan; Scarano, Fulvio; Madsen, Jesper

    2016-05-01

    The flow past a NACA 0018 airfoil with sawtooth trailing edge serrations has been investigated using stereoscopic particle image velocimetry (PIV). The serration flap angle and airfoil incidence are varied in order to study the effect of secondary flow establishing between the suction and pressure sides of the serrations. The flow topology around the serrations is inferred from the analysis of time-averaged streamlines close to the airfoil surface and from the wall-normal flow velocity in between serrations. Additional PIV measurements with a plane in cross-flow highlight the formation of streamwise vortex pairs. The flow behavior is further characterized in terms of its turbulence statistics. Noise emissions are measured with an acoustic phased array in combination with beamforming. The serrations are found to be effective in reducing noise, and their application is studied for different degrees of airfoil incidence and serration flap angle.

  15. Effects of Angle of Attack and Velocity on Trailing Edge Noise

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2006-01-01

    Trailing edge (TE) noise measurements for a NACA 63-215 airfoil model are presented, providing benchmark experimental data for a cambered airfoil. The effects of flow Mach number and angle of attack of the airfoil model with different TE bluntnesses are shown. Far-field noise spectra and directivity are obtained using a directional microphone array. Standard and diagonal removal beamforming techniques are evaluated employing tailored weighting functions for quantitatively accounting for the distributed line character of TE noise. Diagonal removal processing is used for the primary database as it successfully removes noise contaminates. Some TE noise predictions are reported to help interpret the data, with respect to flow speed, angle of attack, and TE bluntness on spectral shape and peak levels. Important findings include the validation of a TE noise directivity function for different airfoil angles of attack and the demonstration of the importance of the directivity function s convective amplification terms.

  16. Effects of Angle of Attack and Velocity on Trailing Edge Noise

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2004-01-01

    Trailing edge (TE) noise measurements for a NACA 63-215 airfoil model are presented, providing benchmark experimental data for a cambered airfoil. The effects of flow Mach number and angle of attack of the airfoil model with different TE bluntnesses are shown. Far-field noise spectra and directivity are obtained using a directional microphone array. Standard and diagonal removal beamforming techniques are evaluated employing tailored weighting functions for quantitatively accounting for the distributed line character of TE noise. Diagonal removal processing is used for the primary database as it successfully removes noise contaminates. Some TE noise predictions are reported to help interpret the data, with respect to flow speed, angle of attack, and TE bluntness on spectral shape and peak levels. Important findings include the validation of a TE noise directivity function for different airfoil angles of attack and the demonstration of the importance of the directivity function s convective amplification terms.

  17. Heat transfer, pressure drop, and mass flow rate in pin fin channels with long and short trailing edge ejection holes

    NASA Astrophysics Data System (ADS)

    Lau, S. C.; Han, J. C.; Batten, T.

    1988-06-01

    The turbulent heat transfer and friction characteristics in the pin fin channels with small trailing edge ejection holes found in internally-cooled turbine airfoils have been experimentally investigated. It is found that the overall heat transfer increases when the length of the trailing edge ejection holes is increased and when the trailing edge ejection holes are configured such that much of the cooling air is forced to flow further downstream in the radial flow direction prior to exiting. The increase in the overall heat transfer is shown to be accompanied by an increase in the overall pressure drop.

  18. Blade-mounted trailing edge flap control for BVI noise reduction

    NASA Technical Reports Server (NTRS)

    Hassan, A. A.; Charles, B. D.; Tadghighi, H.; Sankar, L. N.

    1992-01-01

    Numerical procedures based on the 2-D and 3-D full potential equations and the 2-D Navier-Stokes equations were developed to study the effects of leading and trailing edge flap motions on the aerodynamics of parallel airfoil-vortex interactions and on the aerodynamics and acoustics of the more general self-generated rotor blade vortex interactions (BVI). For subcritical interactions, the 2-D results indicate that the trailing edge flap can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, the results show the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented. For the OLS model rotor, contours of a BVI noise metric were used to quantify the effects of the trailing edge flap on the size and directivity of the high/low intensity noise region(s). Average reductions in the BVI noise levels on the order of 5 dB with moderate power penalties on the order of 18 pct. for a four bladed rotor and 58 pct. for a two bladed rotor were obtained.

  19. Blade-mounted trailing edge flap control for BVI noise reduction

    NASA Astrophysics Data System (ADS)

    Hassan, A. A.; Charles, B. D.; Tadghighi, H.; Sankar, L. N.

    1992-02-01

    Numerical procedures based on the 2-D and 3-D full potential equations and the 2-D Navier-Stokes equations were developed to study the effects of leading and trailing edge flap motions on the aerodynamics of parallel airfoil-vortex interactions and on the aerodynamics and acoustics of the more general self-generated rotor blade vortex interactions (BVI). For subcritical interactions, the 2-D results indicate that the trailing edge flap can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, the results show the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented. For the OLS model rotor, contours of a BVI noise metric were used to quantify the effects of the trailing edge flap on the size and directivity of the high/low intensity noise region(s). Average reductions in the BVI noise levels on the order of 5 dB with moderate power penalties on the order of 18 pct. for a four bladed rotor and 58 pct. for a two bladed rotor were obtained.

  20. Wavy flow cooling concept for turbine airfoils

    DOEpatents

    Liang, George

    2010-08-31

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

  1. 7. Roaring Fork Motor Nature Trail, rocks along edge of ...

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

    7. Roaring Fork Motor Nature Trail, rocks along edge of road. - Great Smoky Mountains National Park Roads & Bridges, Roaring Fork Motor Nature Trail, Between Cherokee Orchard Road & U.S. Route 321, Gatlinburg, Sevier County, TN

  2. The Effect of Nozzle Trailing Edge Thickness on Jet Noise

    NASA Technical Reports Server (NTRS)

    Henderson, Brenda; Kinzie, Kevin; Haskin, Henry

    2004-01-01

    The effect of nozzle trailing edge thickness on broadband acoustic radiation and the production of tones is investigated for coannular nozzles. Experiments were performed for a core nozzle trailing edge thickness between 0.38 mm and 3.17 mm. The on-set of discrete tones was found to be predominantly affected by the velocity ratio, the ratio of the fan velocity to the core velocity, although some dependency on trailing edge thickness was also noted. For a core nozzle trailing edge thickness greater than or equal to 0.89 mm, tones were produced for velocity ratios between 0.91 and 1.61. For a constant nozzle trailing edge thickness, the frequency varied almost linearly with the core velocity. The Strouhal number based on the core velocity changed with nozzle trailing edge thickness and varied between 0.16 and 0.2 for the core nozzles used in the experiments. Increases in broadband noise with increasing trailing edge thickness were observed for tone producing and non-tone producing conditions. A variable thickness trailing edge (crenellated) nozzle resulted in no tonal production and a reduction of the broadband trailing edge noise relative to that of the corresponding constant thickness trailing edge.

  3. Dynamic Stall Characteristics of Drooped Leading Edge Airfoils

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

  4. Identification of secondary instabilities in the near wake of a blunt trailing edge profiled body

    NASA Astrophysics Data System (ADS)

    Cruikshank, Ross; Zhao, Wenyi; Lavoie, Philippe

    2015-11-01

    Aerodynamic research into blunt trailing edge (BTE) airfoils is driven by their structural and aerodynamic advantages over sharp trailing edge airfoils. However, the wake of BTE airfoils is dominated by a vortex street, which causes increased drag. One method to reduce the spanwise coherence of the vortex street is to generate streamwise vorticity in the wake. Recent evidence suggests that the efficiency of this control method can be improved by forcing at the same wavelength as a secondary instability (SI) of the vortex street, present at Reynolds numbers (based on airfoil thickness, d) above 470. The objective of the present study was to investigate the variation of the SI wavelength at 2000 < Red < 35 , 000 , and to examine the effect of forcing on the wake topology. The velocity field in the wake of a BTE profiled model was measured using particle image velocimetry, and proper orthogonal decomposition was applied as a filter for measurement noise. It was found that, for a laminar boundary layer, the SI wavelength decreased as Red increased. Following boundary layer transition to turbulence, the SI wavelength was insensitive to Red . This study will also examine the effect of forcing at different wavelengths on the dominant spanwise wavelength of the wake velocity field. The authors gratefully acknowledge the support of the Natural Sciences and Engineering Research Council of Canada.

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

    NASA Astrophysics Data System (ADS)

    Phan, Nhan Huu

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

  6. Wind turbine trailing edge aerodynamic brakes

    SciTech Connect

    Migliore, P G; Miller, L S; Quandt, G A

    1995-04-01

    Five trailing-edge devices were investigated to determine their potential as wind-turbine aerodynamic brakes, and for power modulation and load alleviation. Several promising configurations were identified. A new device, called the spoiler-flap, appears to be the best alternative. It is a simple device that is effective at all angles of attack. It is not structurally intrusive, and it has the potential for small actuating loads. It is shown that simultaneous achievement of a low lift/drag ratio and high drag is the determinant of device effectiveness, and that these attributes must persist up to an angle of attack of 45{degree}. It is also argued that aerodynamic brakes must be designed for a wind speed of at least 45 m/s (100 mph).

  7. Partially turbulated trailing edge cooling passages for gas turbine nozzles

    DOEpatents

    Thatcher, Jonathan Carl; Burdgick, Steven Sebastian

    2001-01-01

    A plurality of passages are spaced one from the other along the length of a trailing edge of a nozzle vane in a gas turbine. The passages lie in communication with a cavity in the vane for flowing cooling air from the cavity through the passages through the tip of the trailing edge into the hot gas path. Each passage is partially turbulated and includes ribs in an aft portion thereof to provide enhanced cooling effects adjacent the tip of the trailing edge. The major portions of the passages are smooth bore. By this arrangement, reduced temperature gradients across the trailing edge metal are provided. Additionally, the inlets to each of the passages have a restriction whereby a reduced magnitude of compressor bleed discharge air is utilized for trailing edge cooling purposes.

  8. Experimental Investigation of Base Pressure on Blunt-Trailing-Edge Wings at Supersonic Velocities

    NASA Technical Reports Server (NTRS)

    Chapman, Dean R; Wimbrow, William R; Kester, Robert H

    1952-01-01

    Measurements of base pressure are presented for 29 blunt-trailing-edge wings having an aspect ratio of 3.0 and various airfoil profiles. The different profiles comprised thickness ratios between 0.05 and 0.10, boattail angles between -2.9 degrees and 20 degrees, and ratios of trailing-edge thickness to airfoil thickness between 0.2 and 1.0. The tests were conducted at Mach numbers of 1.25, 1.5, 2.0, and 3.1. For each Mach number, the Reynolds number and angle of attack were varied. The lowest Reynolds number investigated was 0.2 x 10(6) and the highest was 3.5 x 10(6). Measurements on each wing were obtained separately with turbulent flow and laminar flow in the boundary layer. Span-wise surveys of the base pressure were conducted on several wings. The results with turbulent boundary-layer flow showed only small effects on base pressure of variations in Reynolds number, airfoil profile shape, boattail angle, and angle of attack. The principal variable affecting the base pressure for turbulent flow was the Mach number.

  9. Modeling the transient aerodynamic effects during the motion of a flexible trailing edge

    NASA Astrophysics Data System (ADS)

    Wolff, T.; Seume, J. R.

    2016-09-01

    Wind turbine blades have been becoming longer and more slender during the last few decades. The longer lever arm results in higher stresses at the blade root. Hence, the unsteady loads induced by turbulence, gust, or wind shear increase. One promising way to control these loads is to use flexible trailing edges near the blade tip. The unsteady effects which appear during the motion of a flexible trailing edge must be considered for the load calculation during the design process because of their high influence on aeroelastic effects and hence on the fatigue loads. This is not yet possible in most of the wind turbine simulation environments. Consequently, an empirical model is developed in the present study which accounts for unsteady effects during the motion of the trailing edge. The model is based on Fourier analyses of results generated with Reynolds-Averaged Navier-Stokes (RANS) simulations of a typical thin airfoil with a deformable trailing edge. The validation showed that the model fits Computational Fluid Dynamics (CFD) results simulated with a random time series of the deflection angle.

  10. Development of Variable Camber Continuous Trailing Edge Flap System

    NASA Technical Reports Server (NTRS)

    Urnes, Jim, Sr.; Nguyen, Nhan T.; Dykman, John

    2012-01-01

    This presentation describes the current status of the joint NASA/Boeing collaboration on the development of a variable camber continuous trailing edge flap system for use in wing shaping control for cruise drag reduction.

  11. 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.

  12. Elastically Deformable Side-Edge Link for Trailing-Edge Flap Aeroacoustic Noise Reduction

    NASA Technical Reports Server (NTRS)

    Khorrami, Mehdi R. (Inventor); Lockard, David P. (Inventor); Moore, James B. (Inventor); Su, Ji (Inventor); Turner, Travis L. (Inventor); Lin, John C. (Inventor); Taminger, Karen M. (Inventor); Kahng, Seun K. (Inventor); Verden, Scott A. (Inventor)

    2014-01-01

    A system is provided for reducing aeroacoustic noise generated by an aircraft having wings equipped with trailing-edge flaps. The system includes a plurality of elastically deformable structures. Each structure is coupled to and along one of the side edges of one of the trailing-edge flaps, and is coupled to a portion of one of the wings that is adjacent to the one of the side edges. The structures elastically deform when the trailing-edge flaps are deployed away from the wings.

  13. Development of a Wind Turbine Test Rig and Rotor for Trailing Edge Flap Investigation: Static Flap Angles Case

    NASA Astrophysics Data System (ADS)

    Abdelrahman, Ahmed; Johnson, David A.

    2014-06-01

    One of the strategies used to improve performance and increase the life-span of wind turbines is active flow control. It involves the modification of the aerodynamic characteristics of a wind turbine blade by means of moveable aerodynamic control surfaces. Trailing edge flaps are relatively small moveable control surfaces placed at the trailing edge of a blade's airfoil that modify the lift of a blade or airfoil section. An instrumented wind turbine test rig and rotor were specifically developed to enable a wide-range of experiments to investigate the potential of trailing edge flaps as an active control technique. A modular blade based on the S833 airfoil was designed to allow accurate instrumentation and customizable settings. The blade is 1.7 meters long, had a constant 178mm chord and a 6° pitch. The modular aerodynamic parts were 3D printed using plastic PC-ABS material. The blade design point was within the range of wind velocities in the available large test facility. The wind facility is a large open jet wind tunnel with a maximum velocity of 11m/s in the test area. The capability of the developed system was demonstrated through an initial study of the effect of stationary trailing edge flaps on blade load and performance. The investigation focused on measuring the changes in flapwise bending moment and power production for different trailing edge flap spanwise locations and deflection angles. The relationship between the load reduction and deflection angle was linear as expected from theory and the highest reduction was caused by the flap furthest from the rotor center. Overall, the experimental setup proved to be effective in measuring small changes in flapwise bending moment within the wind turbine blade and will provide insight when (active) flap control is targeted.

  14. 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.

  15. Application of Passive Porous Treatment to Slat Trailing Edge Noise

    NASA Technical Reports Server (NTRS)

    Khorrami, Mehdi R.; Choudhari, Meelan M.

    2003-01-01

    Porous trailing-edge treatment is investigated as a passive means for slat noise reduction by using time-accurate simulations based on Reynolds-averaged Navier-Stokes equations. For the model scale high-lift configuration used during previous experiments in the Low-Turbulence Pressure Tunnel at NASA Langley Research Center, application of the proposed treatment over a minute fraction of the slat surface area is shown to mitigate the noise impact of the trailing edge, with no measurable aerodynamic penalty. Assessment of the pressure fluctuations in the vicinity of the treated edge indicates a potential noise reduction in excess of 20 dB. The primary mechanism underlying this reduction is related to the reduced strength of Strouhal shedding from the finite thickness trailing edge. A secondary effect of the treatment involves an upward shift in the Strouhal-shedding frequency to a frequency band of reduced auditory sensitivity in a full-scale application.

  16. Design and performance prediction of swashplateless helicopter rotors with trailing edge flaps and tabs

    NASA Astrophysics Data System (ADS)

    Falls, Jaye

    This work studies the design of trailing edge controls for swashplateless helicopter primary control, and examines the impact of those controls on the performance of the rotor. The objective is to develop a comprehensive aeroelastic analysis for swashplateless rotors in steady level flight. The two key issues to be solved for this swashplateless control concept are actuation of the trailing edge controls and evaluating the performance of the swashplateless rotor compared to conventionally controlled helicopters. Solving the first requires simultaneous minimization of trailing flap control angles and hinge moments to reduce actuation power. The second issue requires not only the accurate assessment of swashplateless rotor power, but also similar or improved performance compared to conventional rotors. The analysis consists of two major parts, the structural model and the aerodynamic model. The inertial contributions of the trailing edge flap and tab are derived and added to the system equations in the structural model. Two different aerodynamic models are used in the analysis, a quasi-steady thin airfoil theory that includes arbitrary hinge positions for the flap and the tab, and an unsteady lifting line model with airfoil table lookup based on wind tunnel test data and computational fluid dynamics simulation. The design aspect of the problem is investigated through parametric studies of the trailing edge flap and tab for a Kaman-type conceptual rotor and a UH-60A swashplateless variant. The UH-60A model is not changed except for the addition of a trailing edge flap to the rotor blade, and the reduction of pitch link stiffness to imitate a soft root spring. Study of the uncoupled blade response identifies torsional stiffness and flap hinge stiffness as important design features of the swashplateless rotor. Important trailing edge flap and tab design features including index angle, aerodynamic overhang, chord and length are identified through examination of coupled

  17. Tuning of turbulent boundary layer anisotropy for improved surface pressure and trailing-edge noise modeling

    NASA Astrophysics Data System (ADS)

    Bertagnolio, Franck; Fischer, Andreas; Jun Zhu, Wei

    2014-02-01

    The modeling of the surface pressure spectrum beneath a turbulent boundary layer is investigated, focusing on the case of airfoil flows and associated trailing edge noise prediction using the so-called TNO model. This type of flow is characterized by the presence of an adverse pressure gradient along the airfoil chord. It is shown that discrepancies between measurements and results from the TNO model increase as the pressure gradient increases. The original model is modified by introducing anisotropy in the definition of the turbulent vertical velocity spectrum across the boundary layer and by considering a frequency-dependent vertical correlation length. The degree of anisotropy is directly related to the strength of the pressure gradient. It is shown that by appropriately normalizing the pressure gradient and by tuning the degree of anisotropy, experimental results can be closely reproduced by the modified model. The model is validated against Large Eddy Simulation results and additional wind tunnel measurements. It is further validated in the context of trailing edge noise for which the model formulation makes use of the above surface pressure spectrum.

  18. Structural design of morphing trailing edge actuated by SMA

    NASA Astrophysics Data System (ADS)

    Wang, Qi; Xu, Zhiwei; Zhu, Qian

    2013-09-01

    In this paper, the morphing trailing edge is designed to achieve the up and down deflection under the aerodynamic load. After a detailed and accurate computational analysis to determine the SMA specifications and layout programs, a solid model is created in CATIA and the structures of the morphing wing trailing edge are produced by CNC machining. A set of DSP measurement and control system is designed to accomplish the controlling experiment of the morphing wing trailing edge. At last, via the force analysis, the trailing edge is fabricated with four sections of aluminum alloy, and the arrangement scheme of SMA wires is determined. Experiment of precise control integral has been performed to survey the control effect. The experiment consists of deflection angle tests of the third joint and the integral structure. Primarily, the ultimate deflection angle is tested in these two experiments. Therefore, the controlling experiment of different angles could be performed within this range. The results show that the deflection error is less than 4%and response time is less than 6.7 s, the precise controlling of the morphing trailing edge is preliminary realized.

  19. On the wake flow of asymmetrically beveled trailing edges

    NASA Astrophysics Data System (ADS)

    Guan, Yaoyi; Pröbsting, Stefan; Stephens, David; Gupta, Abhineet; Morris, Scott C.

    2016-05-01

    Trailing edge and wake flows are of interest for a wide range of applications. Small changes in the design of asymmetrically beveled or semi-rounded trailing edges can result in significant difference in flow features which are relevant for the aerodynamic performance, flow-induced structural vibration and aerodynamically generated sound. The present study describes in detail the flow field characteristics around a family of asymmetrically beveled trailing edges with an enclosed trailing-edge angle of 25° and variable radius of curvature R. The flow fields over the beveled trailing edges are described using data obtained by particle image velocimetry (PIV) experiments. The flow topology for different trailing edges was found to be strongly dependent on the radius of curvature R, with flow separation occurring further downstream as R increases. This variation in the location of flow separation influences the aerodynamic force coefficients, which were evaluated from the PIV data using a control volume approach. Two-point correlations of the in-plane velocity components are considered to assess the structure in the flow field. The analysis shows large-scale coherent motions in the far wake, which are associated with vortex shedding. The wake thickness parameter yf is confirmed as an appropriate length scale to characterize this large-scale roll-up motion in the wake. The development in the very near wake was found to be critically dependent on R. In addition, high-speed PIV measurements provide insight into the spectral characteristics of the turbulent fluctuations. Based on the time-resolved flow field data, the frequency range associated with the shedding of coherent vortex pairs in the wake is identified. By means of time-correlation of the velocity components, turbulent structures are found to convect from the attached or separated shear layers without distinct separation point into the wake.

  20. Simulation and Optimization of an Airfoil with Leading Edge Slat

    NASA Astrophysics Data System (ADS)

    Schramm, Matthias; Stoevesandt, Bernhard; Peinke, Joachim

    2016-09-01

    A gradient-based optimization is used in order to improve the shape of a leading edge slat upstream of a DU 91-W2-250 airfoil. The simulations are performed by solving the Reynolds-Averaged Navier-Stokes equations (RANS) using the open source CFD code OpenFOAM. Gradients are computed via the adjoint approach, which is suitable to deal with many design parameters, but keeping the computational costs low. The implementation is verified by comparing the gradients from the adjoint method with gradients obtained by finite differences for a NACA 0012 airfoil. The simulations of the leading edge slat are validated against measurements from the acoustic wind tunnel of Oldenburg University at a Reynolds number of Re = 6 • 105. The shape of the slat is optimized using the adjoint approach resulting in a drag reduction of 2%. Although the optimization is done for Re = 6 • 105, the improvements also hold for a higher Reynolds number of Re = 7.9 • 106, which is more realistic at modern wind turbines.

  1. 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.

  2. Reduction of the shock wave intensity by modifying the transonic blade trailing edge

    NASA Technical Reports Server (NTRS)

    Doerffer, P.

    1983-01-01

    It is shown that the shock wave intensity can be decreased by using modified trailing edge configurations to reduce or even completely compensate for the effect of the finite thickness of the trailing edge. A theoretical analysis is presented together with numerical results for two supersonic streams flowing off the trailing edge at different velocities. The analysis is based on an ideal gas model.

  3. Implementation of a Trailing-Edge Flap Analysis Model in the NASA Langley CAMRAD.MOD1/Hires Program

    NASA Technical Reports Server (NTRS)

    Charles, Bruce

    1999-01-01

    Continual advances in rotorcraft performance, vibration and acoustic characteristics are being sought by rotary-wing vehicle manufacturers to improve efficiency, handling qualities and community noise acceptance of their products. The rotor system aerodynamic and dynamic behavior are among the key factors which must be addressed to meet the desired goals. Rotor aerodynamicists study how airload redistribution impacts performance and noise, and seek ways to achieve better airload distribution through changes in local aerodynamic response characteristics. One method currently receiving attention is the use of trailing-edge flaps mounted on the rotor blades to provide direct control of a portion of the spanwise lift characteristics. The following work describes the incorporation of a trailing-edge flap model in the CAMRAD.Mod1/FHUS comprehensive rotorcraft analysis code. The CAM-RAD.Mod1/HIRES analysis consists of three separate executable codes. These include the comprehensive trim analysis, CAMRAD.Mod1, the Indicial Post-Processor, IPP, for high resolution airloads, and AIRFOIL, which produces the rotor airfoil tables from input airfoil section characteristics. The modifications made to these components permitting analysis of flapped rotor configurations are documented herein along with user instructions detailing the new input variables and operational notes.

  4. Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream

    DOEpatents

    Myers, Robert B.; Yagiela, Anthony S.

    1990-12-25

    An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member.

  5. Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream

    DOEpatents

    Myers, R.B.; Yagiela, A.S.

    1990-12-25

    An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member. 3 figs.

  6. Continuous Trailing-Edge Flaps for Primary Flight Control of a Helicopter Main Rotor

    NASA Technical Reports Server (NTRS)

    Thornburgh, Robert P.; Kreshock, Andrew R.; Wilbur, Matthew L.; Sekula, Martin K.; Shen, Jinwei

    2014-01-01

    The use of continuous trailing-edge flaps (CTEFs) for primary flight control of a helicopter main rotor is studied. A practical, optimized bimorph design with Macro-Fiber Composite actuators is developed for CTEF control, and a coupled structures and computational fluid dynamics methodology is used to study the fundamental behavior of an airfoil with CTEFs. These results are used within a comprehensive rotorcraft analysis model to study the control authority requirements of the CTEFs when utilized for primary flight control of a utility class helicopter. A study of the effect of blade root pitch index (RPI) on CTEF control authority is conducted, and the impact of structural and aerodynamic model complexity on the comprehensive analysis results is presented. The results show that primary flight control using CTEFs is promising; however, a more viable option may include the control of blade RPI, as well.

  7. Transition duct assembly with modified trailing edge in turbine system

    DOEpatents

    McMahan, Kevin Weston; Schott, Carl Gerard; Ingram, Clint Luigie; Siden, Gunnar Leif; Pierre, Sylvain

    2016-10-04

    Transition duct assemblies for turbine systems and turbomachines are provided. In one embodiment, a transition duct assembly includes a plurality of transition ducts disposed in a generally annular array and comprising a first transition duct and a second transition duct. Each of the plurality of transition ducts includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of each transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct assembly further includes an aerodynamic structure defined by the passages of the first transition duct and the second transition duct. The aerodynamic structure includes a pressure side, a suction side, and a trailing edge, the trailing edge having a modified aerodynamic contour.

  8. Stator Loading Measurements Behind a Fan With Trailing Edge Blowing

    NASA Technical Reports Server (NTRS)

    Waitz, Ian A.

    2000-01-01

    The problem of aircraft noise pollution around airports has become increasingly important as those areas have become more densely populated. Currently, the removal of older noisier aircraft from operation is reducing noise levels around airports; however, with air traffic projected to increase by about 5% over the next decade the number of commercial aircraft operating in the world is expected to be about 17,700 by the year 2007. To keep noise levels around airports from increasing as a result of traffic increases, it is important to investigate new methods of noise reduction. The objective of this work is to provide a better understanding of the effects that trailing edge blowing has on stator unsteady loading. This is done by presenting flowfield and stator loading data from experiments conducted with and without trailing edge blowing.

  9. Low-Speed Fan Noise Reduction With Trailing Edge Blowing

    NASA Technical Reports Server (NTRS)

    Sutliff, Daniel L.; Tweedt, Daniel L.; Fite, E. Brian; Envia, Edmane

    2002-01-01

    An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a trailing edge slot. Composite hollow rotor blades with internal flow passages were designed based on analytical codes modeling the internal flow. The hollow blade with interior guide vanes creates flow channels through which externally supplied air flows from the root of the blade to the trailing edge. The impact of the rotor wake-stator interaction on the acoustics was also predicted analytically. The Active Noise Control Fan, located at the NASA Glenn Research Center, was used as the proof- of-concept test bed. In-duct mode and farfield directivity acoustic data were acquired at blowing rates (defined as mass supplied to trailing edge blowing system divided by fan mass flow) ranging from 0.5 to 2.0 percent. The first three blade passing frequency harmonics at fan rotational speeds of 1700 to 1900 rpm were analyzed. The acoustic tone power levels (PWL) in the inlet and exhaust were reduced 11.5 and -0.1, 7.2 and 11.4, 11.8 and 19.4 PWL dB, respectively. The farfield tone power levels at the first three harmonics were reduced 5.4, 10.6, and 12.4 dB PWL. At selected conditions, two-component hotwire and stator vane unsteady surface pressures were acquired. These measurements illustrate the physics behind the noise reduction.

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

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Hatton, Kenneth S.

    1999-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  12. The effect of acoustic forcing on an airfoil tonal noise mechanism.

    PubMed

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

    2014-08-01

    The response of the boundary layer over an airfoil with cavity to external acoustic forcing, across a sweep of frequencies, was measured. The boundary layer downstream of the cavity trailing edge was found to respond strongly and selectively at the natural airfoil tonal frequencies. This is considered to be due to enhanced feedback. However, the shear layer upstream of the cavity trailing edge did not respond at these frequencies. These findings confirm that an aeroacoustic feedback loop exists between the airfoil trailing edge and a location near the cavity trailing edge.

  13. Computational Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.; Walker, Charlotte E.

    1999-01-01

    Computational test cases have been selected from the data set for a clipped delta wing with a six-percent-thick circular-arc airfoil section that was tested in the NASA Langley Transonic Dynamics Tunnel. The test cases include parametric variation of static angle of attack, pitching oscillation frequency, trailing-edge control surface oscillation frequency, and Mach numbers from subsonic to low supersonic values. Tables and plots of the measured pressures are presented for each case. This report provides an early release of test cases that have been proposed for a document that supplements the cases presented in AGARD Report 702.

  14. Measurement of Trailing Edge Noise Using Directional Array and Coherent Output Power Methods

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2002-01-01

    The use of a directional (or phased) array of microphones for the measurement of trailing edge (TE) noise is described and tested. The capabilities of this method arc evaluated via measurements of TE noise from a NACA 63-215 airfoil model and from a cylindrical rod. This TE noise measurement approach is compared to one that is based on thc cross spectral analysis of output signals from a pair of microphones placed on opposite sides of an airframe model (COP method). Advantages and limitations of both methods arc examined. It is shown that the microphone array can accurately measures TE noise and captures its two-dimensional characteristic over a large frequency range for any TE configuration as long as noise contamination from extraneous sources is within bounds. The COP method is shown to also accurately measure TE noise but over a more limited frequency range that narrows for increased TE thickness. Finally, the applicability and generality of an airfoil self-noise prediction method was evaluated via comparison to the experimental data obtained using the COP and array measurement methods. The predicted and experimental results are shown to agree over large frequency ranges.

  15. Measurement Of Trailing Edge Noise using Directional Array and Coherent Output Power Methods

    NASA Technical Reports Server (NTRS)

    Hutcheson, Florence V.; Brooks, Thomas F.

    2002-01-01

    The use of a directional array of microphones for the measurement of trailing edge (TE) noise is described. The capabilities of this method are evaluated via measurements of TE noise from a NACA 63-215 airfoil model and from a cylindrical rod. This TE noise measurement approach is compared to one that is based on the cross spectral analysis of output signals from a pair of microphones (COP method). Advantages and limitations of both methods are examined. It is shown that the microphone array can accurately measures TE noise and captures its two-dimensional characteristic over a large frequency range for any TE configuration as long as noise contamination from extraneous sources is within bounds. The COP method is shown to also accurately measure TE noise but over a more limited frequency range that narrows for increased TE thickness. Finally, the applicability and generality of an airfoil self-noise prediction method was evaluated via comparison to the experimental data obtained using the COP and array measurement methods. The predicted and experimental results are shown to agree over large frequency ranges.

  16. Second-order subsonic airfoil theory including edge effects

    NASA Technical Reports Server (NTRS)

    Van Dyke, Milton D

    1956-01-01

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

  17. A study of the effects of Reynolds number and Mach number on constant pressure coefficient jump for shock-induced trailing-edge separation

    NASA Technical Reports Server (NTRS)

    Cunningham, Atlee M., Jr.; Spragle, Gregory S.

    1987-01-01

    The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied. It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant. The influence of Reynolds number was found to be important but was not strong. Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge. It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge. This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied.

  18. Aeroelastic Response of the Adaptive Compliant Trailing Edge Transtition Section

    NASA Technical Reports Server (NTRS)

    Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat

    2016-01-01

    The Adaptive Compliant Trailing Edge demonstrator was a joint task under the Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan), chartered by the National Aeronautics and Space Administration to develop advanced technologies that enable environmentally friendly aircraft, such as continuous mold-line technologies. The Adaptive Compliant Trailing Edge demonstrator encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys, Inc., a pair of uniquely-designed, unconventional flaps to be used as lifting surfaces during flight-testing to substantiate their structural effectiveness. The unconventional flaps consisted of a main flap section and two transition sections, inboard and outboard, which demonstrated the continuous mold-line technology. Unique characteristics of the transition sections provided a challenge to the airworthiness assessment for this part of the structure. A series of build-up tests and analyses were conducted to ensure the data required to support the airworthiness assessment were acquired and applied accurately. The transition sections were analyzed both as individual components and as part of the flight-test article assembly. Instrumentation was installed in the transition sections based on the analysis to best capture the in-flight aeroelastic response. Flight-testing was conducted and flight data were acquired to validate the analyses. This paper documents the details of the aeroelastic assessment and in-flight response of the transition sections of the unconventional Adaptive Compliant Trailing Edge flaps.

  19. Oscillatory blowing airfoil

    NASA Technical Reports Server (NTRS)

    1997-01-01

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

  20. 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.

  1. Computational Aeroacoustic Analysis of Slat Trailing-Edge Flow

    NASA Technical Reports Server (NTRS)

    Singer, Bart A.; Lockhard, David P.; Brentner, Kenneth S.; Khorrami, Mehdi R.; Berkman, Mert E.; Choudhari, Meelan

    2000-01-01

    An acoustic analysis based on the Ffowcs Williams and Hawkings equation was performed for a high-lift system. As input, the acoustic analysis used un- steady flow data obtained from a highly resolved, time-dependent, Reynolds-averaged Navier-Stokes calculation. The analysis strongly suggests that vor- tex shedding from the trailing edge of the slat results in a high-amplitude, high-frequency acoustic signal, similar to that which was observed in a correspond- ing experimental study of the high-lift system.

  2. Wind-tunnel investigations of wings with serrated sharp trailing edges

    NASA Technical Reports Server (NTRS)

    Vijgen, P. M. H. W.; Van Dam, C. P.; Holmes, B. J.; Howard, F. G.

    1989-01-01

    Exploratory wind-tunnel force measurements are presented for two wing geometries with small-scale planar and nonplanar serrated trailing-edge devices (chord-Reynolds numbers ranged from 1.0 to 3.7 million). The planar serrated trailing-edge extensions reduced the drag at conditions when trailing-edge separation occurred at low angles of attack. The introduction of serrations reduced or eliminated the drag penalty, due to the small (1-2 percent of the chord length) nonplanar trailing-edge flaps, while maintaining the effects of increase in camber. The presence of streamwise vortices immediately downstream of the serrated trailing edges is believed to have favorably affected the boundary-layer flow approaching the trailing edge and the near-wake development, resulting in reduced pressure (form) drag.

  3. Active flow control for a blunt trailing edge profiled body

    NASA Astrophysics Data System (ADS)

    Naghib Lahouti, Arash

    Flow in the wake of nominally two-dimensional bluff bodies is dominated by vortex shedding, beyond a very small threshold Reynolds number. Vortex shedding poses challenges in the design of structures, due to its adverse effects such as cyclic aerodynamic loads and fatigue. The wake vortices are often accompanied by large- and small-scale secondary instabilities, which manifest as dislocations in the primary wake vortices, and/or pairs of counter-rotating streamwise vortices, depending on the dominant instability mode(s), which in turn depends on the profile geometry and Reynolds number. The secondary instabilities interact with the wake vortices through several mechanisms. Therefore, manipulation of the secondary instabilities can be used as a means to alter the wake vortices, in order to reduce their adverse effects. In the present study, flow in the wake of a blunt trailing edge profiled body, composed of an elliptical leading edge and a rectangular trailing edge, has been studied at Reynolds numbers ranging from Re(d) = 500 to 2150 where d is thickness of the body, to identify the secondary instabilities. Various tools, including numerical simulations, Laser Induced Fluorescence (LIF), and Particle Image Velocimetry (PIV) have been used for this study. Proper Orthogonal Decomposition (POD) has been applied to analyze the velocity field data. The results indicate the existence of small-scale instabilities with a spanwise wavelength of 2.0d to 2.5d in the near wake. The mechanism of the instability is similar to the Mode-A instability of a circular cylinder; however, it displays features that are specific to the blunt trailing edge profiled body. An active three-dimensional flow control mechanism based on the small-scale instabilities has been designed and evaluated. The mechanism comprises a series of trailing edge injection ports, with a spanwise spacing equal to the wavelength of the small-scale instabilities. Following preliminary evaluation of the control

  4. A review of the theory of trailing edge noise

    NASA Technical Reports Server (NTRS)

    Howe, M. S.

    1978-01-01

    Literature on the theory of the generation of sound by the interaction of low Mach number turbulent flow with the edge of a semi-infinite rigid plate is critically reviewed. Three different approaches to the subject are identified, consisting of theories based on (1) Lighthill's acoustic analogy; (2) the solution of special, linearized hydroacoustic problems; and (3) ad hoc aerodynamic source models. When appropriately interpreted, all relevant theories produce essentially identical predictions in the limit of very small Mach numbers. None of the theories discusses the implications of the Kutta condition, however, nor of the effect of forward flight and source motion relative to the trailing edge. An outline of a redevelopment of the theory is included to give a unified view of the problem, exhibit the significance of the various approximations, and incorporate the effect of mean motion and of the Kutta condition.

  5. A review of the theory of trailing edge noise

    NASA Technical Reports Server (NTRS)

    Howe, M. S.

    1978-01-01

    A critical review is presented of the literature on the theory of the generation of sound by the interaction of low Mach number turbulent flow with the edge of a semi-infinite rigid plate. Three distinct approaches to the subject are identified, consisting of theories based on (1) Lighthill's acoustic analogy, (2) the solution of special, linearized hydroacoustic problems and (3) ad hoc aerodynamic source models. It is shown that, when appropriately interpreted, all relevant theories produce essentially identical predictions in the limit of very small Mach numbers. In none of the theories, however, are the implications of the Kutta condition discussed, nor of the effect of forward flight and source motion relative to the trailing edge. Accordingly this review includes an outline of a redevelopment of the theory which gives a unified view of the problem, exhibits the significance of the various approximations, and incorporates the effect of mean motion and of the Kutta condition.

  6. Flutter Stability Verified for the Trailing Edge Blowing Fan

    NASA Technical Reports Server (NTRS)

    Bakhle, Milind A.; Srivastava, Rakesh

    2005-01-01

    The TURBO-AE aeroelastic code has been used to verify the flutter stability of the trailing edge blowing (TEB) fan, which is a unique technology demonstrator being designed and fabricated at the NASA Glenn Research Center for testing in Glenn s 9- by 15-Foot Low-Speed Wind Tunnel. Air can be blown out of slots near the trailing edges of the TEB fan blades to fill in the wakes downstream of the rotating blades, which reduces the rotor-stator interaction (tone) noise caused by the interaction of wakes with the downstream stators. The TEB fan will demonstrate a 1.6-EPNdB reduction in tone noise through wake filling. Furthermore, the reduced blade-row interaction will decrease the possibility of forced-response vibrations and enable closer spacing of blade rows, thus reducing engine length and weight. The detailed aeroelastic analysis capability of the three-dimensional Navier-Stokes TURBO-AE code was used to check the TEB fan rotor blades for flutter stability. Flutter calculations were first performed with no TEB flow; then select calculations were repeated with TEB flow turned on.

  7. Airfoil shape for a turbine nozzle

    DOEpatents

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

    2002-01-01

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

  8. Experimental aerodynamics of mesoscale trailing-edge actuators

    NASA Astrophysics Data System (ADS)

    Solovitz, Stephen Adam

    Uninhabited air vehicles (UAVs) are commonly designed with high-aspect ratio wings, which can be susceptible to significant aeroelastic vibrations. These modes can result in a loss of control or structural failure, and new techniques are necessary to alleviate them. A multidisciplinary effort at Stanford developed a distributed flow control method that used small trailing-edge actuators to alter the aerodynamic loads at specific spanwise locations along an airplane wing. This involved design and production of the actuators, computational and experimental study of their characteristics, and application to a flexible wing. This project focused on the experimental response. The actuators were based on a Gurney flap, which is a trailing-edge flap of small size and large deflection, typically about 2% of the chord and 90 degrees, respectively. Because of the large deflection, there is a significant change to the wing camber, increasing the lift. However, due to the small size, the drag does not increase substantially, and the performance is actually improved for high lift conditions. For this project, a 1.5% flap was divided into small span segments (5.2% of the chord), each individually controllable. These devices are termed microflaps or Micro Trailing-edge Effectors (MiTEs). The aerodynamic response was examined to determine the effects of small flap span, the influence of the device structure, and the transient response to relatively rapid MiTE actuation. Measurements included integrated loads, pressure profiles, wake surveys, and near-wake studies using particle image velocimetry. The basic response was similar to a Gurney flap, as full-span actuation of the devices produced a lift increment of about +0.25 when applied towards the pressure surface. For partial actuated spans, the load increment was approximately linear with the actuated span, regardless of configuration. The primary effects occurred within two device spans, indicating that most of the load was

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

    PubMed Central

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

    2016-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Soderman, P. T.

    1972-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Li, Liguo; Li, Zhaohui

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

  12. Atmospheric testing of wind turbine trailing edge aerodynamic brakes

    SciTech Connect

    Miller, L.S.; Migliore, P.G.; Quandt, G.A.

    1997-12-31

    An experimental investigation was conducted using an instrumented horizontal-axis wind turbine that incorporated variable span trailing-edge aerodynamic brakes. A primary goal was to directly compare study results with (infinite-span) wind tunnel data and to provide information on how to account for device span effects during turbine design or analysis. Comprehensive measurements were utilized to define effective changes in the aerodynamic coefficients, as a function of angle of attack and control deflection, for three device spans and configurations. Differences in the lift and drag behavior are most pronounced near stall and for device spans of less than 15%. Drag performance is affected only minimally (<70%) for 15% or larger span devices. Interestingly, aerodynamic controls with characteristic vents or openings appear most affected by span reductions and three-dimensional flow.

  13. Spanwise morphing trailing edge on a finite wing

    NASA Astrophysics Data System (ADS)

    Pankonien, Alexander M.; Inman, Daniel J.

    2015-04-01

    Unmanned Aerial Vehicles are prime targets for morphing implementation as they must adapt to large changes in flight conditions associated with locally varying wind or large changes in mass associated with payload delivery. The Spanwise Morphing Trailing Edge concept locally varies the trailing edge camber of a wing or control surface, functioning as a modular replacement for conventional ailerons without altering the spar box. Utilizing alternating active sections of Macro Fiber Composites (MFCs) driving internal compliant mechanisms and inactive sections of elastomeric honeycombs, the SMTE concept eliminates geometric discontinuities associated with shape change, increasing aerodynamic performance. Previous work investigated a representative section of the SMTE concept and investigated the effect of various skin designs on actuation authority. The current work experimentally evaluates the aerodynamic gains for the SMTE concept for a representative finite wing as compared with a conventional, articulated wing. The comparative performance for both wings is evaluated by measuring the drag penalty associated with achieving a design lift coefficient from an off-design angle of attack. To reduce experimental complexity, optimal control configurations are predicted with lifting line theory and experimentally measured control derivatives. Evaluated over a range of off-design flight conditions, this metric captures the comparative capability of both concepts to adapt or "morph" to changes in flight conditions. Even with this simplistic model, the SMTE concept is shown to reduce the drag penalty due to adaptation up to 20% at off-design conditions, justifying the increase in mass and complexity and motivating concepts capable of larger displacement ranges, higher fidelity modelling, and condition-sensing control.

  14. Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

    NASA Technical Reports Server (NTRS)

    Bennett, Robert M.

    2000-01-01

    Steady and unsteady measured pressures for a Clipped Delta Wing (CDW) undergoing pitching oscillations and trailing-edge control surface oscillations have been presented . From the several hundred compiled data points, 22 static cases, 12 pitching-oscillation cases, and 12 control-surface-oscillation cases have been proposed for Computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The planform for this wing was derived by simplifying the planform of a proposed design for a supersonic transport which is described as the Boeing 2707-300. The strake was deleted, the resulting planform was approximated by a trapezoid with an unswept trailing edge, and the twist and camber were removed. In order to facilitate pressure instrumentation, the thickness was increased to 6 percent from the typical 2.5 to 3 percent for the supersonic transport. The airfoil is thus a symmetrical circular arc section with t/c = 0.06. A wing of similar planform but with a thinner airfoil of t/c = 0.03 was used in the flutter investigations, and the buffet and stall flutter investigation . Flutter results are also reported both for the 3 per cent thick simplified wing and for a more complex SST model. One of the consequences of the increased thickness of the clipped delta wing is that transonic effects are enhanced for Mach numbers near one. They are significantly stronger than would be the case for the thinner wing. Also, with the combination of high leading edge sweep of 50.5, and the sharp leading edge, a leading edge vortex forms on the wing at relatively low angles of attack, on the order of three degrees. The Appendix discusses some of the vortex flow effects. In addition, a shock develops over the aft portion of the wing at transonic speeds such that at some angles of attack, there is both a leading edge vortex and a shock wave on the wing. Such cases are a computational challenge. Some previous applications of this data set have been

  15. Drag Coefficient of Water Droplets Approaching the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia

    2013-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Results are presented and discussed for drag coefficients of droplets with diameters in the range of 300 to 1800 micrometers, and airfoil velocities of 50, 70 and 90 meters/second. The effect of droplet oscillation on the drag coefficient is discussed.

  16. Mechanism of Water Droplet Breakup Near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida, Garcia

    2012-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 microns, and airfoil velocities of 70 and 90 m/sec.

  17. Mechanism of Water Droplet Breakup near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Sor, Suthyvann; Magarino, Adelaida Garcia

    2012-01-01

    This work presents results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de T cnica Aeroespacial (INTA) in Madrid, Spain. The airfoil model was placed at the end of the rotating arm and a monosize droplet generator produced droplets that fell from above, perpendicular to the path of the airfoil. The interaction between the droplets and the airfoil was captured with high speed imaging and allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. Image processing software was used to measure the position of the droplet centroid, equivalent diameter, perimeter, area, and the major and minor axes of an ellipse superimposed over the deforming droplet. The horizontal and vertical displacement of each droplet against time was also measured, and the velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of the droplet to the beginning of breakup. Droplet deformation is defined and studied against main parameters. The high speed imaging allowed observation of the actual mechanism of breakup and identification of the sequence of configurations from the initiation of the breakup to the disintegration of the droplet. Results and comparisons are presented for droplets of diameters in the range of 500 to 1800 micrometers, and airfoil velocities of 70 and 90 meters/second.

  18. The Transient Aerodynamic Forces Effected by Trailing Edge Active Flow Control

    NASA Astrophysics Data System (ADS)

    Brzozowski, Dan; Culp, John; Glezer, Ari

    2012-11-01

    The transient aerodynamic forces effected by trailing edge flow control are investigated in wind tunnel experiments using a 2-DOF traverse which enables application of time-dependent external torque and forces by servo motors. The global aerodynamic forces and moments are regulated by controlling vorticity generation and accumulation near the surface using hybrid synthetic jet actuators. The time-histories of surface pressure and aerodynamic lift and pitching moment immediately following the application of flow control are measured using simultaneous pressure, force and velocity measurements that are taken phase-locked to the commanded actuation waveform. Circulation time history that is estimated from a PIV wake survey shows that the entire flow over the airfoil readjusts within about 1 . 5TCONV , which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales that are commensurate with the flow's convective time scale, and that the maneuver response is primarily limited by the inertia of the platform.

  19. 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.

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

    NASA Technical Reports Server (NTRS)

    Stack, John; Lindsey, W F

    1938-01-01

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

  1. Passive drag control of a blunt trailing edge cylinder

    NASA Astrophysics Data System (ADS)

    Thiria, B.; Cadot, O.; Beaudoin, J.-F.

    2009-07-01

    The turbulent drag of a 2-D bluff body with a blunt trailing edge is investigated for Reynolds numbers in the range 5×10

  2. Atmospheric tests of trailing-edge aerodynamic devices

    SciTech Connect

    Miller, L S; Huang, S; Quandt, G A

    1998-01-01

    An experiment was conducted at the National Renewable Energy Laboratory`s (NREL`s) National Wind Technology Center (NWTC) using an instrumented horizontal-axis wind turbine that incorporated variable-span, trailing-edge aerodynamic brakes. The goal of the investigation was to directly compare results with (infinite-span) wind tunnel data and to provide information on how to account for device span effects during turbine design or analysis. Comprehensive measurements were used to define effective changes in the aerodynamic and hinge-moment coefficients, as a function of angle of attack and control deflection, for three device spans (7.5%, 15%, and 22.5%) and configurations (Spoiler-Flap, vented sileron, and unvented aileron). Differences in the lift and drag behavior are most pronounced near stall and for device spans of less than 15%. Drag performance is affected only minimally (about a 30% reduction from infinite-span) for 15% or larger span devices. Interestingly, aerodynamic controls with vents or openings appear most affected by span reductions and three-dimensional flow.

  3. Structural and aerodynamic considerations of an active piezoelectric trailing-edge tab on a helicopter rotor

    NASA Astrophysics Data System (ADS)

    Murray, Gabriel Jon

    This dissertation is concerned with an active tab for use on a rotorcraft for noise and vibration reduction. The tab is located at the trailing edge of the airfoil. The tab consists of a shim sandwiched by layers of the piezoelectric actuators, macro fiber composites, of varying length. This configuration is similar to a bimorph. The modus operandi is similar to that of a trailing edge flap. The actuators deform the tab, bending it to achieve a tip displacement. This provides a change in the lift, moment, and drag coefficients of the airfoil. By actuating the system at 3/rev to 5/rev, reductions in noise and vibration can be realized. The system was examined and designed around using the UH-60 Blackhawk as the model rotorcraft. The tab is envisioned to operate between 65% to 85% of the main rotor span. The tab's chordwise dimensions considered were 20% and 15% of the blade chord. In order to assess the potential of the tab to change the lift and moment coefficients of the airfoil-tab system, a steady computational fluid dynamics study was conducted. The results were generated via the University of Maryland's Transonic Unsteady Navier-Stokes code. Various tab deflection angles, Mach numbers, and angle-of-attack values were computed. These results were compared to a trailing edge flap of similar size. The comparison shows that the tab produces lift and moment increments similar to that of the trailing edge flap. The design of the tab---composed of both active piezoelectric actuators and passive materials---was conducted using finite element analysis. The objectives were to maximize the tip deflection due to the actuators, while minimizing the deformation due to inertial and aerodynamic forces and loads. The inertial loads (acceleration terms) come from both blade motion, such as flapping and pitch, as well as the rotation of the rotor (centrifugal force). All of these previously mentioned terms cause the tab to undergo undesirable deflections. The original concept

  4. Vibration and local edge buckling of thermally stressed, wedge airfoil cantilever wings.

    NASA Technical Reports Server (NTRS)

    Bailey, C. D.

    1973-01-01

    The local edge buckling phenomena that can occur along the heated thin edge of a wedge shape airfoil is calculated. Qualitative comparison (qualitative only because the experimental temperature distribution was not measured) is made to the experimentally observed phenomena. The consequences of the assumption of identical vibration and buckling modes is shown by a comparison of results with and without the assumption of mode identity. Computer plots of the elastic surface as local buckling develops with increasing temperature are shown. The calculated, fully developed local edge buckling is compared to a photograph of a fully developed buckling as observed in the laboratory.

  5. Experimental Measurement and CFD Model Development of Thick Wind Turbine Airfoils with Leading Edge Erosion

    NASA Astrophysics Data System (ADS)

    Maniaci, David C.; White, Edward B.; Wilcox, Benjamin; Langel, Christopher M.; van Dam, C. P.; Paquette, Joshua A.

    2016-09-01

    Leading edge erosion and roughness accumulation is an issue observed with great variability by wind plant operators, but with little understanding of the effect on wind turbine performance. In wind tunnels, airfoil models are typically tested with standard grit roughness and trip tape to simulate the effects of roughness and erosion observed in field operation, but there is a lack of established relation between field measurements and wind tunnel test conditions. A research collaboration between lab, academic, and industry partners has sought to establish a method to estimate the effect of erosion in wind turbine blades that correlates to roughness and erosion measured in the field. Measurements of roughness and erosion were taken off of operational utility wind turbine blades using a profilometer. The field measurements were statistically reproduced in the wind tunnel on representative tip and midspan airfoils. Simultaneously, a computational model was developed and calibrated to capture the effect of roughness and erosion on airfoil transition and performance characteristics. The results indicate that the effects of field roughness fall between clean airfoil performance and the effects of transition tape. Severe leading edge erosion can cause detrimental performance effects beyond standard roughness. The results also indicate that a heavily eroded wind turbine blade can reduce annual energy production by over 5% for a utility scale wind turbine.

  6. The Prediction of Noise Due to Jet Turbulence Convecting Past Flight Vehicle Trailing Edges

    NASA Technical Reports Server (NTRS)

    Miller, Steven A. E.

    2014-01-01

    High intensity acoustic radiation occurs when turbulence convects past airframe trailing edges. A mathematical model is developed to predict this acoustic radiation. The model is dependent on the local flow and turbulent statistics above the trailing edge of the flight vehicle airframe. These quantities are dependent on the jet and flight vehicle Mach numbers and jet temperature. A term in the model approximates the turbulent statistics of single-stream heated jet flows and is developed based upon measurement. The developed model is valid for a wide range of jet Mach numbers, jet temperature ratios, and flight vehicle Mach numbers. The model predicts traditional trailing edge noise if the jet is not interacting with the airframe. Predictions of mean-flow quantities and the cross-spectrum of static pressure near the airframe trailing edge are compared with measurement. Finally, predictions of acoustic intensity are compared with measurement and the model is shown to accurately capture the phenomenon.

  7. Unsteady Airloads on Airfoils in Reverse Flow

    NASA Astrophysics Data System (ADS)

    Lind, Andrew; Jones, Anya

    2014-11-01

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

  8. Aeroelastic Airworthiness Assesment of the Adaptive Compliant Trailing Edge Flaps

    NASA Technical Reports Server (NTRS)

    Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat; Ervin, Gregory; Flick, Peter

    2015-01-01

    The Adaptive Compliant Trailing Edge (ACTE) demonstrator is a joint task under the National Aeronautics and Space Administration Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan). The project goal is to develop advanced technologies that enable environmentally friendly aircraft, such as adaptive compliant technologies. The ACTE demonstrator flight-test program encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a modified Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys. The control surfaces developed by FlexSys are a pair of uniquely-designed unconventional flaps to be used as lifting surfaces during flight-testing to validate their structural effectiveness. The unconventional flaps required a multidisciplinary airworthiness assessment to prove they could withstand the prescribed flight envelope. Several challenges were posed due to the large deflections experienced by the structure, requiring non-linear analysis methods. The aeroelastic assessment necessitated both conventional and extensive testing and analysis methods. A series of ground vibration tests (GVTs) were conducted to provide modal characteristics to validate and update finite element models (FEMs) used for the flutter analyses for a subset of the various flight configurations. Numerous FEMs were developed using data from FlexSys and the ground tests. The flap FEMs were then attached to the aircraft model to generate a combined FEM that could be analyzed for aeroelastic instabilities. The aeroelastic analysis results showed the combined system of aircraft and flaps were predicted to have the required flutter margin to successfully demonstrate the adaptive compliant technology. This paper documents the details of the aeroelastic airworthiness assessment described, including the ground testing and analyses, and subsequent flight

  9. Numerical prediction of vortex cores of the leading and trailing edges of delta wings

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.

    1980-01-01

    The purpose of the present paper is to predict the roll-up of the vortex sheets emanating from the leading- and trailing-edges of delta wings with emphasis on the interaction of vortex cores beyond the trailing edge. The motivation behind the present work is the recent experimental data published by Hummel. The Nonlinear Discrete-Vortex method (NDV-method) is modified and extended to predict the leading- and trailing-vortex cores beyond the trailing edge. The present model alleviates the problems previously encountered in predicting satisfactory pressure distributions. This is accomplished by lumping the free-vortex lines during the iteration procedure. The leading- and trailing-edge cores and their feeding sheets are obtained as parts of the solution. The numerical results show that the NDV-method is successful in confirming the formation of a trailing-edge core with opposite circulation and opposite roll-up to those of the leading-edge core. This work is a breakthrough in the high angle of attack aerodynamics and moreover, it is the first numerical prediction done on this problem

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

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T. (Inventor)

    1976-01-01

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

  11. 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.

  12. Trailing edges projected to move faster than leading edges for large pelagic fish habitats under climate change

    NASA Astrophysics Data System (ADS)

    Robinson, L. M.; Hobday, A. J.; Possingham, H. P.; Richardson, A. J.

    2015-03-01

    There is mounting evidence to suggest that many species are shifting their ranges in concordance with the climate velocity of their preferred environmental conditions/habitat. While accelerated rates in species' range shifts have been noted in areas of intense warming, due to climate change, few studies have considered the influence that both spatial temperature gradients and rates of warming (i.e., the two components of climate velocity) could have on rates of movement in species habitats. We compared projected shifts in the core habitat of nine large pelagic fish species (five tuna, two billfish and two shark species) off the east coast of Australia at different spatial points (centre, leading and trailing edges of the core habitat), during different seasons (summer and winter), in the near-(2030) and long-term (2070), using independent species distribution models and habitat suitability models. Model projections incorporated depth integrated temperature data from 11 climate models with a focus on the IPCC SRES A2 general emission scenario. Projections showed a number of consistent patterns: southern (poleward) shifts in all species' core habitats; trailing edges shifted faster than leading edges; shifts were faster by 2070 than 2030; and there was little difference in shifts among species and between seasons. Averaging across all species and climate models, rates of habitat shifts for 2030 were 45-60 km decade-1 at the trailing edge, 40-45 km decade-1 at the centre, and 20-30 km decade-1 at the leading edge. Habitat shifts for 2070 were 60-70 km decade-1 at the trailing edge, 50-55 km decade-1 at the centre, and 30-40 km decade-1 at the leading edge. It is often assumed that the leading edge of a species range will shift faster than the trailing edge, but there are few projections or observations in large pelagic fish to validate this assumption. We found that projected shifts at the trailing edge were greater than at the centre and leading of core habitats in

  13. 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.

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

    NASA Astrophysics Data System (ADS)

    Daniel, Libin; Jacob, Jamey

    2013-11-01

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

  15. Navier-Stokes analysis of airfoils with leading edge ice accretions

    NASA Technical Reports Server (NTRS)

    Potapczuk, Mark G.

    1993-01-01

    A numerical analysis of the flowfield characteristics and the performance degradation of an airfoil with leading edge ice accretions was performed. The important fluid dynamic processes were identified and calculated. Among these were the leading edge separation bubble at low angles of attack, complete separation on the low pressure surface resulting in premature shell, drag rise due to the ice shape, and the effects of angle of attack on the separated flow field. Comparisons to experimental results were conducted to confirm these calculations. A computer code which solves the Navier-Stokes equations in two dimensions, ARC2D, was used to perform the calculations. A Modified Mixing Length turbulence model was developed to produce grids for several ice shape and airfoil combinations. Results indicate that the ability to predict overall performance characteristics, such as lift and drag, at low angles of attack is excellent. Transition location is important for accurately determining separation bubble shape. Details of the flowfield in and downstream of the separated regions requires some modifications. Calculations for the stalled airfoil indicate periodic shedding of vorticity that was generated aft of the ice accretion. Time averaged pressure values produce results which compare favorably with experimental information. A turbulence model which accounts for the history effects in the flow may be justified.

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

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  17. Physics behind vortex-induced vibration reduction using an oblique trailing edge hydrofoil

    NASA Astrophysics Data System (ADS)

    Zobeiri, Amirreza; Avellan, Francois; Farhat, Mohamed

    2010-11-01

    The issue of vortex-induced vibration based on the phenomenon of vortex shedding behind a bluff body is a major problem in hydraulic machinery. Resulting fluctuating forces may lead to excessive vibrations and premature cracks. It is well known that a hydrofoil with an oblique trailing edge reduces vibration as compared to that with a blunt trailing edge. However physics behind this is not fully understood. The purpose of the present work is to conduct an experimental investigation of vortex shedding dynamics in the wake of an oblique trailing edge hydrofoil to understand the phenomena and the reasons for vibration reduction. This could help optimize the trailing edge shape and diminish the induced vibration. A velocity survey in the hydrofoil wake is performed via Laser-Doppler and Particle Image velocimetry using the Proper-Orthogonal-Decomposition technique for post-processing. In addition, flow induced vibration measurements and high speed visualization are performed. The high-speed videos clearly demonstrate alternate shedding of the vortices transforming into nearly simultaneous shedding at the hydrofoil trailing edge. As a result, partial cancellation is observed for upper and lower vortices, accompanied by the thickening of the lower vortex core that is believed to be the primary reason of the vibration reduction.

  18. 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.

  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. Simulated big sagebrush regeneration supports predicted changes at the trailing and leading edges of distribution shifts

    USGS Publications Warehouse

    Schlaepfer, Daniel R.; Taylor, Kyle A.; Pennington, Victoria E.; Nelson, Kellen N.; Martin, Trace E.; Rottler, Caitlin M.; Lauenroth, William K.; Bradford, John B.

    2015-01-01

    Many semi-arid plant communities in western North America are dominated by big sagebrush. These ecosystems are being reduced in extent and quality due to economic development, invasive species, and climate change. These pervasive modifications have generated concern about the long-term viability of sagebrush habitat and sagebrush-obligate wildlife species (notably greater sage-grouse), highlighting the need for better understanding of the future big sagebrush distribution, particularly at the species' range margins. These leading and trailing edges of potential climate-driven sagebrush distribution shifts are likely to be areas most sensitive to climate change. We used a process-based regeneration model for big sagebrush, which simulates potential germination and seedling survival in response to climatic and edaphic conditions and tested expectations about current and future regeneration responses at trailing and leading edges that were previously identified using traditional species distribution models. Our results confirmed expectations of increased probability of regeneration at the leading edge and decreased probability of regeneration at the trailing edge below current levels. Our simulations indicated that soil water dynamics at the leading edge became more similar to the typical seasonal ecohydrological conditions observed within the current range of big sagebrush ecosystems. At the trailing edge, an increased winter and spring dryness represented a departure from conditions typically supportive of big sagebrush. Our results highlighted that minimum and maximum daily temperatures as well as soil water recharge and summer dry periods are important constraints for big sagebrush regeneration. Overall, our results confirmed previous predictions, i.e., we see consistent changes in areas identified as trailing and leading edges; however, we also identified potential local refugia within the trailing edge, mostly at sites at higher elevation. Decreasing

  1. Performance of S-cambered profiles with cut-off trailing edges

    SciTech Connect

    Chacko, B.; Balabaskaran, V.; Tulapurkara, E.G.; Krishna, H.C.R. )

    1994-09-01

    The aerodynamic characteristics of an S-cambered profile are studies under forward and reversed flow conditions. The profile chord is cut by 3, 6, and 9 percent of the chord at the sharp trailing edge end and the performances of these profiles are compared. It is found that with increase in length of cutting the lift coefficient increases in forward direction and decreases in reverse direction of flow. cutting off the sharp trailing edge improves the lift-drag characteristics in forward mode and deteriorates in the reverse mode.

  2. Experimental Observations on the Deformation and Breakup of Water Droplets Near the Leading Edge of an Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Feo, Alex

    2011-01-01

    This work presents the results of an experimental study on droplet deformation and breakup near the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the Instituto Nacional de Tecnica Aeroespacial (INTA) in Madrid, Spain. An airfoil model placed at the end of the rotating arm was moved at speeds of 50 to 90 m/sec. A monosize droplet generator was employed to produce droplets that were allowed to fall from above, perpendicular to the path of the airfoil at a given location. High speed imaging was employed to observe the interaction between the droplets and the airfoil. The high speed imaging allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. A tracking software program was used to measure from the high speed movies the horizontal and vertical displacement of the droplet against time. The velocity, acceleration, Weber number, Bond number, Reynolds number, and the drag coefficients were calculated along the path of a given droplet from beginning of deformation to breakup and/or hitting the airfoil. Results are presented for droplets with a diameter of 490 micrometers at airfoil speeds of 50, 60, 70, 80 and 90 m/sec

  3. An experimental investigation of convective heat transfer at the leading edge of a gas turbine airfoil

    NASA Astrophysics Data System (ADS)

    Gendron, S.; Marchand, N. J.; Korn, C.; Immarigeon, J. P.; Kacprzynski, J. J.

    1992-06-01

    This paper describes the experimental methods used to determine the surface temperatures and heat-transfer coefficients at the leading edge, and elsewhere over the surface, of a specially designed double-edge wedge shell specimen subjected to cyclic heating in a high velocity hot gas stream generated by a burner rig. The methods included temperature measurements with thermocouples (embedded below the surface) as well as surface temperature measurements by optical pyrometry. The experiments were carried out at gas temperatures between 806 to 1323 C and velocities in the range from Mach 0.32 to Mach 0.39. The calibration procedures for each method, the various testing conditions to which the airfoil-like specimen was exposed and the results pertaining to the determination of the surface temperatures and heat-transfer coefficients are described and discussed.

  4. Wind turbine trailing-edge aerodynamic brake design

    SciTech Connect

    Quandt, G.

    1996-01-01

    This report describes the design of a centrifugally actuated aerodynamic-overspeed device for a horizontal-axis wind turbine. The device will meet the following criteria; (1) It will be effective for airfoil angles of attack 0{degrees} to 45{degrees}. (2) It will be stowed inside the blade profile prior to deployment. (3) It will be capable of offsetting the positive torque produced by the overall blade. (4) Hinge moments will be minimized to lower actuator loads and cost. (5) It will be evaluated as a potential power modulating active rotor-control system. A literature review of aerodynamic braking devices was conducted. Information from the literature review was used to conceptualize the most effective devices for subsequent testing and design. Wind-tunnel test data for several braking devices are presented in this report. Using the data for the most promising configuration, a preliminary design was developed for a MICON 65/13 wind turbine with Phoenix 7.9-m rotor blades.

  5. Cascade trailing-edge noise modeling using a mode-matching technique and the edge-dipole theory

    NASA Astrophysics Data System (ADS)

    Roger, Michel; François, Benjamin; Moreau, Stéphane

    2016-11-01

    An original analytical approach is proposed to model the broadband trailing-edge noise produced by high-solidity outlet guide vanes in an axial turbomachine. The model is formulated in the frequency domain and first in two dimensions for a preliminary assessment of the method. In a first step the trailing-edge noise sources of a single vane are shown to be equivalent to the onset of a so-called edge dipole, the direct field of which is expanded in a series of plane-wave modes. A criterion for the distance of the dipole to the trailing-edge and a scaling of its amplitude is defined to yield a robust model. In a second step the diffraction of each plane-wave mode is derived considering the cascade as an array of bifurcated waveguides and using a mode-matching technique. The cascade response is finally synthesized by summing the diffracted fields of all cut-on modes to yield upstream and downstream sound power spectral densities. The obtained spectral shapes are physically consistent and the present results show that upstream radiation is typically 3 dB higher than downstream radiation, which has been experimentally observed previously. Even though the trailing-edge noise sources are not vane-to-vane correlated their radiation is strongly determined by a cascade effect that consequently must be accounted for. The interest of the approach is that it can be extended to a three-dimensional annular configuration without resorting to a strip theory approach. As such it is a promising and versatile alternative to previously published methods.

  6. Incidence angle effects on convected gust airfoil noise

    NASA Astrophysics Data System (ADS)

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

    1983-04-01

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

  7. Broadband Noise Reduction of a Low-Speed Fan Noise Using Trailing Edge Blowing

    NASA Technical Reports Server (NTRS)

    Sutliff, Daniel L.

    2005-01-01

    An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through the use of rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a continuous trailing edge slot. Hollow blades with interior guide vanes create flow channels through which externally supplied air flows from the blade root to the trailing edge. A previous paper documented the substantial tonal reductions of this Trailing Edge Rotor Blowing (TERB) fan. This report documents the broadband characteristics of TERB. The Active Noise Control Fan (ANCF), located at the NASA Glenn Research Center, was used as the proof-of-concept test bed. Two-component hotwire data behind the rotor, unsteady surface pressures on the stator vane, and farfield directivity acoustic data were acquired at blowing rates of 1.1, 1.5, and 1.8 percent of the total fan mass flow. The results indicate a substantial reduction in the rotor wake turbulent velocity and in the stator vane unsteady surface pressures. Based on the physics of the noise generation, these indirect measurements indicate the prospect of broadband noise reduction. However, since the broadband noise generated by the ANCF is rotor-dominated, any change in the rotor-stator interaction broadband noise levels is barely distinguishable in the farfield measurements.

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

    SciTech Connect

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

    1987-07-01

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

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

    NASA Astrophysics Data System (ADS)

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

    1987-07-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  12. Suppression of dynamic stall with a leading-edge slat on a VR-7 airfoil

    NASA Technical Reports Server (NTRS)

    Mcalister, K. W.; Tung, C.

    1993-01-01

    The VR-7 airfoil was experimentally studied with and without a leading-edge slat at fixed angles of attack from 0 deg to 30 deg at Re = 200,000 and for unsteady pitching motions described by alpha equals alpha(sub m) + 10 deg(sin(wt)). The models were two dimensional, and the test was performed in a water tunnel at Ames Research Center. The unsteady conditions ranged over Re equals 100,000 to 250,000, k equals 0.001 to 0.2, and alpha(sub m) = 10 deg to 20 deg. Unsteady lift, drag, and pitching-moment measurements were obtained along with fluorescent-dye flow visualizations. The addition of the slat was found to delay the static-drag and static-moment stall by about 5 degrees and to eliminate completely the development of a dynamic-stall vortex during unsteady motions that reached angles as high as 25 degrees. In all of the unsteady cases studied, the slat caused a significant reduction in the force and moment hysteresis amplitudes. The reduced frequency was found to have the greatest effect on the results, whereas the Reynolds number had little effect on the behavior of either the basic or the slatted airfoil. The slat caused a slight drag penalty at low angles of attack, but generally increased the lift/drag ratio when averaged over the full cycle of oscillation.

  13. Active Management of Flap-Edge Trailing Vortices

    NASA Technical Reports Server (NTRS)

    Greenblatt, David; Yao, Chung-Sheng; Vey, Stefan; Paschereit, Oliver C.; Meyer, Robert

    2008-01-01

    The vortex hazard produced by large airliners and increasingly larger airliners entering service, combined with projected rapid increases in the demand for air transportation, is expected to act as a major impediment to increased air traffic capacity. Significant reduction in the vortex hazard is possible, however, by employing active vortex alleviation techniques that reduce the wake severity by dynamically modifying its vortex characteristics, providing that the techniques do not degrade performance or compromise safety and ride quality. With this as background, a series of experiments were performed, initially at NASA Langley Research Center and subsequently at the Berlin University of Technology in collaboration with the German Aerospace Center. The investigations demonstrated the basic mechanism for managing trailing vortices using retrofitted devices that are decoupled from conventional control surfaces. The basic premise for managing vortices advanced here is rooted in the erstwhile forgotten hypothesis of Albert Betz, as extended and verified ingeniously by Coleman duPont Donaldson and his collaborators. Using these devices, vortices may be perturbed at arbitrarily long wavelengths down to wavelengths less than a typical airliner wingspan and the oscillatory loads on the wings, and hence the vehicle, are small. Significant flexibility in the specific device has been demonstrated using local passive and active separation control as well as local circulation control via Gurney flaps. The method is now in a position to be tested in a wind tunnel with a longer test section on a scaled airliner configuration. Alternatively, the method can be tested directly in a towing tank, on a model aircraft, a light aircraft or a full-scale airliner. The authors believed that this method will have significant appeal from an industry perspective due to its retrofit potential with little to no impact on cruise (devices tucked away in the cove or retracted); low operating power

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

    NASA Astrophysics Data System (ADS)

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

    2014-04-01

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

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

    SciTech Connect

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

    2010-06-01

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

  16. Effect of Trailing Edge Flow Injection on Fan Noise and Aerodynamic Performance

    NASA Technical Reports Server (NTRS)

    Fite, E. Brian; Woodward, Richard P.; Podboy, Gary G.

    2006-01-01

    An experimental investigation using trailing edge blowing for reducing fan rotor/guide vane wake interaction noise was completed in the NASA Glenn 9- by 15-foot Low Speed Wind Tunnel. Data were acquired to measure noise, aerodynamic performance, and flow features for a 22" tip diameter fan representative of modern turbofan technology. The fan was designed to use trailing edge blowing to reduce the fan blade wake momentum deficit. The test objective was to quantify noise reductions, measure impacts on fan aerodynamic performance, and document the flow field using hot-film anemometry. Measurements concentrated on approach, cutback, and takeoff rotational speeds as those are the primary conditions of acoustic interest. Data are presented for a 2% (relative to overall fan flow) trailing edge injection rate and show a 2 dB reduction in Overall Sound Power Level (OAPWL) at all fan test speeds. The reduction in broadband noise is nearly constant and is approximately 1.5 dB up to 20 kHz at all fan speeds. Measurements of tone noise show significant variation, as evidenced by reductions of up to 6 dB in the 2 BPF tone at 6700 rpm.: and increases of nearly 2 dB for the 4 BPF tone at approach speed. Aerodynamic performance measurements show the fan with 2 % injection has an overall efficiency that is comparable to the baseline fan and operates, as intended, with nearly the same pressure ratio and mass flow parameters. Hot-film measurements obtained at the approach operating condition indicate that mean blade wake filling in the tip region was not as significant as expected. This suggests that additional acoustic benefits could be realized if the trailing edge blowing could be modified to provide better filling of the wake momentum deficit. Nevertheless, the hot-film measurements indicate that the trailing edge blowing provided significant reductions in blade wake turbulence. Overall, these results indicate that further work may be required to fully understand the proper

  17. Wake instabilities of a blunt trailing edge profiled body at intermediate Reynolds numbers

    NASA Astrophysics Data System (ADS)

    Naghib-Lahouti, A.; Lavoie, P.; Hangan, H.

    2014-07-01

    Experiments have been conducted to identify and characterize the instabilities in the wake of a blunt trailing edge profiled body, comprised of an elliptical leading edge and a rectangular trailing edge, for a broad range of Reynolds numbers ( based on the thickness of the body). These experiments, which include measurements of the wake velocity field using hot-wire anemometry and particle image velocimetry, complement previous studies of the wake flow for the same geometry at lower and higher Reynolds numbers. The spatial characteristics of the primary wake instability (the von Kármán vortex street) are found to have relatively little variation in the range of Reynolds numbers investigated, in spite of the transition of the boundary layer upstream of the trailing edge from a laminar to a turbulent state. The dominant secondary instability, identified based on the structure of velocity and vorticity fields in the wake extracted using proper orthogonal decomposition, is found to have features similar to the ones described numerically and experimentally by Ryan et al. (J Fluid Mech 538:1-29, 2005), and Naghib-Lahouti et al. (Exp Fluids 52:1547-1566, 2012) at lower Reynolds numbers. The findings suggest that the spatial characteristics of the dominant primary and secondary wake flow instabilities have little dependence on the state of the flow upstream of the separation points, in spite of the distinct change in the normalized vortex shedding frequency upon the transition of the boundary layer.

  18. Secondary wake instabilities of a blunt trailing edge profiled body as a basis for flow control

    NASA Astrophysics Data System (ADS)

    Naghib-Lahouti, Arash; Doddipatla, Lakshmana Sampat; Hangan, Horia

    2012-06-01

    Flow in the wake of a blunt trailing edge profiled body, comprised of an elliptical leading edge and a rectangular trailing edge, has been investigated experimentally, to identify and characterize the secondary instabilities accompanying the von Kármán vortices. The experiments, which involve laser-induced fluorescence for visualization and particle image velocimetry for quantitative measurement of the wake instabilities, cover Reynolds numbers ranging from 250 to 2,150 based on thickness of the body, to include the wake transition regime. The dominant secondary instability appears as spanwise undulations in von Kármán vortices, which evolve into pairs of counter-rotating vortices, with features resembling the instability mechanism predicted by Ryan et al. (J Fluid Mech 538:1-29, 2005). Feasibility of a flow control approach based on interaction with the secondary instability using a series of discrete trailing edge injectors has also been investigated. The control approach mitigates the adverse effects of vortex shedding in certain conditions, where it is able to amplify the secondary instability effectively.

  19. Wake Control of a Blunt Trailing Edge Profiled Body Using Dielectric Barrier Discharge Plasma Actuators

    NASA Astrophysics Data System (ADS)

    Naghib-Lahouti, Arash; Lavoie, Philippe

    2012-11-01

    The periodic shedding of von Karman vortices is the source of cyclic aerodynamic forces on nominally 2D bluff bodies. Beyond a threshold Reynolds number, which can be as high as 700 depending on profile geometry, secondary instabilities, appearing as undulations in the von Karman vortices and pairs of counter-rotating streamwise vortices, emerge in the wake. The secondary instabilities are found to persist at Reynolds numbers in the order of 104. It has been shown that amplification of the secondary instabilities can lead to disorganization of the von Karman vortices, and attenuation of the cyclic forces. In the present study, this relationship is used as the basis of a wake flow control approach for a blunt trailing edge profiled body, comprised of an elliptical leading edge and a rectangular trailing edge. An array of dielectric barrier discharge plasma actuators placed at the trailing edge is used for control actuation, with a spanwise spacing based on the wavelength of the secondary instabilities, to achieve maximum amplification of the instabilities. PIV and hot-wire measurements have been conducted at Reynolds numbers between 2,000 and 24,000 to determine the effect of flow control on the wake characteristics, and the total drag. Funded in part by the Government of Ontario, Mitacs, and Bombardier Aerospace.

  20. Flow visualization of vortices locked by spanwise blowing over wings featuring a unique leading and trailing-edge flap system

    NASA Technical Reports Server (NTRS)

    Erickson, G. E.; Campbell, J. F.

    1975-01-01

    Flow visualization studies were conducted to qualitatively determine the effects of active generation and augmentation of vortex flow over wings by blowing a discrete jet in a spanwise direction in the channel formed by extension of upper surface leading- and trailing-edge flaps. Spanwise blowing from a reflection plane over a rectangular wing was found to generate and lock a dual corotating vortex system within the channel and, at sufficient blowing rates, cause the separated flow off the upper end of the leading-edge flap to reattach to the trailing-edge flap. Test parameters included wing angle of attack, jet momentum coefficient, leading- and trailing-edge flap deflection angle, and jet location above the wing surface. Effects due to removal of the leading- and trailing-edge flap were also investigated.

  1. Performance of Swashplateless Ultralight Helicopter Rotor with Trailing-edge Flaps for Primary Flight Control

    NASA Technical Reports Server (NTRS)

    Shen, Jin-Wei; Chopra, Inderjit

    2003-01-01

    The objective of present study is to evaluate the rotor performance, trailing-edge deflections and actuation requirement of a helicopter rotor with trailing-edge flap system for primary flight control. The swashplateless design is implemented by modifying a two-bladed teetering rotor of an production ultralight helicopter through the use of plain flaps on the blades, and by replacing the pitch link to fixed system control system assembly with a root spring. A comprehensive rotorcraft analysis based on UMARC is carried out to obtain the results for both the swashplateless and a conventional baseline rotor configuration. The predictions show swashplateless configuration achieve superior performance than the conventional rotor attributed from reduction of parasite drag by eliminating swashplate mechanic system. It is indicated that optimal selection of blade pitch index angle, flap location, length, and chord ratio reduces flap deflections and actuation requirements, however, has virtually no effect on rotor performance.

  2. Trailing edge cooling using angled impingement on surface enhanced with cast chevron arrangements

    SciTech Connect

    Lee, Ching-Pang; Heneveld, Benjamin E.; Brown, Glenn E.; Klinger, Jill

    2015-05-26

    A gas turbine engine component, including: a pressure side (12) having an interior surface (34); a suction side (14) having an interior surface (36); a trailing edge portion (30); and a plurality of suction side and pressure side impingement orifices (24) disposed in the trailing edge portion (30). Each suction side impingement orifice is configured to direct an impingement jet (48) at an acute angle (52) onto a target area (60) that encompasses a tip (140) of a chevron (122) within a chevron arrangement (120) formed in the suction side interior surface. Each pressure side impingement orifice is configured to direct an impingement jet at an acute angle onto an elongated target area that encompasses a tip of a chevron within a chevron arrangement formed in the pressure side interior surface.

  3. Numerical study of improving aerodynamic performance of low solidity LPT cascade through increasing trailing edge thickness

    NASA Astrophysics Data System (ADS)

    Li, Chao; Yan, Peigang; Wang, Xiangfeng; Han, Wanjin; Wang, Qingchao

    2016-08-01

    This paper presents a new idea to reduce the solidity of low-pressure turbine (LPT) blade cascades, while remain the structural integrity of LPT blade. Aerodynamic performance of a low solidity LPT cascade was improved by increasing blade trailing edge thickness (TET). The solidity of the LPT cascade blade can be reduced by about 12.5% through increasing the TET of the blade without a significant drop in energy efficiency. For the low solidity LPT cascade, increasing the TET can decrease energy loss by 23.30% and increase the flow turning angle by 1.86% for Reynolds number (Re) of 25,000 and freestream turbulence intensities (FSTI) of 2.35%. The flow control mechanism governing behavior around the trailing edge of an LPT cascade is also presented. The results show that appropriate TET is important for the optimal design of high-lift load LPT blade cascades.

  4. The effect of a cavity on airfoil tones

    NASA Astrophysics Data System (ADS)

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

    2014-03-01

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

  5. New Fan Engine Noise-Reduction Concept Using Trailing Edge Blowing of Fan Blades Demonstrated

    NASA Technical Reports Server (NTRS)

    Heidelberg, Laurence J.

    2002-01-01

    A major source of noise in commercial turbofan engines is the interaction of the fan blade wakes with the fan exit vanes (stators). These wakes can be greatly reduced by filling them with air blown out of the blade trailing edge. Extensive testing of this concept has demonstrated significant noise reductions. These tests were conducted on a low-speed, 4- ft-diameter fan using hollow blades at NASA Glenn Research Center's Aeroacoustic Propulsion Laboratory (AAPL).

  6. Flat Plate Wake Velocity Statistics Obtained With Circular And Elliptic Trailing Edges

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan

    2016-01-01

    The near wake of a flat plate with circular and elliptic trailing edges is investigated with data from direct numerical simulations. The plate length and thickness are the same in both cases. The separating boundary layers are turbulent and statistically identical. Therefore the wake is symmetric in the two cases. The emphasis in this study is on a comparison of the wake-distributions of velocity components, normal intensity and fluctuating shear stress obtained in the two cases.

  7. Serrated trailing edges for improving lift and drag characteristics of lifting surfaces

    NASA Technical Reports Server (NTRS)

    Vijgen, Paul M. H. W. (Inventor); Howard, Floyd G. (Inventor); Bushnell, Dennis M. (Inventor); Holmes, Bruce J. (Inventor)

    1992-01-01

    An improvement in the lift and drag characteristics of a lifting surface is achieved by attaching a serrated panel to the trailing edge of the lifting surface. The serrations may have a saw-tooth configuration, with a 60 degree included angle between adjacent serrations. The serrations may vary in shape and size over the span-wise length of the lifting surface, and may be positioned at fixed or adjustable deflections relative to the chord of the lifting surface.

  8. Aeroelastic control of flutter using trailing edge control surfaces powered by piezoelectric actuators

    NASA Astrophysics Data System (ADS)

    Ardelean, Emil Valentin

    Flutter is a rather spectacular phenomenon of aeroelastic instability that affects lifting and control surfaces, yet can also lead to catastrophic consequences for the aircraft. The idea of controlling flutter by using the same energy that causes it, namely airflow energy, through changing the aerodynamics in a controlled manner is not new. In the case of fixed wings, the use of trailing edge control surfaces (flaps) is an extremely effective method to alter the aerodynamics. This research presents the development of an actuation system for trailing edge control surfaces (flaps) used for aeroelastic flutter control of a typical section wing model. In order to be effective for aeroelastic control of flutter, flap deflection of +/-5-6° with adequate bandwidth (up to 25--30 Hz) is required. Classical solutions for flap actuation do not have the capabilities required for this task. Therefore actuation systems using active materials became the focus of this investigation. A new piezoelectric actuator (V-Stack Piezoelectric Actuator) was developed. This actuator meets the requirements for trailing edge flap actuation in both stroke and force over the bandwidth of interest. It is compact, simple, sturdy, and leverages stroke geometrically with minimum force penalties, while displaying linearity over a wide range of stroke. Integration of the actuator inside an existing structure requires minimal modifications of the structure. The shape of the actuator makes it very suitable for trailing edge flap actuation, eliminating the need for a push rod. The actuation solution presented here stands out because of its simplicity, compactness, small mass (compared to that of the actuated structure) and high reliability. Although the actuator was designed for flap actuation, other applications can also benefit from its capabilities. In order to demonstrate the actuation concept, a typical section prototype was constructed and tested experimentally in the wind tunnel at Duke

  9. The symmetric turbulent plane wake downstream of a sharp trailing edge

    NASA Technical Reports Server (NTRS)

    Bogucz, E. A.

    1991-01-01

    The analysis and modeling of the symmetric turbulent plane wake downstream of a sharp trailing edge is addressed. A compact description of the flow near the trailing edge is formulated using the results of a previous asymptotic analysis. The new description retains the two-layered structure identified in the previous work, and it clarifies the principal dynamics of the flow in the near-wake outer layer, away from the wake centerline. For zero-pressure-gradient flow, the near-wake outer layer is shown to be represented to leading order by the similarity solution that governs the outer region of the surface boundary layer. The leading perturbation in the outer layer due to the developing near-wake inner-layer flow is identified, and this is shown to be asymptotically smaller than undetermined higher-order terms associated with the surface boundary-layer flow. Results of the new near-wake analysis are used to formulate an algebraic eddy viscosity model for wake flow predictions at arbitrary distances from the trailing edge. The model is used in a numerical solution of the boundary layer equations, and computed velocity and Reynolds stress profiles are shown to compare well with experimental data.

  10. Unsteady flow past an airfoil pitched at constant rate

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

    2013-11-01

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

  12. On the acoustic radiation of a pitching airfoil

    NASA Astrophysics Data System (ADS)

    Manela, A.

    2013-07-01

    We examine the acoustic far field of a thin elastic airfoil, immersed in low-Mach non-uniform stream flow, and actuated by small-amplitude sinusoidal pitching motion. The near-field fluid-structure interaction problem is analyzed using potential thin-airfoil theory, combined with a discrete vortex model to describe the evolution of airfoil trailing edge wake. The leading order dipole-sound signature of the system is investigated using Powell-Howe acoustic analogy. Compared with a pitching rigid airfoil, the results demonstrate a two-fold effect of structure elasticity on airfoil acoustic field: at actuation frequencies close to the system least stable eigenfrequency, elasticity amplifies airfoil motion amplitude and associated sound levels; however, at frequencies distant from this eigenfrequency, structure elasticity acts to absorb system kinetic energy and reduce acoustic radiation. In the latter case, and with increasing pitching frequency ωp, a rigid-airfoil setup becomes significantly noisier than an elastic airfoil, owing to an ω _p^{5/2} increase of its direct motion noise component. Unlike rigid airfoil signature, it is shown that wake sound contribution to elastic airfoil radiation is significant for all ωp. Remarkably, this contribution contains, in addition to the fundamental pitching frequency, its odd multiple harmonics, which result from nonlinear interactions between the airfoil and the wake. The results suggest that structure elasticity may serve as a viable means for design of flapping flight noise control methodologies.

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

    NASA Astrophysics Data System (ADS)

    Liu, Yixiong; Yang, Ce; Song, Xiancheng

    2015-04-01

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

  14. 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.

  15. 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.

  16. Structural and acoustic noise produced by turbulent flow over an elastic trailing edge

    NASA Astrophysics Data System (ADS)

    Howe, M. S.

    1993-09-01

    An analysis is made of the sound and vibration produced by turbulent flow at low Mach number over the trailing edge of an elastic plate. General formulas are developed for the structural and acoustic edge-noise when the control surface is modeled by a semiinfinite thin elastic plate which can support bending waves. Numerical results are given for steel plates in air and in water. In the latter case it is shown that, when the frequency is smaller than the coincidence frequency, the bending wave power exceeds the total sound power generated at the edge by 20-40 dB, independently of the mean flow velocity, so that sound generated by secondary scattering may then be the dominant source of acoustic radiation.

  17. Validation and Analysis of Numerical Results for a Two-Pass Trapezoidal Channel With Different Cooling Configurations of Trailing Edge.

    PubMed

    Siddique, Waseem; El-Gabry, Lamyaa; Shevchuk, Igor V; Fransson, Torsten H

    2013-01-01

    High inlet temperatures in a gas turbine lead to an increase in the thermal efficiency of the gas turbine. This results in the requirement of cooling of gas turbine blades/vanes. Internal cooling of the gas turbine blade/vanes with the help of two-pass channels is one of the effective methods to reduce the metal temperatures. In particular, the trailing edge of a turbine vane is a critical area, where effective cooling is required. The trailing edge can be modeled as a trapezoidal channel. This paper describes the numerical validation of the heat transfer and pressure drop in a trapezoidal channel with and without orthogonal ribs at the bottom surface. A new concept of ribbed trailing edge has been introduced in this paper which presents a numerical study of several trailing edge cooling configurations based on the placement of ribs at different walls. The baseline geometries are two-pass trapezoidal channels with and without orthogonal ribs at the bottom surface of the channel. Ribs induce secondary flow which results in enhancement of heat transfer; therefore, for enhancement of heat transfer at the trailing edge, ribs are placed at the trailing edge surface in three different configurations: first without ribs at the bottom surface, then ribs at the trailing edge surface in-line with the ribs at the bottom surface, and finally staggered ribs. Heat transfer and pressure drop is calculated at Reynolds number equal to 9400 for all configurations. Different turbulent models are used for the validation of the numerical results. For the smooth channel low-Re k-ɛ model, realizable k-ɛ model, the RNG k-ω model, low-Re k-ω model, and SST k-ω models are compared, whereas for ribbed channel, low-Re k-ɛ model and SST k-ω models are compared. The results show that the low-Re k-ɛ model, which predicts the heat transfer in outlet pass of the smooth channels with difference of +7%, underpredicts the heat transfer by -17% in case of ribbed channel compared to

  18. Validation and Analysis of Numerical Results for a Two-Pass Trapezoidal Channel With Different Cooling Configurations of Trailing Edge.

    PubMed

    Siddique, Waseem; El-Gabry, Lamyaa; Shevchuk, Igor V; Fransson, Torsten H

    2013-01-01

    High inlet temperatures in a gas turbine lead to an increase in the thermal efficiency of the gas turbine. This results in the requirement of cooling of gas turbine blades/vanes. Internal cooling of the gas turbine blade/vanes with the help of two-pass channels is one of the effective methods to reduce the metal temperatures. In particular, the trailing edge of a turbine vane is a critical area, where effective cooling is required. The trailing edge can be modeled as a trapezoidal channel. This paper describes the numerical validation of the heat transfer and pressure drop in a trapezoidal channel with and without orthogonal ribs at the bottom surface. A new concept of ribbed trailing edge has been introduced in this paper which presents a numerical study of several trailing edge cooling configurations based on the placement of ribs at different walls. The baseline geometries are two-pass trapezoidal channels with and without orthogonal ribs at the bottom surface of the channel. Ribs induce secondary flow which results in enhancement of heat transfer; therefore, for enhancement of heat transfer at the trailing edge, ribs are placed at the trailing edge surface in three different configurations: first without ribs at the bottom surface, then ribs at the trailing edge surface in-line with the ribs at the bottom surface, and finally staggered ribs. Heat transfer and pressure drop is calculated at Reynolds number equal to 9400 for all configurations. Different turbulent models are used for the validation of the numerical results. For the smooth channel low-Re k-ɛ model, realizable k-ɛ model, the RNG k-ω model, low-Re k-ω model, and SST k-ω models are compared, whereas for ribbed channel, low-Re k-ɛ model and SST k-ω models are compared. The results show that the low-Re k-ɛ model, which predicts the heat transfer in outlet pass of the smooth channels with difference of +7%, underpredicts the heat transfer by -17% in case of ribbed channel compared to

  19. Leading- and trailing-edge effects on the aeromechanics of membrane aerofoils

    NASA Astrophysics Data System (ADS)

    Arbós-Torrent, Sara; Ganapathisubramani, Bharathram; Palacios, Rafael

    2013-04-01

    This study explores the effect that geometry of silver steel supports have on the aeromechanic performance of membrane aerofoils. Tests are performed at low Reynolds numbers, Re=9×104, and incidences of 2°-25° High-speed photogrammetry as well as force measurements are carried out to explore the effects of four different leading-edge (LE) and trailing-edge (TE) designs on the performance of membrane aerofoils. Results indicate that the mean camber as well as membrane vibrations (both mode shape and frequency) change with geometry and size of the LE and TE supports. The LE/TE supports with a rectangular cross-section consistently provide higher lift forces and higher mean camber deformations compared to the support with circular cross-section. The membrane vibrations are also found to be higher for aerofoils with LE/TE supports with rectangular cross-section. Moreover, it is shown that the LE/TE supports deflect under aerodynamic loading and consequently alter the performance of the aerofoil. Furthermore, some of the supports are found to vibrate at their resonance frequency. In all, this study quantifies the impact of the leading- and trailing-edge support on the membrane and provides guidelines for geometry selection for future studies.

  20. KRISTINA: Kinematic rib-based structural system for innovative adaptive trailing edge

    NASA Astrophysics Data System (ADS)

    Pecora, R.; Amoroso, F.; Magnifico, M.; Dimino, I.; Concilio, A.

    2016-04-01

    Nature teaches that the flight of the birds succeeds perfectly since they are able to change the shape of their wings in a continuous manner. The careful observation of this phenomenon has re-introduced in the recent research topics the study of "metamorphic" wing structures; these innovative architectures allow for the controlled wing shape adaptation to different flight conditions with the ultimate goal of getting desirable improvements such as the increase of aerodynamic efficiency or load control effectiveness. In this framework, the European research project SARISTU aimed at combining morphing and smart ideas to the leading edge, the trailing edge and the winglet of a large commercial airplane (EASA CS25 category) while assessing integrated technologies validation through high-speed wind tunnel test on a true scale outer wing segment. The design process of the adaptive trailing edge (ATED) addressed by SARISTU is here outlined, from the conceptual definition of the camber-morphing architecture up to the assessment of the device executive layout. Rational design criteria were implemented in order to preliminarily define ATED structural layout and the general configuration of the embedded mechanisms enabling morphing under the action of aerodynamic loads. Advanced FE analyses were then carried out and the robustness of adopted structural arrangements was proven in compliance with applicable airworthiness requirements.

  1. Active interrogation of helicopter main rotor faults using trailing edge flap actuation

    NASA Astrophysics Data System (ADS)

    Stevens, Patricia Lynn

    Over the past decade, the helicopter community has become increasingly interested in health monitoring. The rotor system, however, is not sufficiently covered in the current Health and Usage Monitoring Systems (HUMS). This dissertation describes the development and evaluation of a new approach for detecting helicopter rotor faults in which active trailing edge flaps are used to interrogate the system. This work is based on the presumption that trailing edge flaps would be installed for the primary purpose of vibration and/or noise control; health monitoring is a secondary use. Using this approach, the blade is excited by an interrogation signal, which is a low amplitude oscillation at a few discrete frequencies. The blade response is measured and the health of the system is determined using a frequency domain damage identification algorithm. Damage detection and location are achieved via the residual force vector. The residual force vector, coupled with an understanding of the system physics, also provides nature characterization. Quantification of damage extent is achieved via a frequency domain adaptation of the Asymmetric Minimum Rank Perturbation Theory. The active interrogation system is evaluated using an aeroelastic finite element model of the rotor system in hover, including an advanced unsteady aerodynamic model to predict the trailing edge flap loads. Realistic damage models, including distributed bending stiffness damage, torsional stiffness damage, control system stiffness damage, cracks and ballistic damage, are seeded in the rotor system model. Results demonstrate detection, location and quantification of extent of all of the faults tested. The effects of noise and modeling errors are discussed and mitigation techniques are developed. Additionally, a measurability study is included. Benefits of this work include both improved health monitoring for rotorcraft as well as insights into the application of structural damage detection algorithms to a

  2. Modeling Creep-Induced Stress Relaxation at the Leading Edge of SiC/SiC Airfoils

    NASA Technical Reports Server (NTRS)

    Lang, Jerry; DiCarlo, James A.

    2007-01-01

    Anticipating the implementation of advanced SiC/SiC composites into internally cooled airfoil components within the turbine section of future aero-propulsion engines, the primary objective of this study was to develop physics-based analytical and finite-element modeling tools to predict the effects of composite creep and stress relaxation at the airfoil leading edges, which will generally experience large thermal gradients at high temperatures. A second objective was to examine how some advanced NASA-developed SiC/SiC systems coated with typical EBC materials would behave as leading edge materials in terms of long-term steady-state operating temperatures. Because of the complexities introduced by mechanical stresses inherent in internally cooled airfoils, a simple cylindrical thin-walled tube model subjected to thermal stresses only is employed for the leading edge, thereby obtaining a best-case scenario for the material behavior. In addition, the SiC/SiC composite materials are assumed to behave as isotropic materials with temperature-dependent viscoelastic creep behavior as measured in-plane on thin-walled panels. Key findings include: (1) without mechanical stresses and for typical airfoil geometries, as heat flux is increased through the leading edge, life-limiting tensile crack formation will occur first in the hoop direction on the inside wall of the leading edge; (2) thermal gradients through all current SiC/SiC systems should be kept below approx.300 F at high temperatures to avoid this cracking; (3) at temperatures near the maximum operating temperatures of advanced SiC/SiC systems, thermal stresses induced by the thermal gradients will beneficially relax with time due to creep; (4) although stress relaxation occurs, the maximum gradient should still not exceed 300oF because of residual tensile stress buildup on the airfoil outer wall during cool-down; and (5) without film cooling and mechanical stresses, the NASA-developed N26 SiC/SiC system with thru

  3. 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%.

  4. Approach for Structurally Clearing an Adaptive Compliant Trailing Edge Flap for Flight

    NASA Technical Reports Server (NTRS)

    Miller, Eric J.; Lokos, William A.; Cruz, Josue; Crampton, Glen; Stephens, Craig A.; Kota, Sridhar; Ervin, Gregory; Flick, Pete

    2015-01-01

    The Adaptive Compliant Trailing Edge (ACTE) flap was flown on the NASA Gulfstream GIII test bed at the NASA Armstrong Flight Research Center. This smoothly curving flap replaced the existing Fowler flaps creating a seamless control surface. This compliant structure, developed by FlexSys Inc. in partnership with Air Force Research Laboratory, supported NASA objectives for airframe structural noise reduction, aerodynamic efficiency, and wing weight reduction through gust load alleviation. A thorough structures airworthiness approach was developed to move this project safely to flight.

  5. Structure of unsteady flows at leading- and trailing-edges: Flow visualization and its interpretation

    NASA Technical Reports Server (NTRS)

    Rockwell, D.; Atta, R.; Kramer, L.; Lawson, R.; Lusseyran, D.; Magness, C.; Sohn, D.; Staubli, T.

    1987-01-01

    Unsteady two- and three-dimensional flow structure at leading and trailing edges of bodies can be characterized effectively using recently developed techniques for acquisition and interpretation of flow visualization. The techniques addressed here include: flow image/surface pressure correlations; 3-D reconstruction of flow structure from flow images; and interactive interpretation of flow images with theoretical simulations. These techniques can be employed in conjunction with: visual correlation and ensemble-averaging, both within a given image and between images; recognition of patterns from images; and estimates of velocity eigenfunctions from images.

  6. Blowing Circulation Control on a Seaplane Airfoil

    NASA Astrophysics Data System (ADS)

    Guo, B. D.; Liu, P. Q.; Qu, Q. L.

    2011-09-01

    RANS simulations are presented for blowing circulation control on a seaplane airfoil. Realizable k-epsilon turbulent model and pressure-based coupled algorithm with second-order discretization were adopted to simulate the compressible flow. Both clear and simple flap configuration were simulated with blowing momentum coefficient Cμ = 0, 0.15 and 0.30. The results show that blowing near the airfoil trailing edge could enhance the Coanda effect, delay the flow separation, and increase the lift coefficient dramatically. The blowing circulation control is promising to apply to taking off and landing of an amphibious aircraft or seaplane.

  7. 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.

  8. 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.

  9. Near-wall serpentine cooled turbine airfoil

    DOEpatents

    Lee, Ching-Pang

    2013-09-17

    A serpentine coolant flow path (54A-54G) formed by inner walls (50, 52) in a cavity (49) between pressure and suction side walls (22, 24) of a turbine airfoil (20A). A coolant flow (58) enters (56) an end of the airfoil, flows into a span-wise channel (54A), then flows forward (54B) over the inner surface of the pressure side wall, then turns behind the leading edge (26), and flows back along a forward part of the suction side wall, then follows a loop (54E) forward and back around an inner wall (52), then flows along an intermediate part of the suction side wall, then flows into an aft channel (54G) between the pressure and suction side walls, then exits the trailing edge (28). This provides cooling matched to the heating topography of the airfoil, minimizes differential thermal expansion, revives the coolant, and minimizes the flow volume needed.

  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. 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

  12. Distributed electromechanical actuation system design for a morphing trailing edge wing

    NASA Astrophysics Data System (ADS)

    Dimino, I.; Diodati, G.; Concilio, A.; Volovick, A.; Zivan, L.

    2016-04-01

    Next-generation flight control actuation technology will be based on "more electric" concepts to ensure benefits in terms of efficiency, weight and maintenance. This paper is concerned with the design of an un-shafted distributed servo-electromechanical actuation system, suited for morphing trailing edge wings of large commercial aircraft. It aims at producing small wing camber variations in the range between -5° and +5° in cruise, to enable aerodynamic efficiency improvements. The deployment kinematics is based on multiple "direct-drive" actuation, each made of light-weight compact lever mechanisms, rigidly connected to compliant ribs and sustained by load-bearing motors. Navier-Stokes computations are performed to estimate the pressure distribution over the interested wing region and the resulting hinge moments. These transfer to the primary structure via the driving mechanism. An electro-mechanical Matlab/Simulink model of the distributed actuation architecture is developed and used as a design tool, to preliminary evaluate the complete system performance. Implementing a multi-shaft strategy, each actuator is sized for the torque acting on the respective adaptive rib, following the effect of both the aerodynamic pressure and the morphing skin stiffness. Elastic trailing edge rotations and power needs are evaluated in operative conditions. Focus is finally given to the key challenges of the proposed concept: targeting quantifiable performance improvements while being compliant to the demanding requirements in terms of reliability and safety.

  13. Flow Phenomena in the Very Near Wake of a Flat Plate with a Circular Trailing Edge

    NASA Technical Reports Server (NTRS)

    Rai, Man Mohan

    2014-01-01

    The very near wake of a flat plate with a circular trailing edge, exhibiting pronounced shedding of wake vortices, is investigated with data from a direct numerical simulation. The separating boundary layers are turbulent and statistically identical thus resulting in a wake that is symmetric in the mean. The focus here is on the instability of the detached shear layers, the evolution of rib-vortex induced localized regions of reverse flow that detach from the main body of reverse flow in the trailing edge region and convect downstream, and phaseaveraged velocity statistics in the very near wake. The detached shear layers are found to exhibit unstable behavior intermittently, including the development of shear layer vortices as in earlier cylinder flow investigations with laminar separating boundary layers. Only a small fraction of the separated turbulent boundary layers undergo this instability, and form the initial shed vortices. Pressure spectra within the shear layers show a broadband peak at a multiple of shedding frequency. Phase-averaged intensity and shear stress distributions of the randomly fluctuating component of velocity are compared with those obtained in the near wake. The distributions of the production terms in the transport equations for the turbulent stresses are also provided.

  14. The structure of separated flow regions occurring near the leading edge of airfoils, including transition

    NASA Technical Reports Server (NTRS)

    Mueller, T. J.

    1985-01-01

    The structure and behavior of the separation bubble including transition and the redeveloping boundary layer after reattachment over an airfoil at low Reynolds numbers was studied. The intent is to further the understanding of the complex flow phenomena 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.

  15. 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.

  16. Flow conditions around the exit and downstream of certain stator blading with various trailing edge thicknesses and geometries

    NASA Technical Reports Server (NTRS)

    Prust, H. W., Jr.; Helon, R. M.

    1972-01-01

    The blading investigated was curved back with thick profiles. The variations in flow conditions considered were flow angle and isentropic energy. Experimental data were obtained in a two-dimensional cascade from surveys with a combined angle, total-, and static-pressure probe and from an array of end-wall static pressure taps. Analytical results were obtained from ideal flow theory. The results showed large variations in flow conditions close to the plane of the trailing edge that were largely attenuated at a distance a little greater than one blade pitch downstream of the trailing edge in the direction of flow. The results were affected by the geometry and thickness of the trailing edge. The agreement between experimental and analytical results is generally fair to excellent.

  17. High-flaps for natural laminar flow airfoils

    NASA Technical Reports Server (NTRS)

    Morgan, Harry L.

    1986-01-01

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

  18. Study on Busemann Biplane Airfoil in Low-Speed Smoke Wind Tunnel

    NASA Astrophysics Data System (ADS)

    Kashitani, Masashi; Yamaguchi, Yutaka; Kai, Yoshiharu; Hirata, Kenichi; Kusunose, Kazuhiro

    The Busemann biplane airfoil is considered one of the candidates for reducing sonic boom. In aircraft designs utilizing the biplane concept, high-lift devices must be used for takeoff and landing in low-speed conditions. In this work, flow visualizations were performed around a Busemann biplane airfoil equipped with leading and trailing edge flaps in a smoke wind tunnel. The lift coefficient of the biplane airfoil was estimated by utilizing a method based on measurements of smoke line patterns. The aspect ratio of the baseline Busemann biplane model was 0.75, the thickness ratio of the single element was 5%, and the wave cancellation condition was designed for Mach number 1.7. The length of each of the flap chords was 30% of the baseline. The Reynolds number, which is based on the chord length of the airfoil, is about 2.8×105. The results of the study are summarized as follows. For the baseline Busemann airfoil without flaps, the lift coefficient increases linearly as the angle of attack increases. The slope of the lift coefficient cl is 0.062 (1/deg.), which is in good agreement with reference data. This indicates that measuring smoke line patterns is a valid method for estimating the lift coefficient of biplane airfoils. Based on the visualization of the flow around the biplane model equipped with deflected leading and trailing edge flaps, confirmed that the separation bubble is smaller than in the baseline model due to the effective increase in camber. When the deflection angle of the trailing edge flap is increased, the lift coefficient also increases. The trend of the increasing cl is similar to that of conventional monoplane airfoil models with trailing edge flaps. Therefore, such flaps can be considered effective high-lift devices for Busemann biplane airfoils.

  19. Investigation of trailing-edge-flap, spanwise-blowing concepts on an advanced fighter configuration

    NASA Technical Reports Server (NTRS)

    Paulson, J. W., Jr.; Quinto, P. F.; Banks, D. W.

    1984-01-01

    The aerodynamic effects of spanwise blowing on the trailing edge flap of an advanced fighter aircraft configuration were determined in the 4 by 7 Meter Tunnel. A series of tests were conducted with variations in spanwise-blowing vector angle, nozzle exit area, nozzle location, thrust coefficient, and flap deflection in order to determine a superior configuration for both an underwing cascade concept and an overwing port concept. This screening phase of the testing was conducted at a nominal approach angle of attack from 12 deg to 16 deg; and then the superior configurations were tested over a more complete angle of attack range from 0 deg to 20 deg at tunnel free stream dynamic pressures from 20 to 40 lbf/sq ft at thrust coefficients from 0 to 2.

  20. The calculated effect of trailing-edge flaps on the take-off of flying boats

    NASA Technical Reports Server (NTRS)

    Parkinson, J E; Bell, J W

    1934-01-01

    The results of take-off calculations are given for an application of simple trailing-edge flaps to two hypothetical flying boats, one having medium wing and power loading and consequently considerable excess of thrust over total resistance during the take-off run, the other having high wing and power loading and a very low excess thrust. For these seaplanes the effect of downward flap settings was: (1) to increase the total resistance below the stalling speed, (2) to decrease the get-away speed, (3) to improve the take-off performance of the seaplane having considerable excess thrust, and (4) to hinder the take-off of the seaplane having low excess thrust. It is indicated that flaps would allow a decrease in the high angles of wing setting necessary with most seaplanes, provided that the excess thrust is not too low.

  1. An evaluation of several wind turbine trailing-edge aerodynamic brakes

    SciTech Connect

    Miller, L.S.; Migliore, P.G.; Quandt, G.A.

    1995-09-01

    An investigation was undertaken to identify the aerodynamic performance of five separate trailing-edge control devices and to evaluate their potential for wind turbine overspeed control applications. A modular two-dimensional wind tunnel model was constructed and evaluated during extensive wind tunnel testing. Aerodynamic lift, drag and suction coefficient data were acquired and analyzed for various control configurations and angles-of-attack. To further interpret their potential performance, the controls were evaluated numerically using a generic wind turbine geometry and a performance analysis computer program. On the basis of the results of the investigation, the Spoiler-Flap control configuration was deemed best suited for turbine braking applications. This particular control exhibited a good negative suction coefficient behavior over a broad angle-of-attack range and good turbine braking capabilities, especially at low tip-speed ratio conditions.

  2. Approach for Structurally Clearing an Adaptive Compliant Trailing Edge Flap for Flight

    NASA Technical Reports Server (NTRS)

    Miller, Eric J.; Lokos, William A.; Cruz, Josue; Crampton, Glen; Stephens, Craig A.; Kota, Sridhar; Ervin, Gregory; Flick, Pete

    2015-01-01

    The Adaptive Compliant Trailing Edge (ACTE) flap was flown on the National Aeronautics and Space Administration (NASA) Gulfstream GIII testbed at the NASA Armstrong Flight Research Center. This smoothly curving flap replaced the existing Fowler flaps creating a seamless control surface. This compliant structure, developed by FlexSys Inc. in partnership with the Air Force Research Laboratory, supported NASA objectives for airframe structural noise reduction, aerodynamic efficiency, and wing weight reduction through gust load alleviation. A thorough structures airworthiness approach was developed to move this project safely to flight. A combination of industry and NASA standard practice require various structural analyses, ground testing, and health monitoring techniques for showing an airworthy structure. This paper provides an overview of compliant structures design, the structural ground testing leading up to flight, and the flight envelope expansion and monitoring strategy. Flight data will be presented, and lessons learned along the way will be highlighted.

  3. Trailing edge devices to improve performance and increase lifetime of wind-electric water pumping systems

    SciTech Connect

    Vick, B.D.; Clark, R.N.

    1996-12-31

    Trailing edge flaps were applied to the blades of a 10 kW wind turbine used for water pumping to try to improve the performance and decrease the structural fatigue on the wind turbine. Most small wind turbines (10 kW and below) use furling (rotor turns out of wind similar to a mechanical windmill) to protect the wind turbine from overspeed during high winds. Some small wind turbines, however, do not furl soon enough to keep the wind turbine from being off line part of the time in moderately high wind speeds (10 - 16 m/s). As a result, the load is disconnected and no water is pumped at moderately high wind speeds. When the turbine is offline, the frequency increases rapidly often causing excessive vibration of the wind turbine and tower components. The furling wind speed could possibly be decreased by increasing the offset between the tower centerline and the rotor centerline, but would be a major and potentially expensive retrofit. Trailing edge flaps (TEF) were used as a quick inexpensive method to try to reduce the furling wind speed and increase the on time by reducing the rotor RPM. One TEF configuration improved the water pumping performance at moderately high wind speeds, but degraded the pumping performance at low wind speeds which resulted in little change in daily water volume. The other TEF configuration differed very little from the no flap configuration. Both TEF configurations however, reduced the rotor RPM in high wind conditions. The TEF, did not reduce the rotor RPM by lowering the furling wind speed as hoped, but apparently did so by increasing the drag which also reduced the volume of water pumped at the lower wind speeds. 6 refs., 9 figs.

  4. Measurement of aeroacoustic noise generated on wind turbine blades modified by trailing edge brushes

    NASA Astrophysics Data System (ADS)

    Asheim, Michael J.

    As wind technology becomes a larger portion of the energy production picture, the problematic interactions between the machines and society will continue to become more pronounced. Of these problems, wind turbine noise is one of the most important to the future of wind turbine development. This study looks at the effect trailing edge brushes mounted on the 2 bladed Controls Advance Research Turbine (CART 2), located at the National Wind Technology Center, have on the overall acoustic and aerodynamic performance of the blades. The use of trailing edge brushes reduced the aeroacoustic noise by 1.0 to 5.0 dB over the baseline blade, depending on wind speed. This acoustic performance comes at a cost to the aerodynamic performance of the blades. The aerodynamic performance indicators, such as turbine power and root bending moments show that increased drag due to the brushes is the main contributor to the reduction in power production. An economic analysis also investigated how to best use noise mitigation devices to optimize acoustic, power performance and loads of a 600 kW baseline turbine, such as the CART 2. The analysis shows that the use of up a noise mitigation device of 4 dB is best used by increasing the rotor diameter and the power rating of the machine, from a 43.3 m diameter, 600 kW machine to a 68.8 m diameter, 886.7 kW machine. This increase resulted in an annual energy production increase of 414% when using a Rayleigh wind distribution with at a mean annual wind speed of 8.5 m/s. This is a reduction of cost of energy from 0.0463 per kWh to 0.0422 kWh. This reduction in energy production costs helps to explain the continuing trend of turbine machine growth in both rotor diameter and power rating.

  5. 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.

  6. Airfoil noise in a uniform flow

    NASA Astrophysics Data System (ADS)

    Garcia, P.

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

  7. Feedback in separated flows over symmetric airfoils

    NASA Technical Reports Server (NTRS)

    Atassi, H. M.

    1984-01-01

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

  8. SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

    NASA Technical Reports Server (NTRS)

    Amari, S.; Foote, J.; Simon, Charles G.; Swan, P.; Walker, R. M.; Zinner, E.; Jessberger, E. K.; Lange, G.; Stadermann, F.

    1992-01-01

    The Long Duration Exposure Facility (LDEF) Experiment AO187-2 consisted of 237 capture cells, 120 on the leading edge and 117 on the trailing edge. Each cell was made of polished Ge plates covered with 2.5 micron thick mylar foil at 200 microns from the Ge. Although all leading edge cells and 105 trailing edge cells had lost their plastic covers during flight, optical and electron microscope examination revealed extended impacts in bare cells from either edge that apparently were produced by high velocity projectiles while the plastic foils were still in place. Detailed optical scanning yielded 53 extended impacts on 100 bare cells from the trailing edge that were selected for SIMS chemical analysis. Lateral multi-element ion probe profiles were obtained on 40 of these impacts. Material that can be attributed to the incoming projectiles was found in all analyzed extended compact features and most seem to be associated with cosmic dust particles. However, LDEF deposits are systematically enriched in the refractory elements Al, Ca, and Ti relative to Mg and Fe when compared to IDP's collected in the stratosphere and to chondritic compositions. These differences are most likely due to elemental fractionation effects during the high velocity impact but real differences between interplanetary particles captured on LDEF and stratospheric IDP's cannot be excluded. Recently we extended our studies to cells from the leading edge and the covered cells from the trailing edge. The 12 covered cells contain 20 extended impact candidates. Ion probe analysis of 3 yielded results similar to those obtained for impacts on the bare cells from the trailing edge. Optical scanning of the bare leading edge cell also reveals many extended impacts (42 on 22 cells scanned to date), demonstrating that the cover foils remained intact at least for some time. However, SIMS analysis showed elements that can reasonably be attributed to micrometeoroids in only 2 out of 11 impacts. Eight impacts

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

    PubMed

    Fuchiwaki, Masaki; Tanaka, Kazuhiro

    2002-10-01

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

  10. SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

    NASA Technical Reports Server (NTRS)

    Amari, S.; Foote, J.; Swan, P.; Walker, R. M.; Zinner, E.; Lange, G.

    1993-01-01

    Numerous 'extended impacts' found in both leading and trailing edge capture cells were successfully analyzed for the chemical composition of projectile residues by secondary ion mass spectrometry (SIMS). Most data were obtained from the trailing edge cells where 45 of 58 impacts were classified as 'probably natural' and the remainder as 'possibly man-made debris.' This is in striking contrast to leading edge cells where 9 of 11 impacts so far measured are definitely classified as orbital debris. Although all the leading edge cells had lost their plastic entrance foils during flight, the rate of foil failure was similar to that of the trailing edge cells, 10 percent of which were recovered intact. Ultraviolet embrittlement is suspected as the major cause of failure on both leading and trailing edges. The major impediment to the accurate determination of projectile chemistry is the fractionation of volatile and refractory elements in the hypervelocity impact and redeposition processes. This effect had been noted in a simulation experiment but is more pronounced in the LDEF capture cells, probably due to the higher average velocities of the space impacts. Surface contamination of the pure Ge surfaces with a substance rich in Si, but also containing Mg and Al, provides an additional problem for the accurate determination of impactor chemistry. The effect is variable, being much larger on surfaces that were exposed to space than in those cells that remained intact. Future work will concentrate on the analyses of more leading edge impacts and the development of new SIMS techniques for the measurement of elemental abundances in extended impacts.

  11. On bimodal flutter behavior of a flexible airfoil

    NASA Astrophysics Data System (ADS)

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

    2014-02-01

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

  12. Noise Benefits of Rotor Trailing Edge Blowing for a Model Turbofan

    NASA Technical Reports Server (NTRS)

    Woodward, Richard P.; Fite, E. Brian; Podboy, Gary G.

    2007-01-01

    An advanced model turbofan was tested in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects associated with rotor Trailing-Edge-Blowing (TEB) for a modern, 1.294 stage pressure ratio turbofan model. The TEB rotor (Fan9) was designed to be aerodynamically similar to the previously tested Fan1, and used the same stator and nacelle hardware. Fan9 was designed with trailing edge blowing slots using an external air supply directed through the rotor hub. The TEB flow was heated to approximate the average fan exit temperature at each fan test speed. Rotor root blockage inserts were used to block TEB to all but the outer 40 and 20% span in addition to full-span blowing. A configuration with full-span TEB on alternate rotor blades was also tested. Far field acoustic data were taken at takeoff/approach conditions at 0.10 tunnel Mach. Far-field acoustic results showed that full-span blowing near 2.0% of the total flow could reduce the overall sound power level by about 2 dB. This noise reduction was observed in both the rotor-stator interaction tones and for the spectral broadband noise levels. Blowing only the outer span region was not very effective for lowering noise, and actually increased the far field noise level in some instances. Full-span blowing of alternate blades at 1.0% of the overall flow rate (equivalent to full-span blowing of all blades at 2.0% flow) showed a more modest noise decrease relative to full-span blowing of all blades. Detailed hot film measurements of the TEB rotor wake at 2.0% flow showed that TEB was not every effective for filling in the wake defect at approach fan speed toward the tip region, but did result in overfilling the wake toward the hub. Downstream turbulence measurements supported this finding, and support the observed reduction in spectral broadband noise.

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

    NASA Astrophysics Data System (ADS)

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

    2016-09-01

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

  14. Theory of viscous transonic flow over airfoils at high Reynolds number

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

    This paper considers viscous flows with unseparated turbulent boundary layers over two-dimensional airfoils at transonic speeds. Conventional theoretical methods are based on boundary layer formulations which do not account for the effect of the curved wake and static pressure variations across the boundary layer in the trailing edge region. In this investigation an extended viscous theory is developed that accounts for both effects. The theory is based on a rational analysis of the strong turbulent interaction at airfoil trailing edges. The method of matched asymptotic expansions is employed to develop formal series solutions of the full Reynolds equations in the limit of Reynolds numbers tending to infinity. Procedures are developed for combining the local trailing edge solution with numerical methods for solving the full potential flow and boundary layer equations. Theoretical results indicate that conventional boundary layer methods account for only about 50% of the viscous effect on lift, the remaining contribution arising from wake curvature and normal pressure gradient effects.

  15. CFD Analysis of the Aerodynamics of a Business-Jet Airfoil with Leading-Edge Ice Accretion

    NASA Technical Reports Server (NTRS)

    Chi, X.; Zhu, B.; Shih, T. I.-P.; Addy, H. E.; Choo, Y. K.

    2004-01-01

    For rime ice - where the ice buildup has only rough and jagged surfaces but no protruding horns - this study shows two dimensional CFD analysis based on the one-equation Spalart-Almaras (S-A) turbulence model to predict accurately the lift, drag, and pressure coefficients up to near the stall angle. For glaze ice - where the ice buildup has two or more protruding horns near the airfoil's leading edge - CFD predictions were much less satisfactory because of the large separated region produced by the horns even at zero angle of attack. This CFD study, based on the WIND and the Fluent codes, assesses the following turbulence models by comparing predictions with available experimental data: S-A, standard k-epsilon, shear-stress transport, v(exp 2)-f, and differential Reynolds stress.

  16. Icing tunnel tests of a glycol-exuding porous leading edge ice protection system on a general aviation airfoil

    NASA Technical Reports Server (NTRS)

    Kohlman, D. L.; Schweikhard, W. G.; Albright, A. E.; Evanich, P.

    1981-01-01

    A glycol-exuding porous leading edge ice protection system was tested. Results show that the system is very effective in preventing ice accretion (anti-ice mode) or removing ice from an airfoil. Minimum glycol flow rates required for anti-icing are a function of velocity, liquid water content in the air, ambient temperature, and droplet size. Large ice caps were removed in only a few minutes using anti-ice flow rates. It was found that the shed time is a function of the type of ice, size of the ice cap, angle of attack, and glycol flow rate. Wake survey measurements show that there is no significant drag penalty for the installation or operation of the system tested.

  17. On experimental sensitivity analysis of the turbulent wake from an axisymmetric blunt trailing edge

    NASA Astrophysics Data System (ADS)

    Grandemange, M.; Gohlke, M.; Parezanović, V.; Cadot, O.

    2012-03-01

    The sensitivity to steady local disturbances of the turbulent wake over a 3D blunt body with an axisymmetric trailing edge is investigated at Re = 2.1 × 104. The flow presents a favored m = 2 azimuthal periodicity due to the mounting of the body. It is proved that two asymmetric flows cohabit and generate a statistically symmetric natural flow. Topology shifts are random but are visible after a large number of global mode periods. The statistical symmetry is highly sensitive to any steady antisymmetric disturbance. As a consequence, depending on its position, a small control cylinder (considered as a disturbance with m = 1 azimuthal periodicity) in the close wake fixes the flow to one asymmetric topology affecting shedding activity as well as drag. On the contrary, a thin control ring (i.e., m = 0 disturbance) preserves the mean flow symmetry and has a greater impact on global mode frequency and drag. Whatever the disturbance, the sensitivity of the wake to control devices is concentrated into the mixing layers; the observed effects seem to depend more upon their local turbulent characteristics than on the global inviscid dynamics of vorticity.

  18. Characterization of base pressure fluctuations in a blunt trailing edge wake with three-dimensional forcing

    NASA Astrophysics Data System (ADS)

    Clark, Heather; Lavoie, Philippe

    2015-11-01

    The wakes of many nominally two-dimensional bluff bodies exhibit multiple intrinsic three-dimensional instabilities whose spatiotemporal structure and growth rate depend on geometry and Reynolds number. Here, these features are investigated experimentally for a blunt trailing edge profiled body using simultaneous measurements of velocity and fluctuating surface pressure on the model rear face near separation. Passive three-dimensional forcing of the wake is implemented with an array of vortex generators that are distributed according to the characteristic spanwise wavelength of the dominant secondary instability. For a Reynolds number of 8000 based on model thickness, the control strategy is found to increase the base pressure coefficient by 26% while globally reducing the amplitude of base pressure fluctuations, relative to the unforced flow. Additionally, amplitude modulation of the pressure signals that is observed in the natural wake decreases in strength with distributed forcing as a result of the modified three-dimensional flow structure. The spanwise distribution of pressure will be further examined for the baseline and controlled flows via temporal spectral analysis and spatial modal decomposition.

  19. Near Wake Flow Topology of a Blunt Trailing Edge Profiled Flat Plate

    NASA Astrophysics Data System (ADS)

    Sampat Doddipatla, Lakshmana; Naghib Lahouti, Arash; Hangan, Horia; Siddiqui, Kamran

    2009-11-01

    Wake flows behind two dimensional bodies are unstable due to formation of spanwise von Karman vortices accompanied by three dimensional streamwise instabilities, also referred to as rolls and ribs, respectively. These three dimensional instabilities lead to two distinct instability modes (Mode A and Mode B), or a combination of the two, depending on the flow Reynolds number and the profile geometry. It has been observed that the ribs wrap around the rolls, progressively distorting them. Therefore, enhancing the action of streamwise vortices can lead to early suppression of the spanwise von Karman vortices accompanied by the reduction of fluctuating lift and base drag. The present investigation seeks to better understand these near wake instabilities for blunt trailing edge profiled bodies of various aspect ratios, for flow Reynolds numbers ranging from Re(d)=500 to Re(d)=2200, and various inlet conditions. Planar Laser Induced Fluorescence (PLIF) visualizations and measurements are performed in the near wake to study and characterize the topology of streamwise and spanwise vortices.

  20. Experimental Study of Wake Instabilities of a Blunt Trailing Edge Profiled Body at Intermediate Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Naghib-Lahouti, Arash; Lavoie, Philippe; Hangan, Horia

    2011-11-01

    The periodic shedding of von Kármán vortices is the primary instability in the wake of nominally 2D bluff bodies, beyond a critical Reynolds number around 45-49. When Reynolds number passes a second threshold, which can be as high as 700 depending on profile geometry, secondary instabilities emerge and accompany the von Kármán vortices. For most bluff bodies, these instabilities appear as pairs of counter-rotating streamwise vortices, and spanwise undulations of the von Kármán vortices. The mechanism and scale of these instabilities depend on the bluff body geometry and Reynolds number. The focus of the present study is to identify and characterize the dominant secondary instability in the wake of a blunt trailing edge profiled body at intermediate Reynolds numbers between 8,000 and 20,000 based on the body thickness. The experiments, which include PIV and hot-wire measurements in the wake, complement previous studies involving the same bluff body at higher and lower Reynolds numbers, and make it possible to determine the scale and mechanism of the secondary instability at intermediate Reynolds numbers. Funded partly by the Government of Ontario, MITACS and Bombardier Aerospace.

  1. Predicting Modes of the Unsteady Vorticity Field near the Trailing Edge of a Blade

    NASA Technical Reports Server (NTRS)

    Devenport, William J.; Spitz, Nicolas; Envia, Edmane (Technical Monitor)

    2003-01-01

    Progress on predicting modes of the unsteady velocity/vorticity field of a turbulent boundary layer from Reynolds stress statistics is described. Prediction of these modes, that provide the source terms for trailing edge noise predictions in aircraft engine fans and other configurations, will allow for the first time detailed viscous flow effects to be included in such noise calculations. The key accomplishments of this work in FY02 are: (1) The development of a Matlab code for the prediction of modes in two- and three-dimensional boundary layers, previously applied to plane wakes; (2) Predictions with the code using a constant lengthscale formulation in a fully developed turbulence channel flow. Comparison of these boundary layer predictions with available DNS simulation results; and (3) Formulation of an improved model using a variable lengthscale proportional to mixing length. Turbulent channel flow predictions and comparison with DNS results. This work is being carried out in continuous communication and collaboration with the Glegg research group at Florida Atlantic University, which will be incorporating mode predictions into engine noise calculations.

  2. Dynamics of a pneumatic artificial muscle actuation system driving a trailing edge flap

    NASA Astrophysics Data System (ADS)

    Woods, Benjamin K. S.; Kothera, Curt S.; Wang, Gang; Wereley, Norman M.

    2014-09-01

    This study presents a time domain dynamic model of an antagonistic pneumatic artificial muscle (PAM) driven trailing edge flap (TEF) system for next generation active helicopter rotors. Active rotor concepts are currently being widely researched in the rotorcraft community as a means to provide a significant leap forward in performance through primary aircraft control, vibration mitigation and noise reduction. Recent work has shown PAMs to be a promising candidate for active rotor actuation due to their combination of high force, large stroke, light weight, and suitable bandwidth. When arranged into biologically inspired agonist/antagonist muscle pairs they can produce bidirectional torques for effectively driving a TEF. However, there are no analytical dynamic models in the literature that can accurately capture the behavior of such systems across the broad range of frequencies required for this demanding application. This work combines mechanical, pneumatic, and aerodynamic component models into a global flap system model developed for the Bell 407 rotor system. This model can accurately predict pressure, force, and flap angle response to pneumatic control valve inputs over a range of operating frequencies from 7 to 35 Hz (1/rev to 5/rev for the Bell 407) and operating pressures from 30 to 90 psi.

  3. Performance of a short annular dump diffuser using wall trailing-edge suction

    NASA Technical Reports Server (NTRS)

    Juhasz, A. J.

    1974-01-01

    A short annular dump (abrupt flow area change) diffuser was tested with suction through wall trailing-edge slots at inlet Mach numbers of 0.19 and 0.27 and at near ambient inlet temperature and pressure, with suction flow varied from zero to 10 percent of the inlet air mass-flow rate. The overall ratio of diffuser exit area to inlet area was 4.0, and the ratio of length to inlet height was 2.0. By applying suction flow separately on either wall or to both walls simultaneously, the original annular jet profile could be altered to either a hub- or tip-biased profile. Diffuser effectiveness was increased from about 25 percent with no suction to 50 percent at 6 percent outer-wall suction and to 52 percent at a combined suction rate on both walls of 10.25 percent. At the same time, diffuser total pressure loss was reduced by one-fourth.

  4. Evaluation of Load Analysis Methods for NASAs GIII Adaptive Compliant Trailing Edge Project

    NASA Technical Reports Server (NTRS)

    Cruz, Josue; Miller, Eric J.

    2016-01-01

    The Air Force Research Laboratory (AFRL), NASA Armstrong Flight Research Center (AFRC), and FlexSys Inc. (Ann Arbor, Michigan) have collaborated to flight test the Adaptive Compliant Trailing Edge (ACTE) flaps. These flaps were installed on a Gulfstream Aerospace Corporation (GAC) GIII aircraft and tested at AFRC at various deflection angles over a range of flight conditions. External aerodynamic and inertial load analyses were conducted with the intention to ensure that the change in wing loads due to the deployed ACTE flap did not overload the existing baseline GIII wing box structure. The objective of this paper was to substantiate the analysis tools used for predicting wing loads at AFRC. Computational fluid dynamics (CFD) models and distributed mass inertial models were developed for predicting the loads on the wing. The analysis tools included TRANAIR (full potential) and CMARC (panel) models. Aerodynamic pressure data from the analysis codes were validated against static pressure port data collected in-flight. Combined results from the CFD predictions and the inertial load analysis were used to predict the normal force, bending moment, and torque loads on the wing. Wing loads obtained from calibrated strain gages installed on the wing were used for substantiation of the load prediction tools. The load predictions exhibited good agreement compared to the flight load results obtained from calibrated strain gage measurements.

  5. The role of shock induced trailing-edge separation in limit cycle oscillations

    NASA Technical Reports Server (NTRS)

    Cunningham, Atlee M., Jr.

    1989-01-01

    The potential role of shock induced trailing edge separation (SITES) in limit cycle oscillations (LCO) was established. It was shown that the flip-flop characteristics of transition to and from SITES as well as its hysteresis could couple with wing modes with torsional motion and low damping. This connection led to the formulation of a very simple nonlinear math model using the linear equations of motion with a nonlinear step forcing function with hysteresis. A finite difference solution with time was developed and calculations were made for the F-111 TACT were used to determine the step forcing function due to SITES transition. Since no data were available for the hysteresis, a parameter study was conducted allowing the hysteresis effect to vary. Very small hysteresis effects, which were within expected bounds, were required to obtain reasonable response levels that essentially agreed with flight test results. Also in agreement with wind tunnel tests, LCO calculations for the 1/6 scale F-111 model showed that the model should have not experienced LCO.

  6. Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade. Part 2:; Simulation Results

    NASA Technical Reports Server (NTRS)

    Rigby, David L.; Bunker, Ronald S.

    2002-01-01

    A combined experimental and numerical study to investigate the heat transfer distribution in a complex blade trailing edge passage was conducted. The geometry consists of a two pass serpentine passage with taper toward the trailing edge, as well as from hub to tip. The upflow channel has an average aspect ratio of roughly 14:1, while the exit passage aspect ratio is about 5:1. The upflow channel is split in an interrupted way and is smooth on the trailing edge side of the split and turbulated on the other side. A turning vane is placed near the tip of the upflow channel. Reynolds numbers in the range of 31,000 to 61,000, based on inlet conditions, were simulated numerically. The simulation was performed using the Glenn-HT code, a full three-dimensional Navier-Stokes solver using the Wilcox k-omega turbulence model. A structured multi-block grid is used with approximately 4.5 million cells and average y+ values on the order of unity. Pressure and heat transfer distributions are presented with comparison to the experimental data. While there are some regions with discrepancies, in general the agreement is very good for both pressure and heat transfer.

  7. A Study of Sound Generation by an Airfoil via Direct Numerical Simulation

    NASA Astrophysics Data System (ADS)

    Tam, Christopher; Ju, Hongbin

    2004-11-01

    Direct numerical simulations of sound generation associated with flow past a NACA0012 airfoil with a blunt trailing edge are performed. The simulations use the high resolution Dispersion-Relation-Preserving (DRP) time marching scheme and a body fitted grid constructed by conformal mapping. It is known experimentally that at Mach number around 0.1 and chord Reynolds number around 200,000 a strong tone is emitted. The primary objective of this study is to investigate the tone generation mechanism. We are also interested to determine the tone frequency and directivity. In the present numerical simulation, vortex shedding at the blunt trailing edge of the airfoil and the emission of a strong tone are observed. The tone frequencies over a range of Reynolds numbers measured in the simulations are in good agreement with experimental measurements. The directivity of the tones resembles closely to that of an oscillating dipole placed with the dipole axis perpendicular to the flow. An examination of the simulation data reveals that vortices are shed at the blunt trailing of the airfoil. However, after a very careful investigation of the wake flow it is believed that the vortex shedding process is not directly responsible for sound generation. The wake flow is highly unstable. A Kelvin-Helmholtz instability wave is excited slightly downstream of the airfoil trailing edge. The processes that generate the instability wave appear to generate the tone as a by-product.

  8. Navier-Stokes computations for circulation control airfoils

    NASA Technical Reports Server (NTRS)

    Pulliam, Thomas H.; Jespersen, Dennis C.; Barth, Timothy J.

    1987-01-01

    Navier-Stokes computations of subsonic to transonic flow past airfoils with augmented lift due to rearward jet blowing over a curved trailing edge are presented. The approach uses a spiral grid topology. Solutions are obtained using a Navier-Stokes code which employs an implicit finite difference method, an algebraic turbulence model, and developments which improve stability, convergence, and accuracy. Results are compared against experiments for no jet blowing and moderate jet pressures and demonstrate the capability to compute these complicated flows.

  9. Navier-Stokes computations for circulation controlled airfoils

    NASA Technical Reports Server (NTRS)

    Pulliam, T. H.; Jesperen, D. C.; Barth, T. J.

    1986-01-01

    Navier-Stokes computations of subsonic to transonic flow past airfoils with augmented lift due to rearward jet blowing over a curved trailing edge are presented. The approach uses a spiral grid topology. Solutions are obtained using a Navier-Stokes code which employs an implicit finite difference method, an algebraic turbulence model, and developments which improve stability, convergence, and accuracy. Results are compared against experiments for no jet blowing and moderate jet pressures and demonstrate the capability to compute these complicated flows.

  10. Aerodynamic Flow Control of a Maneuvering Airfoil

    NASA Astrophysics Data System (ADS)

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

    2010-11-01

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

  11. 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.

  12. Gross separation approaching a blunt trailing edge as the turbulence intensity increases.

    PubMed

    Scheichl, B

    2014-07-28

    A novel rational description of incompressible two-dimensional time-mean turbulent boundary layer (BL) flow separating from a bluff body at an arbitrarily large globally formed Reynolds number, Re, is devised. Partly in contrast to and partly complementing previous approaches, it predicts a pronounced delay of massive separation as the turbulence intensity level increases. This is bounded from above by a weakly decaying Re-dependent gauge function (hence, the BL approximation stays intact locally), and thus the finite intensity level characterizing fully developed turbulence. However, it by far exceeds the moderate level found in a preceding study which copes with the associated moderate delay of separation. Thus, the present analysis bridges this self-consistent and another forerunner theory, proposing extremely retarded separation by anticipating a fully attached external potential flow. Specifically, it is shown upon formulation of a respective distinguished limit at which rate the separation point and the attached-flow trailing edge collapse as [Formula: see text] and how on a short streamwise scale the typical small velocity deficit in the core region of the incident BL evolves to a large one. Hence, at its base, the separating velocity profile varies generically with the one-third power of the wall distance, and the classical triple-deck problem describing local viscous-inviscid interaction crucial for moderately retarded separation is superseded by a Rayleigh problem, governing separation of that core layer. Its targeted solution proves vital for understanding the separation process more close to the wall. Most importantly, the analysis does not resort to any specific turbulence closure. A first comparison with the available experimentally found positions of separation for the canonical flow past a circular cylinder is encouraging. PMID:24936020

  13. Symbolic regression modeling of noise generation at porous airfoils

    NASA Astrophysics Data System (ADS)

    Sarradj, Ennes; Geyer, Thomas

    2014-07-01

    Based on data sets from previous experimental studies, the tool of symbolic regression is applied to find empirical models that describe the noise generation at porous airfoils. Both the self noise from the interaction of a turbulent boundary layer with the trailing edge of an porous airfoil and the noise generated at the leading edge due to turbulent inflow are considered. Following a dimensional analysis, models are built for trailing edge noise and leading edge noise in terms of four and six dimensionless quantities, respectively. Models of different accuracy and complexity are proposed and discussed. For the trailing edge noise case, a general dependency of the sound power on the fifth power of the flow velocity was found and the frequency spectrum is controlled by the flow resistivity of the porous material. Leading edge noise power is proportional to the square of the turbulence intensity and shows a dependency on the fifth to sixth power of the flow velocity, while the spectrum is governed by the flow resistivity and the integral length scale of the incoming turbulence.

  14. Airfoil for a gas turbine engine

    SciTech Connect

    Liang, George

    2011-05-24

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

  15. A 3-D Computational Study of a Variable Camber Continuous Trailing Edge Flap (VCCTEF) Spanwise Segment

    NASA Technical Reports Server (NTRS)

    Kaul, Upender K.; Nguyen, Nhan T.

    2015-01-01

    Results of a computational study carried out to explore the effects of various elastomer configurations joining spanwise contiguous Variable Camber Continuous Trailing Edge Flap (VCCTEF) segments are reported here. This research is carried out as a proof-of-concept study that will seek to push the flight envelope in cruise with drag optimization as the objective. The cruise conditions can be well off design such as caused by environmental conditions, maneuvering, etc. To handle these off-design conditions, flap deflection is used so when the flap is deflected in a given direction, the aircraft angle of attack changes accordingly to maintain a given lift. The angle of attack is also a design parameter along with the flap deflection. In a previous 2D study,1 the effect of camber was investigated and the results revealed some insight into the relative merit of various camber settings of the VCCTEF. The present state of the art has not advanced sufficiently to do a full 3-D viscous analysis of the whole NASA Generic Transport Model (GTM) wing with VCCTEF deployed with elastomers. Therefore, this study seeks to explore the local effects of three contiguous flap segments on lift and drag of a model devised here to determine possible trades among various flap deflections to achieve desired lift and drag results. Although this approach is an approximation, it provides new insights into the "local" effects of the relative deflections of the contiguous spanwise flap systems and various elastomer segment configurations. The present study is a natural extension of the 2-D study to assess these local 3-D effects. Design cruise condition at 36,000 feet at free stream Mach number of 0.797 and a mean aerodynamic chord (MAC) based Reynolds number of 30.734x10(exp 6) is simulated for an angle of attack (AoA) range of 0 to 6 deg. In the previous 2-D study, the calculations revealed that the parabolic arc camber (1x2x3) and circular arc camber (VCCTEF222) offered the best L

  16. 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.

  17. A general aerodynamic approach to the problem of decaying or growing vibrations of thin, flexible wings with supersonic leading and trailing edges and no side edges

    NASA Technical Reports Server (NTRS)

    Warner, R. W.

    1975-01-01

    Indicial aerodynamic influence coefficients were evaluated from potential theory for a thin, flexible wing with supersonic leading and trailing edges only. The analysis is based on the use of small surface areas in which the downwash is assumed uniform. Within this limitation, the results are exact except for the restriction of linearized theory. The areas are not restricted either to square boxes or Mach boxes. A given area may be any rectangle or square which may or may not be cut by the Mach forecone, and any area can be used anywhere in the forecone without loss of accuracy.

  18. Distributed forcing flow control in the wake of a blunt trailing edge profiled body using plasma actuators

    NASA Astrophysics Data System (ADS)

    Naghib-Lahouti, A.; Hangan, H.; Lavoie, P.

    2015-03-01

    A modern flow control technique for reducing the drag associated with the periodic shedding of von Kármán vortices in the wake of a blunt trailing edge profiled body is presented. The technique involves distributed forcing of the wake flow using an array of dielectric barrier discharge plasma actuators, with a spanwise spacing matched to the spanwise wavelength of the dominant secondary wake instability. The experiments include measurement of the velocity field in multiple vertical and horizontal planes in the wake using particle image velocimetry, as well as base pressure, at Reynolds numbers of 2000, 3000, and 5000 based on trailing edge thickness. The flow control technique causes elongation of the vortex formation region across the span, and significant reduction of the fluctuating and total drag forces, up to a maximum of 94% and 18%, respectively. The effectiveness of the flow control technique is shown to be dependent on the induced momentum coefficient. Proper orthogonal decomposition analysis is used to investigate the mechanism of interaction of the flow control technique with the wake flow. Two distinct flow regimes are observed depending on the induced momentum coefficient. The effect of the control on the wake flow structure in the first regime is similar to those observed in previous studies involving mild spanwise-periodic geometric perturbations at the trailing edge, where control leads to streamwise displacement of the vortices and a shift in shedding frequency. However, an incremental increase in the momentum coefficient leads to a second flow regime similar to those previously observed in the case of large-amplitude geometric perturbations, with an almost complete attenuation of vortex shedding in the near-wake region.

  19. Shapes for rotating airfoils

    NASA Technical Reports Server (NTRS)

    Bingham, G. J. (Inventor)

    1984-01-01

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

  20. Preparation of eutectic superalloys by EFG. [Edge-defined Film-fed Growth for directional solidification in airfoil structures

    NASA Technical Reports Server (NTRS)

    Hurley, G. F.; Marr, N. W.

    1975-01-01

    An attempt was made to produce airfoil shaped bars of three different eutectic superalloys by means of the edge-defined, film-fed growth (EFG) method. The alloys used were a gamma + delta Ni-Cb alloy, a gamma/gamma prime + delta Ni-Cb-Al alloy and a Co-TaC alloy containing Ni and Cr. The development of a new die material was essential in the investigation since these alloys are reactive toward known die materials. Tantalum carbide was selected as a die material because it exhibited spontaneous capillary rise and slow rate of degradation in the liquid metals. Eutectic bars up to 1 mm thick and 6 mm wide were grown from TaC dies in order to determine the growth characteristics and the thermal gradient. Large bars of the gamma/gamma prime + delta alloy were grown and tensile tested. A die with a blind central cavity was designed and several hollow, tear-shaped bars were grown.

  1. Swashplateless Helicopter Experimental Investigation: Primary Control with Trailing Edge Flaps Actuated with Piezobenders

    NASA Astrophysics Data System (ADS)

    Copp, Peter

    Helicopter rotor primary control is conventionally carried out using a swashplate with pitch links. Eliminating the swashplate promises to reduce the helicopter's parasitic power in high speed forward flight, as well as may lead to a hydraulic-less vehicle. A Mach-scale swashplateless rotor is designed with integrated piezobender-actuated trailing edge flaps and systematically tested on the benchtop, in the vacuum chamber and on the hoverstand. The blade is nominally based on the UH-60 rotor with a hover tip Mach number of 0.64. The blade diameter is 66 inches requiring 2400 RPM for Mach scale simulation. The rotor hub is modified to reduce the blade fundamental torsional frequency to less than 2.0/rev by replacing the rigid pitch links with linear springs, which results in an increase of the blade pitching response to the trailing edge flaps. Piezoelectric multilayer benders provide the necessary bandwidth, stroke and stiffness to drive the flaps for primary control while fitting inside the blade profile and withstanding the high centrifugal forces. This work focuses on several key issues. A piezobender designed from a soft piezoelectric material, PZT-5K4, is constructed. The new material is used to construct multi-layer benders with increased stroke for the same stiffness relative to hard materials such as PZT-5H2. Each layer has a thickness of 10 mils. The soft material with gold electrodes requires a different bonding method than hard material with nickel electrodes. With this new bonding method, the measured stiffness matches precisely the predicted stiffness for a 12 layer bender with 1.26 inch length and 1.0 inch width with a stiffness of 1.04 lb/mil. The final in-blade bender has a length of 1.38 inches and 1.0 inch width with a stiffness of 0.325 lb/mil and stroke of 20.2 mils for an energy output of 66.3 lb-mil. The behavior of piezobenders under very high electric fields is investigated. High field means +18.9 kV/cm (limited by arcing in air) and -3.54k

  2. Electrical breakdown of a dielectric on the voltage pulse trailing edge: Investigation in terms of the incubation time concept

    NASA Astrophysics Data System (ADS)

    Petrov, Yu. V.; Morozov, V. A.; Smirnov, I. V.; Lukin, A. A.

    2015-12-01

    The time effects of electrical breakdown at the leading and trailing edges of a voltage pulse applied to an interelectrode gap are studied. The pulsed dielectric strengths of limestone, sandstone, clay, paraffin, and water are determined experimentally at fixed parameters of the voltage pulse and different lengths of the dielectric-filled interelectrode gap. Experimental data are explained in terms of a structure-time approach based on the incubation time criterion. It is found that breakdown occurs when a sufficient power impulse (electric energy) arises within a characteristic time rather than when the electric field reaches a limit.

  3. The formation mechanism and impact of streamwise vortices on NACA 0021 airfoil's performance with undulating leading edge modification

    NASA Astrophysics Data System (ADS)

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

    2014-10-01

    Wings with tubercles have been shown to display advantageous loading behavior at high attack angles compared to their unmodified counterparts. In an earlier study by the authors, it was shown that an undulating leading-edge configuration, including but not limited to a tubercled model, induces a cyclic variation in circulation along the span that gives rise to the formation of counter-rotating streamwise vortices. While the aerodynamic benefits of full-span tubercled wings have been associated with the presence of such vortices, their formation mechanism and influence on wing performance are still in question. In the present work, experimental and numerical tests were conducted to further investigate the effect of tubercles on the flow structure over full-span modified wings based on the NACA 0021 profile, in the transitional flow regime. It is found that a skew-induced mechanism accounts for the formation of streamwise vortices whose development is accompanied by flow separation in delta-shaped regions near the trailing edge. The presence of vortices is detrimental to the performance of full-span wings pre-stall, however renders benefits post-stall as demonstrated by wind tunnel pressure measurement tests. Finally, primary and secondary vortices are identified post-stall that produce an enhanced momentum transfer effect that reduces flow separation, thus increasing the generated amount of lift.

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

    NASA Astrophysics Data System (ADS)

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

    2014-12-01

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

  5. 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

  6. 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.

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

    NASA Technical Reports Server (NTRS)

    Lawing, P. L.

    1985-01-01

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

  8. Computer programs for smoothing and scaling airfoil coordinates

    NASA Technical Reports Server (NTRS)

    Morgan, H. L., Jr.

    1983-01-01

    Detailed descriptions are given of the theoretical methods and associated computer codes of a program to smooth and a program to scale arbitrary airfoil coordinates. The smoothing program utilizes both least-squares polynomial and least-squares cubic spline techniques to smooth interatively the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. A technique for computing the camber and thickness distribution of the smoothed airfoil is also discussed. The scaling program can then be used to scale the thickness distribution generated by the smoothing program to a specific maximum thickness which is then combined with the camber distribution to obtain the final scaled airfoil contour. Computer listings of the smoothing and scaling programs are included.

  9. 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)

  10. Effects of a trailing edge flap on the aerodynamics and acoustics of rotor blade-vortex interactions

    NASA Technical Reports Server (NTRS)

    Charles, B. D.; Tadghighi, H.; Hassan, A. A.

    1992-01-01

    The use of a trailing edge flap on a helicopter rotor has been numerically simulated to determine if such a device can mitigate the acoustics of blade vortex interactions (BVI). The numerical procedure employs CAMRAD/JA, a lifting-line helicopter rotor trim code, in conjunction with RFS2, an unsteady transonic full-potential flow solver, and WOPWOP, an acoustic model based on Farassat's formulation 1A. The codes were modified to simulate trailing edge flap effects. The CAMRAD/JA code was used to compute the far wake inflow effects and the vortex wake trajectories and strengths which are utilized by RFS2 to predict the blade surface pressure variations. These pressures were then analyzed using WOPWOP to determine the high frequency acoustic response at several fixed observer locations below the rotor disk. Comparisons were made with different flap deflection amplitudes and rates to assess flap effects on BVI. Numerical experiments were carried out using a one-seventh scale AH-1G rotor system for flight conditions simulating BVI encountered during low speed descending flight with and without flaps. Predicted blade surface pressures and acoustic sound pressure levels obtained have shown good agreement with the baseline no-flap test data obtained in the DNW wind tunnel. Numerical results indicate that the use of flaps is beneficial in reducing BVI noise.

  11. AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading-and Trailing-Edge Slot Blowing

    NASA Technical Reports Server (NTRS)

    Horne, Clifton; Burnside, Nathan J.

    2013-01-01

    Aeroacoustic measurements of the 11 % scale full-span AMELIA CESTOL model with leading- and trailing-edge slot blowing circulation control (CCW) wing were obtained during a recent test in the Arnold Engineering Development Center 40- by 80-Ft. Wind Tunnel at NASA Ames Research Center, Sound levels and spectra were acquired with seven in-flow microphones and a 48-element phased microphone array for a variety of vehicle configurations, CCW slot flow rates, and forward speeds, Corrections to the measurements and processing are in progress, however the data from selected configurations presented in this report confirm good measurement quality and dynamic range over the test conditions, Array beamform maps at 40 kts tunnel speed show that the trailing edge flap source is dominant for most frequencies at flap angles of 0deg and 60deg, The overall sound level for the 60deg flap was similar to the 0deg flap for most slot blowing rates forward of 90deg incidence, but was louder by up to 6 dB for downstream angles, At 100 kts, the in-flow microphone levels were louder than the sensor self-noise for the higher blowing rates, while passive and active background noise suppression methods for the microphone array revealed source levels as much as 20 dB lower than observed with the in-flow microphones,

  12. Robust design of multiple trailing edge flaps for helicopter vibration reduction: A multi-objective bat algorithm approach

    NASA Astrophysics Data System (ADS)

    Mallick, Rajnish; Ganguli, Ranjan; Seetharama Bhat, M.

    2015-09-01

    The objective of this study is to determine an optimal trailing edge flap configuration and flap location to achieve minimum hub vibration levels and flap actuation power simultaneously. An aeroelastic analysis of a soft in-plane four-bladed rotor is performed in conjunction with optimal control. A second-order polynomial response surface based on an orthogonal array (OA) with 3-level design describes both the objectives adequately. Two new orthogonal arrays called MGB2P-OA and MGB4P-OA are proposed to generate nonlinear response surfaces with all interaction terms for two and four parameters, respectively. A multi-objective bat algorithm (MOBA) approach is used to obtain the optimal design point for the mutually conflicting objectives. MOBA is a recently developed nature-inspired metaheuristic optimization algorithm that is based on the echolocation behaviour of bats. It is found that MOBA inspired Pareto optimal trailing edge flap design reduces vibration levels by 73% and flap actuation power by 27% in comparison with the baseline design.

  13. Energy Harvesting of a Flapping Airfoil in a Vortical Wake

    NASA Astrophysics Data System (ADS)

    Zheng, Z. Charlie; Wei, Zhenglun

    2014-11-01

    We study the response of a two-dimensional flapping airfoil in the wake downstream of an oscillating D-shape cylinder. The airfoil has either heaving or pitching motions. The leading edge vortex (LEV) and trailing edge vortex (TEV) of the airfoil play important roles in energy harvesting. Two major interaction modes between the airfoil and incoming vortices, the suppressing mode and the reinforcing mode, are identified. However, distinctions exist between the heaving and pitching motion in terms of their contributions to the interaction modes and the efficiency of the energy extraction. A potential theory and the related fluid dynamics analysis are developed to analytically demonstrate that the topology of the incoming vortices corresponding to the airfoil is the primary factor that determines the interaction modes. Finally, the trade-off between the input and the output is discussed. It is found that appropriate operational parameters for the heaving motion are preferable in order to preserve acceptable input power for energy harvesters, while appropriate parameters for the pitching motion are essential to achieve decent output power.

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

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Winstroth, J.; Seume, J. R.

    2016-09-01

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

  16. Flap-Edge Blowing Experiments

    NASA Technical Reports Server (NTRS)

    Gaeta, R. J.; Englar, R. J.; Ahuja, K. K.

    2003-01-01

    This Appendix documents the salient results from an effort to mitigate the so-called flap-edge noise generated at the split between a flap edge that is deployed and the undeployed flap. Utilizing a Coanda surface installed at the flap edge, steady blowing was used in an attempt to diminish the vortex strength resulting from the uneven lift distribution. The strength of this lifting vortex was augmented by steady blowing over the deployed flap. The test article for this study was the same 2D airfoil used in the steady blowing program reported earlier (also used in pulsed blowing tests, see Appendix G), however its trailing edge geometry was modified. An exact duplicate of the airfoil shape was made out of fiberglass with no flap, and in the clean configuration. It was attached to the existing airfoil to make an airfoil that has half of its flap deployed and half un-deployed. Figure 1 shows a schematic of the planform showing the two areas where steady blowing was introduced. The flap-edge blowing or the auxiliary blowing was in the direction normal to the freestream velocity vector. Slot heights for the blowing chambers were on the order of 0.0 14 inches.

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

    NASA Technical Reports Server (NTRS)

    Adair, D.; Horne, W. C.

    1989-01-01

    Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analyzed. The flow remained attached over the model surfaces, except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 x 10 to the 6th in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. The flow was complicated by the presence of a strong, initially inviscid, jet, emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer.

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

    NASA Technical Reports Server (NTRS)

    Sorenson, R. L.

    1982-01-01

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

  19. Cooled airfoil in a turbine engine

    SciTech Connect

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

    2015-04-21

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

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

    NASA Astrophysics Data System (ADS)

    Zhang, Kai; Wei, Tian; Hu, Hui

    2015-09-01

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

  1. AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading- and Trailing-Edge Slot Blowing

    NASA Technical Reports Server (NTRS)

    Horne, William C.; Burnside, Nathan J.

    2013-01-01

    The AMELIA Cruise-Efficient Short Take-off and Landing (CESTOL) configuration concept was developed to meet future requirements of reduced field length, noise, and fuel burn by researchers at Cal Poly, San Luis Obispo and Georgia Tech Research Institute under sponsorship by the NASA Fundamental Aeronautics Program (FAP), Subsonic Fixed Wing Project. The novel configuration includes leading- and trailing-edge circulation control wing (CCW), over-wing podded turbine propulsion simulation (TPS). Extensive aerodynamic measurements of forces, surfaces pressures, and wing surface skin friction measurements were recently measured over a wide range of test conditions in the Arnold Engineering Development Center(AEDC) National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Ft Wind Tunnel. Acoustic measurements of the model were also acquired for each configuration with 7 fixed microphones on a line under the left wing, and with a 48-element, 40-inch diameter phased microphone array under the right wing. This presentation will discuss acoustic characteristics of the CCW system for a variety of tunnel speeds (0 to 120 kts), model configurations (leading edge(LE) and/or trailing-edge(TE) slot blowing, and orientations (incidence and yaw) based on acoustic measurements acquired concurrently with the aerodynamic measurements. The flow coefficient, Cmu= mVSLOT/qSW varied from 0 to 0.88 at 40 kts, and from 0 to 0.15 at 120 kts. Here m is the slot mass flow rate, VSLOT is the slot exit velocity, q is dynamic pressure, and SW is wing surface area. Directivities at selected 1/3 octave bands will be compared with comparable measurements of a 2-D wing at GTRI, as will as microphone array near-field measurements of the right wing at maximum flow rate. The presentation will include discussion of acoustic sensor calibrations as well as characterization of the wind tunnel background noise environment.

  2. 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.

  3. An experimental study of a turbulent boundary layer in the trailing edge region of a circulation-control airfoil

    NASA Technical Reports Server (NTRS)

    Brown, Jeffrey

    1995-01-01

    This report describes the development of a new near-wall LDV technique and its use to measure the full turbulent stress tensor in a highly three-dimensional separated boundary layer flow. Laser Doppler velocimetry, with small diameter laser beams allowing small beam-crossing angles (i.e., larger fringe spacings), was used, resulting in measurements close to solid surfaces. Dennis Johnson of NASA/Ames combined the advantages offered by FFT signal processors with a beam-steering probe inserted in the flow to devise a new self-aligning, near-wall LDV approach capable of making direct measurements of the transverse velocity component. This new technique was applied to a 2D, zero-pressure-gradient, flat-plate boundary layer in the High Reynolds Number Pilot Channel at Ames. Accurate and repeatable measurements of U and uu to within 30 mm of the wall were obtained. Next, the near-wall LDV technique was employed in a 3D separated boundary layer flow, consisting of a 2-dimensional wall boundary layer approaching a right-angle wedge. The resulting lateral and adverse streamwise pressure gradients generated significant spanwise velocities and turbulent stresses, and a large reverse-flow region. The near-wall LDV was modified further to facilitate measurements in the 3D Wedge Flow.

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

    NASA Technical Reports Server (NTRS)

    Shrewsbury, George D.

    1987-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1979-01-01

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

  6. Theoretical estimation of the transonic aerodynamic characteristics of a supercritical-wing transport model with trailing-edge controls

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.; Mann, M. J.

    1978-01-01

    A method for rapidly estimating the overall forces and moments at supercritical speeds, below drag divergence, of transport configurations with supercritical wings is presented. The method was also used for estimating the rolling moments due to the deflection of wing trailing-edge controls. This analysis was based on a vortex-lattice technique modified to approximate the effects of wing thickness and boundary-layer induced camber. Comparisons between the results of this method and experiment indicate reasonably good correlation of the lift, pitching moment, and rolling moment. The method required much less storage and run time to compute solutions over an angle-of-attack range than presently available transonic nonlinear methods require for a single angle-of-attack solution.

  7. Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

    NASA Technical Reports Server (NTRS)

    Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

    2014-01-01

    This is the presentation to follow conference paper of the same name. The adaptive compliant trailing edge (ACTE) flap experiment safety of flight requires that the flap to wing interface loads be sensed and monitored in real time to ensure that the wing structural load limits are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces will be monitored and each contains four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty one applied test load cases were developed using the predicted in-flight loads for the ACTE experiment.

  8. SIMS chemical analysis of extended impact features from the trailing edge portion of experiment AO187-2

    NASA Technical Reports Server (NTRS)

    Amari, Sachiko; Foote, John; Simon, Charles; Jessberger, Elmar K.; Lange, Gundolf; Stadermann, Frank; Swan, Pat; Walker, Robert M.; Zinner, Ernst

    1992-01-01

    One hundred capture cells from the trailing edge, which had lost their cover foils during flight, were optically scanned for extended impact features caused by high velocity projectiles impinging on the cells while the foils were still intact. Of the 53 candidates, 24 impacts were analyzed by secondary ion mass spectroscopy for the chemical composition of the deposits. Projectile material was found in all impacts, and at least 75 percent of them appear to be caused by interplanetary dust particles. Elemental ratios are fractionated, with refractory elements enriched in the impacts relative to interplanetary dust particles collected in the stratosphere. Although this could be due to systematic differences in the compositions, a more likely explanation is volatility fractionation during the impact process.

  9. Longitudinal Aerodynamic Modeling of the Adaptive Compliant Trailing Edge Flaps on a GIII Airplane and Comparisons to Flight Data

    NASA Technical Reports Server (NTRS)

    Smith, Mark S.; Bui, Trong T.; Garcia, Christian A.; Cumming, Stephen B.

    2016-01-01

    A pair of compliant trailing edge flaps was flown on a modified GIII airplane. Prior to flight test, multiple analysis tools of various levels of complexity were used to predict the aerodynamic effects of the flaps. Vortex lattice, full potential flow, and full Navier-Stokes aerodynamic analysis software programs were used for prediction, in addition to another program that used empirical data. After the flight-test series, lift and pitching moment coefficient increments due to the flaps were estimated from flight data and compared to the results of the predictive tools. The predicted lift increments matched flight data well for all predictive tools for small flap deflections. All tools over-predicted lift increments for large flap deflections. The potential flow and Navier-Stokes programs predicted pitching moment coefficient increments better than the other tools.

  10. 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.

  11. On the impact of multi-axial stress states on trailing edge bondlines in wind turbine rotor blades

    NASA Astrophysics Data System (ADS)

    Noever Castelos, Pablo; Balzani, Claudio

    2016-09-01

    For a reliable design of wind turbine systems all of their components have to be designed to withstand the loads appearing in the turbine's lifetime. When performed in an integral manner this is called systems engineering, and is exceptionally important for components that have an impact on the entire wind turbine system, such as the rotor blade. Bondlines are crucial subcomponents of rotor blades, but they are not much recognized in the wind energy research community. However, a bondline failure can lead to the loss of a rotor blade, and potentially of the entire turbine, and is extraordinarily relevant to be treated with strong emphasis when designing a wind turbine. Modern wind turbine rotor blades with lengths of 80 m and more offer a degree of flexibility that has never been seen in wind energy technology before. Large deflections result in high strains in the adhesive connections, especially at the trailing edge. The latest edition of the DNV GL guideline from end of 2015 demands a three-dimensional stress analysis of bondlines, whereas before an isolated shear stress proof was sufficient. In order to quantify the lack of safety from older certification guidelines this paper studies the influence of multi-axial stress states on the ultimate and fatigue load resistance of trailing edge adhesive bonds. For this purpose, detailed finite element simulations of the IWES IWT-7.5-164 reference wind turbine blades are performed. Different yield criteria are evaluated for the prediction of failure and lifetime. The results show that the multi-axial stress state is governed by span-wise normal stresses. Those are evidently not captured in isolated shear stress proofs, yielding non-conservative estimates of lifetime and ultimate load resistance. This finding highlights the importance to include a three-dimensional stress state in the failure analysis of adhesive bonds in modern wind turbine rotor blades, and the necessity to perform a three-dimensional characterization

  12. Evaluation of the Hinge Moment and Normal Force Aerodynamic Loads from a Seamless Adaptive Compliant Trailing Edge Flap in Flight

    NASA Technical Reports Server (NTRS)

    Miller, Eric J.; Cruz, Josue; Lung, Shun-Fat; Kota, Sridhar; Ervin, Gregory; Lu, Kerr-Jia; Flick, Pete

    2016-01-01

    A seamless adaptive compliant trailing edge (ACTE) flap was demonstrated in flight on a Gulfstream III aircraft at the NASA Armstrong Flight Research Center. The trailing edge flap was deflected between minus 2 deg up and plus 30 deg down in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The interface hardware instrumentation layout and load calibration are discussed. Twenty-one applied calibration test load cases were developed for each individual fitting. The 2-sigma residual errors for the hinge moment was calculated to be 2.4 percent, and for normal force was calculated to be 7.3 percent. The hinge moment and normal force generated by the ACTE flap with a hinge point located at 26-percent wing chord were measured during steady state and symmetric pitch maneuvers. The loads predicted from analysis were compared to the loads observed in flight. The hinge moment loads showed good agreement with the flight loads while the normal force loads calculated from analysis were over-predicted by approximately 20 percent. Normal force and hinge moment loads calculated from the pressure sensors located on the ACTE showed good agreement with the loads calculated from the installed strain gages.

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

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  14. A supercritical airfoil experiment

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Welte, D.

    1978-01-01

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

  16. 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.

  17. 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.

  18. Design and simulative experiment of an innovative trailing edge morphing mechanism driven by artificial muscles embedded in skin

    NASA Astrophysics Data System (ADS)

    Li, Hongda; Liu, Long; Xiao, Tianhang; Ang, Haisong

    2016-09-01

    In this paper, conceptual design of a tailing edge morphing mechanism developed based on a new kind of artificial muscle embedded in skin, named Driving Skin, is proposed. To demonstrate the feasibility of this conceptual design, an experiment using ordinary fishing lines to simulate the function of artificial muscles was designed and carried out. Some measures were designed to ensure measurement accuracy. The experiment result shows that the contraction ratio and force required by the morphing mechanism can be satisfied by the new artificial muscles, and a relationship between contraction ratios and morphing angles can be found. To demonstrate the practical application feasibility of this conceptual design, a wing section using ordinary ropes to simulate the function of the Driving Skin mechanism was designed and fabricated. The demonstration wing section, extremely light in weight and capable of changing thickness, performs well, with a -30^\\circ /+30^\\circ morphing angle achieved. The trailing edge morphing mechanism is efficient in re-contouring the wing profile.

  19. 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.

  20. Artificial Selection Reveals High Genetic Variation in Phenology at the Trailing Edge of a Species Range.

    PubMed

    Sheth, Seema Nayan; Angert, Amy Lauren

    2016-02-01

    Species responses to climate change depend on the interplay of migration and adaptation, yet we know relatively little about the potential for adaptation. Genetic adaptations to climate change often involve shifts in the timing of phenological events, such as flowering. If populations at the edge of a species range have lower genetic variation in phenological traits than central populations, then their persistence under climate change could be threatened. To test this hypothesis, we performed artificial selection experiments using the scarlet monkeyflower (Mimulus cardinalis) and compared genetic variation in flowering time among populations at the latitudinal center, northern edge, and southern edge of the species range. We also assessed whether selection on flowering time yielded correlated responses in functional traits, potentially representing a cost associated with early or late flowering. Contrary to prediction, southern populations exhibited greater responses to selection on flowering time than central or northern populations. Further, selection for early flowering resulted in correlated increases in specific leaf area and leaf nitrogen, whereas selection for late flowering led to decreases in these traits. These results provide critical insights about how spatial variation in the potential for adaptation may affect population persistence under changing climates. PMID:26807746

  1. Aeroelastic Analysis of a Flexible Wing Wind Tunnel Model with Variable Camber Continuous Trailing Edge Flap Design

    NASA Technical Reports Server (NTRS)

    Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia

    2015-01-01

    This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope

  2. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

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

    1945-01-01

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

  3. Parametric Evaluation of Thin, Transonic Circulation-Control Airfoils

    NASA Technical Reports Server (NTRS)

    Schlecht, Robin; Anders, Scott

    2007-01-01

    Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge [03] performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.

  4. Aerodynamic Flight-Test Results for the Adaptive Compliant Trailing Edge

    NASA Technical Reports Server (NTRS)

    Cumming, Stephen B.; Smith, Mark S.; Ali, Aliyah N.; Bui, Trong T.; Ellsworth, Joel C.; Garcia, Christian A.

    2016-01-01

    The aerodynamic effects of compliant flaps installed onto a modified Gulfstream III airplane were investigated. Analyses were performed prior to flight to predict the aerodynamic effects of the flap installation. Flight tests were conducted to gather both structural and aerodynamic data. The airplane was instrumented to collect vehicle aerodynamic data and wing pressure data. A leading-edge stagnation detection system was also installed. The data from these flights were analyzed and compared with predictions. The predictive tools compared well with flight data for small flap deflections, but differences between predictions and flight estimates were greater at larger deflections. This paper describes the methods used to examine the aerodynamics data from the flight tests and provides a discussion of the flight-test results in the areas of vehicle aerodynamics, wing sectional pressure coefficient profiles, and air data.

  5. Prey switching as a means of enhancing persistence in predators at the trailing southern edge.

    PubMed

    Peers, Michael J L; Wehtje, Morgan; Thornton, Daniel H; Murray, Dennis L

    2014-04-01

    Understanding the effects of climate change on species' persistence is a major research interest; however, most studies have focused on responses at the northern or expanding range edge. There is a pressing need to explain how species can persist at their southern range when changing biotic interactions will influence species occurrence. For predators, variation in distribution of primary prey owing to climate change will lead to mismatched distribution and local extinction, unless their diet is altered to more extensively include alternate prey. We assessed whether addition of prey information in climate projections restricted projected habitat of a specialist predator, Canada lynx (Lynx canadensis), and if switching from their primary prey (snowshoe hare; Lepus americanus) to an alternate prey (red squirrel; Tamiasciurus hudsonicus) mitigates range restriction along the southern range edge. Our models projected distributions of each species to 2050 and 2080 to then refine predictions for southern lynx on the basis of varying combinations of prey availability. We found that models that incorporated information on prey substantially reduced the total predicted southern range of lynx in both 2050 and 2080. However, models that emphasized red squirrel as the primary species had 7-24% lower southern range loss than the corresponding snowshoe hare model. These results illustrate that (i) persistence at the southern range may require species to exploit higher portions of alternate food; (ii) selection may act on marginal populations to accommodate phenotypic changes that will allow increased use of alternate resources; and (iii) climate projections based solely on abiotic data can underestimate the severity of future range restriction. In the case of Canada lynx, our results indicate that the southern range likely will be characterized by locally varying levels of mismatch with prey such that the extent of range recession or local adaptation may appear as a geographical

  6. The interbranchial lymphoid tissue of Atlantic Salmon (Salmo salar L) extends as a diffuse mucosal lymphoid tissue throughout the trailing edge of the gill filament.

    PubMed

    Dalum, Alf S; Austbø, Lars; Bjørgen, Håvard; Skjødt, Karsten; Hordvik, Ivar; Hansen, Tom; Fjelldal, Per G; Press, Charles McL; Griffiths, David J; Koppang, Erling O

    2015-09-01

    The teleost gill forms an extensive, semipermeable barrier that must tolerate intimate contact with the surrounding environment and be able to protect the body from external pathogens. The recent discovery of the interbranchial lymphoid tissue (ILT) has initiated an anatomical and functional investigation of the lymphoid tissue of the salmonid gill. In this article, sectioning of gill arches in all three primary planes revealed an elongation of the ILT outward along the trailing edge of the primary filament to the very distal end, a finding not previously described. This newly found lymphoid tissue was investigated using a range of morphological and transcriptional tools. Avoiding potential salinity-related effects, the study focused on two fresh-water life stages-smoltifying juveniles and mature adults. Aggregates of T-cells continuous with the ILT were found within the thick epithelial lining of the trailing edge of the filament in considerably larger numbers than seen in the epithelium of the leading edge and of the interlamellar area. Only a few of these cells were identified as CD8α(+) -cells, and there was a significantly (P < 0.05) higher relative expression of CD4- than of CD8- related genes in all gill segments investigated. Numerous major histocompatibility complex class II(+) -cells were distributed uniformly throughout the filament epithelial tissue. Few Ig(+) -cells were detected. Overall, the morphological features and comparable immune gene expression of the previously described ILT and the filament trailing edge lymphoid tissue suggest a close functional and anatomical relationship. We propose that the anatomical definition of the ILT must be broadened to include both the previously described ILT (to be renamed proximal ILT) and the trailing edge lymphoid tissue (to be named distal ILT). This extended anatomical localisation identifies the ILT as a widely distributed mucosal lymphoid tissue in the gill of Atlantic salmon.

  7. Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

    NASA Technical Reports Server (NTRS)

    Maki, Ralph L.

    1959-01-01

    Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

  8. General perspectives on model construction and evaluation for stochastic estimation, with application to a blunt trailing edge wake

    NASA Astrophysics Data System (ADS)

    Clark, H.; Naghib-Lahouti, A.; Lavoie, P.

    2014-07-01

    The technique of stochastic estimation is examined as a specific application of linear least squares modelling. Factors that are relevant to the objectives of estimation in fluids, such as the number of sensors, the use of multiple time lags, and the strength of linear correlations, are discussed in the context of a general regression formulation. We consolidate the established findings of several research fields in order to outline clearly the potential pitfalls and reasonable performance expectations of these empirical strategies. Experimental measurements of velocity and fluctuating pressure in the wake of a blunt trailing edge body are used for quantitative illustration of key considerations for model construction and performance evaluation. It is emphasized that estimator accuracy is influenced strongly by the physical relationships among the measured variables, in addition to their correlation with the estimated variable. The evaluation of several performance metrics on an independent test set provides valuable information for the selection of a suitably complex model. In particular, "variance inflation" is interpreted as an indicator of the potential amplification of noise by a stochastic estimator.

  9. Vibration reduction in helicopter rotors using an actively controlled partial span trailing edge flap located on the blade

    NASA Technical Reports Server (NTRS)

    Millott, T. A.; Friedmann, P. P.

    1994-01-01

    This report describes an analytical study of vibration reduction in a four-bladed helicopter rotor using an actively controlled, partial span, trailing edge flap located on the blade. The vibration reduction produced by the actively controlled flap (ACF) is compared with that obtained using individual blade control (IBC), in which the entire blade is oscillated in pitch. For both cases a deterministic feedback controller is implemented to reduce the 4/rev hub loads. For all cases considered, the ACF produced vibration reduction comparable with that obtained using IBC, but consumed only 10-30% of the power required to implement IBC. A careful parametric study is conducted to determine the influence of blade torsional stiffness, spanwise location of the control flap, and hinge moment correction on the vibration reduction characteristics of the ACF. The results clearly demonstrate the feasibility of this new approach to vibration reduction. It should be emphasized than the ACF, used together with a conventional swashplate, is completely decoupled from the primary flight control system and thus it has no influence on the airworthiness of the helicopter. This attribute is potentially a significant advantage when compared to IBC.

  10. 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.

  11. 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.

  12. Airfoil in sinusoidal motion in a pulsating stream

    NASA Technical Reports Server (NTRS)

    Greenberg, J Mayo

    1947-01-01

    The forces and moments on a two-dimensional airfoil executing harmonic motions in a pulsating stream are derived on the basis of non-stationary incompressible potential flow theory, with the inclusion of the effect of the continuous sheet of vortices shed from the trailing edge. An assumption as to the form of the wake is made with a certain degree of approximation. A comparison with previous work applicable only to the special case of a stationary airfoil is made by means of a numerical example, and the excellent agreement obtained shows that the wake approximation is quite sufficient. The results obtained are expected to be useful in considerations of forced vibrations and flutter of rotary wing aircraft.

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

    NASA Astrophysics Data System (ADS)

    Lee, B. H. K.

    2001-02-01

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

  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. Design and Experimental Results for the S825 Airfoil; Period of Performance: 1998-1999

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A 17%-thick, natural-laminar-flow airfoil, the S825, for the 75% blade radial station of 20- to 40-meter, variable-speed and variable-pitch (toward feather), horizontal-axis wind turbines has been designed and analyzed theoretically and verified experimentally in the NASA Langley Low-Turbulence Pressure Tunnel. The two primary objectives of high maximum lift, relatively insensitive to roughness and low-profile drag have been achieved. The airfoil exhibits a rapid, trailing-edge stall, which does not meet the design goal of a docile stall. The constraints on the pitching moment and the airfoil thickness have been satisfied. Comparisons of the theoretical and experimental results generally show good agreement.

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

    NASA Technical Reports Server (NTRS)

    Crimi, P.

    1974-01-01

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

  17. Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade - Part 1: Experimental Measurements. Part 1; Experimental Measurements

    NASA Technical Reports Server (NTRS)

    Bunker, Ronald S.; Wetzel, Todd G.; Rigby, David L.; Reddy, D. R. (Technical Monitor)

    2000-01-01

    A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage. The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage. The geometry of the trailing edge passage is that of a two-pass serpentine circuit with a sharp 180-degree turning region at the tip. The upflow channel is split by interrupted ribs into two major subchannels, one of which is turbulated. This channel has an average aspect ratio of roughly 14:1. The spanwise extent of the channel geometry includes both area convergence from root to tip, as well as taper towards the trailing edge apex. The average section Reynolds numbers tested in this upflow channel range from 55,000 to 98,000. The tip section contains a turning vane near the extreme comer. The downflow channel has an aspect ratio of about 5:1, and also includes convergence and taper. Turbulators of varying sizes are included in this channel also. Both detailed heat transfer and pressure distribution measurements are presented. The pressure measurements are incorporated into a flow network model illustrating the major loss contributors.

  18. Parameter study of simplified dragonfly airfoil geometry at Reynolds number of 6000.

    PubMed

    Levy, David-Elie; Seifert, Avraham

    2010-10-21

    Aerodynamic study of a simplified Dragonfly airfoil in gliding flight at Reynolds numbers below 10,000 is motivated by both pure scientific interest and technological applications. At these Reynolds numbers, the natural insect flight could provide inspiration for technology development of Micro UAV's and more. Insect wings are typically characterized by corrugated airfoils. The present study follows a fundamental flow physics study (Levy and Seifert, 2009), that revealed the importance of flow separation from the first corrugation, the roll-up of the separated shear layer to discrete vortices and their role in promoting flow reattachment to the aft arc, as the leading mechanism enabling high-lift, low drag performance of the Dragonfly gliding flight. This paper describes the effect of systematic airfoil geometry variations on the aerodynamic properties of a simplified Dragonfly airfoil at Reynolds number of 6000. The parameter study includes a detailed analysis of small variations of the nominal geometry, such as corrugation placement or height, rear arc and trailing edge shape. Numerical simulations using the 2D laminar Navier-Stokes equations revealed that the flow accelerating over the first corrugation slope is followed by an unsteady pressure recovery, combined with vortex shedding. The latter allows the reattachment of the flow over the rear arc. Also, the drag values are directly linked to the vortices' magnitude. This parametric study shows that geometric variations which reduce the vortices' amplitude, as reduction of the rear cavity depth or the reduction of the rear arc and trailing edge curvature, will reduce the drag values. Other changes will extend the flow reattachment over the rear arc for a larger mean lift coefficients range; such as the negative deflection of the forward flat plate. These changes consequently reduce the drag values at higher mean lift coefficients. The detailed geometry study enabled the definition of a corrugated airfoil

  19. Parameter study of simplified dragonfly airfoil geometry at Reynolds number of 6000.

    PubMed

    Levy, David-Elie; Seifert, Avraham

    2010-10-21

    Aerodynamic study of a simplified Dragonfly airfoil in gliding flight at Reynolds numbers below 10,000 is motivated by both pure scientific interest and technological applications. At these Reynolds numbers, the natural insect flight could provide inspiration for technology development of Micro UAV's and more. Insect wings are typically characterized by corrugated airfoils. The present study follows a fundamental flow physics study (Levy and Seifert, 2009), that revealed the importance of flow separation from the first corrugation, the roll-up of the separated shear layer to discrete vortices and their role in promoting flow reattachment to the aft arc, as the leading mechanism enabling high-lift, low drag performance of the Dragonfly gliding flight. This paper describes the effect of systematic airfoil geometry variations on the aerodynamic properties of a simplified Dragonfly airfoil at Reynolds number of 6000. The parameter study includes a detailed analysis of small variations of the nominal geometry, such as corrugation placement or height, rear arc and trailing edge shape. Numerical simulations using the 2D laminar Navier-Stokes equations revealed that the flow accelerating over the first corrugation slope is followed by an unsteady pressure recovery, combined with vortex shedding. The latter allows the reattachment of the flow over the rear arc. Also, the drag values are directly linked to the vortices' magnitude. This parametric study shows that geometric variations which reduce the vortices' amplitude, as reduction of the rear cavity depth or the reduction of the rear arc and trailing edge curvature, will reduce the drag values. Other changes will extend the flow reattachment over the rear arc for a larger mean lift coefficients range; such as the negative deflection of the forward flat plate. These changes consequently reduce the drag values at higher mean lift coefficients. The detailed geometry study enabled the definition of a corrugated airfoil

  20. Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

    NASA Technical Reports Server (NTRS)

    Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

    2014-01-01

    The safety-of-flight parameters for the Adaptive Compliant Trailing Edge (ACTE) flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces were monitored; each contained four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty-one applied test load cases were developed using the predicted in-flight loads. Pre-test predictions of strain gage responses were produced using finite element method models of the interface fittings. Predicted and measured test strains are presented. A load testing rig and three hydraulic jacks were used to apply combinations of shear, bending, and axial loads to the interface fittings. Hardware deflections under load were measured using photogrammetry and transducers. Due to deflections in the interface fitting hardware and test rig, finite element model techniques were used to calculate the reaction loads throughout the applied load range, taking into account the elastically-deformed geometry. The primary load equations were selected based on multiple calibration metrics. An independent set of validation cases was used to validate each derived equation. The 2-sigma residual errors for the shear loads were less than eight percent of the full-scale calibration load; the 2-sigma residual errors for the bending moment loads were less than three percent of the full-scale calibration load. The derived load equations for shear, bending, and axial loads are presented, with the calculated errors for both the calibration cases and the independent validation load cases.

  1. Airfoil structure

    SciTech Connect

    Frey, Gary A.; Twardochleb, Christopher Z.

    1998-01-01

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

  2. 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.

  3. Airplane wing leading edge variable camber flap

    NASA Technical Reports Server (NTRS)

    Cole, J. B.

    1980-01-01

    The invention and design of an aerodynamic high lift device which provided a solution to an aircraft performance problem are described. The performance problem of converting a high speed cruise airfoil into a low speed aerodynamic shape that would provide landing and take-off characteristics superior to those available with contemporary high lift devices are addressed. The need for an improved wing leading edge device that would complement the high lift performance of a triple slotted trailing edge flap is examined. The mechanical and structural aspects of the variable camber flap are discussed and the aerodynamic performance aspects only as they relate to the invention and design of the device are presented.

  4. Lock-in of elastically mounted airfoils at a 90° angle of attack

    NASA Astrophysics Data System (ADS)

    Ehrmann, R. S.; Loftin, K. M.; Johnson, S.; White, E. B.

    2014-01-01

    Reducing vortex-induced vibration (VIV) of elastically mounted cylinders has applications to petroleum, nuclear, and civil engineering. One simple method is streamlining the cylinder into an airfoil shape. However, if flow direction changes, an elastic airfoil could experience similar oscillations with even more drag. To better understand a general airfoil's response, three elastically mounted airfoil shapes are tested at a 90° angle of attack in a 3 ft by 4 ft wind tunnel. The shapes are a NACA 0018, a sharp leading- and trailing-edge (sharp-sharp) model, and a round leading- and trailing-edge (round-round) model. Mass-damping ranges from 0.96 to 1.44. For comparison to canonical VIV research, a cylinder is also tested. Since lock-in occurs near Rec=125×103, the models are also tested with a trip strip. The NACA 0018 and sharp-sharp configuration show nearly identical responses. The cylinder and round-round airfoil have responses five to eight times larger. Thus, the existence of a single sharp edge is sufficient to greatly reduce VIV at 90° angle of attack. Whereas the cylinder and round-round maximum response amplitudes are similar, cylinder lock-in occurs over a velocity range three times larger than the round-round. The tripped cylinder and round-round models' response is attenuated by 70% compared to their respective clean configurations. Hysteresis is only observed in the circular cylinder and round-round models. Hotwire data indicates the clean cylinder has a unique vortex pattern compared to the other configurations.

  5. AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (CDC VERSION)

    NASA Technical Reports Server (NTRS)

    Darden, C. M.

    1994-01-01

    This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is

  6. AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (IBM PC VERSION)

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.

    1994-01-01

    This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is

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

    NASA Technical Reports Server (NTRS)

    Mccullough, George B; Gault, Donald E

    1947-01-01

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

  8. The aerodynamic design of an advanced rotor airfoil

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  9. Large Eddy Simulation of Airfoil Self-Noise at High Reynolds Number

    NASA Astrophysics Data System (ADS)

    Kocheemoolayil, Joseph; Lele, Sanjiva

    2015-11-01

    The trailing edge noise section (Category 1) of the Benchmark Problems for Airframe Noise Computations (BANC) workshop features five canonical problems. No first-principles based approach free of empiricism and tunable coefficients has successfully predicted trailing edge noise for the five configurations to date. Our simulations predict trailing edge noise accurately for all five configurations. The simulation database is described in detail, highlighting efforts undertaken to validate the results through systematic comparison with dedicated experiments and establish insensitivity to grid resolution, domain size, alleatory uncertainties such as the tripping mechanism used to force transition to turbulence and epistemic uncertainties such as models for unresolved near-wall turbulence. Ongoing efforts to extend the predictive capability to non-canonical configurations featuring flow separation are summarized. A novel, large-span calculation that predicts the flow past a wind turbine airfoil in deep stall with unprecedented accuracy is presented. The simulations predict airfoil noise in the near-stall regime accurately. While the post-stall noise predictions leave room for improvement, significant uncertainties in the experiment might preclude a fair comparison in this regime. We thank Cascade Technologies Inc. for providing access to the CharLES toolkit - a massively-parallel, unstructured large eddy simulation framework.

  10. Heat transfer in internal channel of a blade: Effects of rotation in a trailing edge cooling system

    NASA Astrophysics Data System (ADS)

    Andrei, Luca; Andreini, Antonio; Bonanni, Leonardo; Facchini, Bruno

    2012-06-01

    The aerothermal performance of a trailing edge (TE) internal cooling system of a high pressure gas turbine blade was evaluated under stationary and rotating conditions. The investigated geometry consists of a 30:1 scaled model reproducing a typical wedge shaped discharge duct with one row of enlarged pedestals. The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90 [deg] turning flow from the radial hub inlet to the tangential TE outlet. Two different tip configurations were tested, the first one with a completely closed section, the second one with a 5 holes outlet surfaces discharging at ambient pressure. In order to assess rotation effects, a rotating test rig, composed of a rotating arm holding both the PMMA TE model and the instrumentation, was purposely developed and manufactured. A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements on the blade pressure side. A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus; moreover several slip rings are used for both instrumentation power supply and thermocouple connection. A parallel CFD analysis involving steady-state RANS modeling was conducted to allow an insight of the flow field inside the redirecting channel and the interpedestal ducts to better interpret the developing vortical structures. Low-Reynolds grid clustering permits to integrate up to the wall both the momentum and the thermal boundary layer. Calculations were performed by means of an in-house developed pressure based solver exploiting the k-ω SST turbulence model implemented in the framework of the open-source finite volume discretization toolbox OpenFOAM®. Analyzed flow conditions correspond to Reynolds number of 20000 in the hub inlet section and angular speed varies to obtain rotation numbers in the range from 0 to 0.3. The orientation of the rotation axis is orthogonal to the

  11. 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.

  12. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

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

  14. Application of digital holographic interferometry to pressure measurements of symmetric, supercritical and circulation-control airfoils in transonic flow fields

    NASA Technical Reports Server (NTRS)

    Torres, Francisco J.

    1987-01-01

    Six airfoil interferograms were evaluated using a semiautomatic image-processor system which digitizes, segments, and extracts the fringe coordinates along a polygonal line. The resulting fringe order function was converted into density and pressure distributions and a comparison was made with pressure transducer data at the same wind tunnel test conditions. Three airfoil shapes were used in the evaluation to test the capabilities of the image processor with a variety of flows. Symmetric, supercritical, and circulation-control airfoil interferograms provided fringe patterns with shocks, separated flows, and high-pressure regions for evaluation. Regions along the polygon line with very clear fringe patterns yielded results within 1% of transducer measurements, while poorer quality regions, particularly near the leading and trailing edges, yielded results that were not as good.

  15. Two dimensional aerodynamic interference effects on oscillating airfoils with flaps in ventilated subsonic wind tunnels. [computational fluid dynamics

    NASA Technical Reports Server (NTRS)

    Fromme, J.; Golberg, M.; Werth, J.

    1979-01-01

    The numerical computation of unsteady airloads acting upon thin airfoils with multiple leading and trailing-edge controls in two-dimensional ventilated subsonic wind tunnels is studied. The foundation of the computational method is strengthened with a new and more powerful mathematical existence and convergence theory for solving Cauchy singular integral equations of the first kind, and the method of convergence acceleration by extrapolation to the limit is introduced to analyze airfoils with flaps. New results are presented for steady and unsteady flow, including the effect of acoustic resonance between ventilated wind-tunnel walls and airfoils with oscillating flaps. The computer program TWODI is available for general use and a complete set of instructions is provided.

  16. Noise reduction tests of large-scale-model externally blown flap using trailing-edge blowing and partial flap slot covering. [jet aircraft noise reduction

    NASA Technical Reports Server (NTRS)

    Mckinzie, D. J., Jr.; Burns, R. J.; Wagner, J. M.

    1976-01-01

    Noise data were obtained with a large-scale cold-flow model of a two-flap, under-the-wing, externally blown flap proposed for use on future STOL aircraft. The noise suppression effectiveness of locating a slot conical nozzle at the trailing edge of the second flap and of applying partial covers to the slots between the wing and flaps was evaluated. Overall-sound-pressure-level reductions of 5 db occurred below the wing in the flyover plane. Existing models of several noise sources were applied to the test results. The resulting analytical relation compares favorably with the test data. The noise source mechanisms were analyzed and are discussed.

  17. Theoretical Calculations of the Pressures, Forces, and Moments Due to Various Lateral Motions Acting on Tapered Sweptback Vertical Tails with Supersonic Leading and Trailing Edges

    NASA Technical Reports Server (NTRS)

    Margolis, Kenneth; Elliott, Miriam H.

    1960-01-01

    Based on expressions for the linearized velocity potentials and pressure distributions given in NACA Technical Report 1268, formulas for the span load distribution, forces, and moments are derived for families of thin isolated vertical tails with arbitrary aspect ratio, taper ratio, and sweepback performing the motions constant sideslip, steady rolling, steady yawing, and constant lateral acceleration. The range of Mach number considered corresponds, in general, to the condition that the tail leading and trailing edges are supersonic. To supplement the analytical results, design-type charts are presented which enable rapid estimation of the forces and moments (expressed as stability derivatives) for given combinations of geometry parameters and Mach number.

  18. 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.

  19. Wind Tunnel Evaluation of a Model Helicopter Main-Rotor Blade With Slotted Airfoils at the Tip

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W.; Yeager, William T., Jr.; Singleton, Jeffrey D.; Wilbur, Matthew L.; Mirick, Paul H.

    2001-01-01

    Data for rotors using unconventional airfoils are of interest to permit an evaluation of this technology's capability to meet the U.S. Army's need for increased helicopter mission effectiveness and improved safety and survivability. Thus, an experimental investigation was conducted in the Langley Transonic Dynamics Tunnel (TDT) to evaluate the effect of using slotted airfoils in the rotor blade tip region (85 to 100 percent radius) on rotor aerodynamic performance and loads. Four rotor configurations were tested in forward flight at advance ratios from 0.15 to 0.45 and in hover in-ground effect. The hover tip Mach number was 0.627, which is representative of a design point of 4000-ft geometric altitude and a temperature of 95 F. The baseline rotor configuration had a conventional single-element airfoil in the tip region. A second rotor configuration had a forward-slotted airfoil with a -6 deg slat, a third configuration had a forward-slotted airfoil with a -10 slat, and a fourth configuration had an aft-slotted airfoil with a 3 deg flap (trailing edge down). The results of this investigation indicate that the -6 deg slat configuration offers some performance and loads benefits over the other three configurations.

  20. Numerical analysis of active chordwise flexibility on the performance of non-symmetrical flapping airfoils

    NASA Astrophysics Data System (ADS)

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

    2010-01-01

    This paper investigates the effect of active chordwise flexing on the lift, thrust and propulsive efficiency of three types of airfoils. The factors studied are the flexing center location, standard two-sided flexing as well as a type of single-sided flexing. The airfoils are simulated to flap with four configurations, and the effects of flexing under these configurations are investigated. Results show that flexing is not necessarily beneficial for the performance of the airfoils. However, with the correct parameters, efficiency is as high as 0.76 by placing the flexing centre at the trailing edge. The average thrust coefficient is more than twice as high, from 1.63 to 3.57 with flapping and flexing under the right conditions. Moreover, the single-sided flexing also gives an average lift coefficient as high as 4.61 for the S1020 airfoil. The shape of the airfoil does alter the effect of flexing too. Deviating the flexing phase angle away from 90° does not give a significant improvement to the airfoil’s performance. These results greatly enhance the design of a better performing ornithopter wing.

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

    NASA Technical Reports Server (NTRS)

    Bamber, Millard J

    1932-01-01

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

  2. 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.

  3. Flow and Turbulence Modeling and Computation of Shock Buffet Onset for Conventional and Supercritical Airfoils

    NASA Technical Reports Server (NTRS)

    Bartels, Robert E.

    1998-01-01

    Flow and turbulence models applied to the problem of shock buffet onset are studied. The accuracy of the interactive boundary layer and the thin-layer Navier-Stokes equations solved with recent upwind techniques using similar transport field equation turbulence models is assessed for standard steady test cases, including conditions having significant shock separation. The two methods are found to compare well in the shock buffet onset region of a supercritical airfoil that involves strong trailing-edge separation. A computational analysis using the interactive-boundary layer has revealed a Reynolds scaling effect in the shock buffet onset of the supercritical airfoil, which compares well with experiment. The methods are next applied to a conventional airfoil. Steady shock-separated computations of the conventional airfoil with the two methods compare well with experiment. Although the interactive boundary layer computations in the shock buffet region compare well with experiment for the conventional airfoil, the thin-layer Navier-Stokes computations do not. These findings are discussed in connection with possible mechanisms important in the onset of shock buffet and the constraints imposed by current numerical modeling techniques.

  4. 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.

  5. Development of Advanced High Lift Leading Edge Technology for Laminar Flow Wings

    NASA Technical Reports Server (NTRS)

    Bright, Michelle M.; Korntheuer, Andrea; Komadina, Steve; Lin, John C.

    2013-01-01

    This paper describes the Advanced High Lift Leading Edge (AHLLE) task performed by Northrop Grumman Systems Corporation, Aerospace Systems (NGAS) for the NASA Subsonic Fixed Wing project in an effort to develop enabling high-lift technology for laminar flow wings. Based on a known laminar cruise airfoil that incorporated an NGAS-developed integrated slot design, this effort involved using Computational Fluid Dynamics (CFD) analysis and quality function deployment (QFD) analysis on several leading edge concepts, and subsequently down-selected to two blown leading-edge concepts for testing. A 7-foot-span AHLLE airfoil model was designed and fabricated at NGAS and then tested at the NGAS 7 x 10 Low Speed Wind Tunnel in Hawthorne, CA. The model configurations tested included: baseline, deflected trailing edge, blown deflected trailing edge, blown leading edge, morphed leading edge, and blown/morphed leading edge. A successful demonstration of high lift leading edge technology was achieved, and the target goals for improved lift were exceeded by 30% with a maximum section lift coefficient (Cl) of 5.2. Maximum incremental section lift coefficients ( Cl) of 3.5 and 3.1 were achieved for a blown drooped (morphed) leading edge concept and a non-drooped leading edge blowing concept, respectively. The most effective AHLLE design yielded an estimated 94% lift improvement over the conventional high lift Krueger flap configurations while providing laminar flow capability on the cruise configuration.

  6. Flow and heat transfer in an L-shaped cooling passage with ribs and pin fins for the trailing edge of a gas-turbine vane and blade

    NASA Astrophysics Data System (ADS)

    Pardeshi, Irsha

    Efficient and effective cooling of the trailing edges of gas-turbine vanes and blades is challenging because there is very little space to work with. In this study, CFD simulations based on steady RANS closed by the shear-stress transport turbulence model were performed to study the flow and heat transfer in an L-shaped duct for the trailing edge under two operating conditions. One operating condition, referred to as the laboratory condition, where experimental measurements were made, has a Reynolds number at the duct inlet of ReD = 15,000, coolant inlet temperature of Tinlet = 300 K, wall temperature of Twall = 335 K, a back pressure of Pb = 1 atm. When rotating, the angular speed was O = 1,000 rpm. The other condition, referred to as the engine-relevant condition, has Re D = 150,000 at the duct inlet, Tinlet = 673 K, Twall = 1,173 K, and Pb = 25 atm. When rotating, O was 3,600 rpm. The objective is to understand the nature of the flow and heat transfer in an L-shaped cooling passage for the trailing edge that has a combination of ribs and pin fins under rotating and non-rotating conditions with focus on how pin fins and ribs distribute the flow throughout the passage and to understand what features of the flow and heat transfer can or cannot be extrapolated from the laboratory to the engine-relevant operating conditions. When there is no rotation, results obtained show that for both operating conditions, the pin fins minimized the size of the separation bubble when the flow exits the inlet duct into the expanded portion of the L-shaped duct. The size of the separation bubble at the tip of the L-shaped duct created by the adverse pressure gradient is quite large for the laboratory condition and relatively small for the engine condition. Each rib was found to create two sets of recirculating flows, one just upstream of the rib because of the adverse pressure gradient induced by the rib and one just downstream of the rib because of flow separation from a sharp edge

  7. Drag reduction in flow over a two-dimensional bluff body with a blunt trailing edge using a new passive device

    NASA Astrophysics Data System (ADS)

    Park, Hyungmin; Lee, Dongkon; Jeon, Woo-Pyung; Hahn, Seonghyeon; Kim, Jeonglae; Kim, Jungwoo; Choi, Jin; Choi, Haecheon

    2006-09-01

    In this paper, we present a new passive control device for form-drag reduction in flow over a two-dimensional bluff body with a blunt trailing edge. The device consists of small tabs attached to the upper and lower trailing edges of a bluff body to effectively perturb a two-dimensional wake. Both a wind-tunnel experiment and large-eddy simulation are carried out to examine its drag-reduction performance. Extensive parametric studies are performed experimentally by varying the height and width of the tab and the spanwise spacing between the adjacent tabs at three Reynolds numbers of Re {=} u_infty h/nu {=} 20 000, 40 000 and 80 000, where u_infty is the free-stream velocity and h is the body height. For a wide parameter range, the base pressure increases (i.e. drag reduces) at all three Reynolds numbers. Furthermore, a significant increase in the base pressure by more than 30% is obtained for the optimum tab configuration. Numerical simulations are performed at much lower Reynolds numbers of Re {=} 320 and 4200 to investigate the mechanism responsible for the base-pressure increase by the tab. Results from the velocity measurement and numerical simulations show that the tab introduces the spanwise mismatch in the vortex-shedding process, resulting in a substantial reduction of the vortical strength in the wake and significant increases in the vortex formation length and wake width.

  8. The Effect of Blunt-Trailing-Edge Modifications on the High-Speed Stability and Control Characteristics of a Swept-Wing Fighter Airplane

    NASA Technical Reports Server (NTRS)

    Sadoff, Melvin; Matteson, Frederick H.; Van Dyke, Rudolph D., Jr.

    1954-01-01

    An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

  9. Low-speed aerodynamic characteristics of a wing-canard configuration with underwing spanwise blowing on the trailing-edge flap system

    NASA Technical Reports Server (NTRS)

    Banks, Daniel W.; Paulson, John W., Jr.

    1987-01-01

    An investigation of the effects of spanwise blowing applied to the lower surface of a trailing-edge flap system on a wing-canard configuration has been conducted in the Langley 4- by 7-Meter Tunnel. The investigation studied spanwise-blowing angles of 30 deg., 45 deg., and 60 deg. measured from a perpendicular to the body center-line. The test conditions covered a range of free-stream dynamic pressures up to 50 psf for thrust coefficients up to 2.1 over a range of angles of attack from -2 deg. to 26 deg. Model height above the wind tunnel floor was varied from a height-to-span ratio of 1.70 down to 0.20 (a representative wheel touchdown height). The results indicate that blowing angles of 30 deg. and 45 deg. increase the induced-lift increment produced by spanwise blowing on the lower surface of a trailing-edge flap system. Increasing the blowing angle to 60 deg., in general, produces little further improvement.

  10. Study of viscous flow about airfoils by the integro-differential method

    NASA Technical Reports Server (NTRS)

    Wu, J. C.; Sampath, S.

    1975-01-01

    An integro-differential method was used for numerically solving unsteady incompressible viscous flow problems. A computer program was prepared to solve the problem of an impulsively started 9% thick symmetric Joukowski airfoil at an angle of attack of 15 deg and a Reynolds number of 1000. Some of the results obtained for this problem were discussed and compared with related work completed previously. Two numerical procedures were used, an Alternating Direction Implicit (ADI) method and a Successive Line Relaxation (SLR) method. Generally, the ADI solution agrees well with the SLR solution and with previous results are stations away from the trailing edge. At the trailing edge station, the ADI solution differs substantially from previous results, while the vorticity profiles obtained from the SLR method there are in good qualitative agreement with previous results.

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

    NASA Astrophysics Data System (ADS)

    Kearney, John; Glezer, Ari

    2014-11-01

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

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

    NASA Astrophysics Data System (ADS)

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

    2011-12-01

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

  13. Micro-abrasion package capture cell experiment on the trailing edge of LDEF: Impactor chemistry and whipple bumper shield efficiencies

    NASA Technical Reports Server (NTRS)

    Fitzgerald, Howard J.; Yano, Hajime

    1995-01-01

    Four of the eight available double layer microparticle capture cells, flown as the experiment A0023 on the trailing (West) face of LDEF, have been extensively studied. An investigation of the chemistry of impactors has been made using SEM/EDX techniques and the effectiveness of the capture cells as bumper shields has also been examined. Studies of these capture cells gave positive EDX results, with 53 percent of impact sites indicating the presence of some chemical residues, the predominant residue identified as being silicon in varying quantities.

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

    NASA Astrophysics Data System (ADS)

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

    2014-12-01

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

  15. Flight Measurements of the Lateral Control Characteristics of Narrow-Chord Ailerons on the Trailing Edge of a Full-Span Slotted Flap

    NASA Technical Reports Server (NTRS)

    Sawyer, Richard H

    1947-01-01

    Results are presented of light tests made to determine the effect of flap deflection on the lateral control characteristics of a modified brewster f2a-2 airplane equipped with partial-span narrow-chord ailerons on the trailing edge of a full-span NACA slotted flap. The investigation included determination of the rolling and yawing characteristics of the airplane in abrupt aileron rolls with the slotted flap at various settings ranging from 0 degree to about 40 degrees. The results showed that the effectiveness of the ailerons was greatly reduced at flap deflections greater than about 20 degrees. For flap deflections up to about 20 degrees, the aileron effectiveness was about the same as with flaps retracted, but the adverse yawing velocity developed in the abrupt aileron rolls was somewhat increased.

  16. Blowing-Type Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

    NASA Technical Reports Server (NTRS)

    Kelly, Mark W; Anderson, Seth B; Innis, Robert C

    1958-01-01

    A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.

  17. Persistent source influences on the trailing edge of a groundwater plume, and natural attenuation timeframes: the F-Area Savannah River Site.

    PubMed

    Wan, Jiamin; Tokunaga, Tetsu K; Dong, Wenming; Denham, Miles E; Hubbard, Susan S

    2012-04-17

    At the Savannah River Site's F-Area, wastewaters containing radionuclides were disposed into seepage basins for decades. After closure and capping in 1991, the U.S. Department of Energy (DOE) has being monitoring and remediating the groundwater plume. Despite numerous studies of the plume, its persistence for over 20 years has not been well understood. To better understand the plume dynamics, a limited number of deep boreholes were drilled to determine the current plume characteristics. A mixing model was developed to predict plume tritium and nitrate concentrations. We found that the plume trailing edges have emerged for some contaminants, and that contaminant recharge from the basin's vadose zone is still important. The model's estimated time-dependent basin drainage rates combined with dilution from natural recharge successfully predicted plume tritium and nitrate concentrations. This new understanding of source zone influences can help guide science-based remediation, and improve predictions of the natural attenuation timeframes.

  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. Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 3: Data and performance for stage C

    NASA Technical Reports Server (NTRS)

    Brent, J. A.; Clemmons, D.

    1972-01-01

    Stage C, comprised of tandem-airfoil rotor C and tandem-airfoil stator B, was designed and tested to establish performance data for comparison with the performance of conventional single-airfoil blading. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. At design equivalent rotor speed, rotor C achieved a maximum adiabatic efficiency of 91.8% at a pressure ratio of 1.31. The stage maximum adiabatic efficiency was 86.5% at a pressure ratio of 1.31.

  20. Numerical Investigation of the Control of Separation from Curved and Blunt Trailing Edges Using DNS and LES

    NASA Astrophysics Data System (ADS)

    Fasel, Hermann F.

    2002-07-01

    Wall jets over a curved wall geometry (Coanda flows) are investigated using DNS and turbulence modeling. In experiments large coherent structures have enhanced the effectiveness of wall jets in delaying or preventing flow separation on airfoils. Understanding the behavior of these structures is essential for utilizing wall jets for separation control. The research objective is to investigate curvature effects on large coherent structures, in particular the development of longitudinal (Goertler-type) vortices and their interaction with 2D vortices. The focus is on Coanda cylinders using two computational approaches. With the Flow Simulation Methodology (FSM), a turbulent wall jet is computed over a cylinder segment on a body-fitted grid. In FSM, the contribution of the turbulence model depends on the grid resolution relative to a local turbulent length scale. For a flat-plate reference case, FSM is employed as DNS, LES, and URANS. In all cases the large 2D vortices are captured. For the curved-wall geometry, FSM is employed as a DNS. Goertler-type vortices emerge in the simulation but remain weak due to the narrow computational domain. In the second approach, Coanda flows including nozzle and separated region are computed using immersed boundary techniques (IBT). The feasibility of IBT for Coanda Flows is established.

  1. 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.

  2. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

  6. Supercritical flow about a thick circular-arc airfoil

    NASA Technical Reports Server (NTRS)

    Mcdevitt, J. B.

    1979-01-01

    The supercritical flow about a biconvex circular-arc airfoil is being thoroughly documented at Ames Research Center in order to provide experimental test cases suitable for guiding and evaluating current and future computer codes. The effects of angle of attack, effects of leading and trailing-edge splitter plates, additional unsteady pressure fluctuation (buffeting) measurements and glow-field shadowgraphs, and application of an oil-film technique to display separated-wake streamlines were studied. Computed and measured pressure distributions for steady and unsteady flows, using a recent computer code representative of current methodology, are compared. It was found that the numerical solutions are often fundamentally incorrect in that only strong (shock-polar terminology) shocks are captured, whereas experimentally, both strong and weak shock waves appear.

  7. Recent Turbulence Model Advances Applied to Multielement Airfoil Computations

    NASA Technical Reports Server (NTRS)

    Rumsey, Christopher L.; Gatski, Thomas B.

    2000-01-01

    A one-equation linear turbulence model and a two-equation nonlinear explicit algebraic stress model (EASM) are applied to the flow over a multielement airfoil. The effect of the K-epsilon and K-omega forms of the two-equation model are explored, and the K-epsilon form is shown to be deficient in the wall-bounded regions of adverse pressure gradient flows. A new K-omega form of EASM is introduced. Nonlinear terms present in EASM are shown to improve predictions of turbulent shear stress behind the trailing edge of the main element and near midflap. Curvature corrections are applied to both the one- and two-equation turbulence models and yield only relatively small local differences in the flap region, where the flow field undergoes the greatest curvature. Predictions of maximum lift are essentially unaffected by the turbulence model variations studied.

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

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Sewall, William G.

    1995-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Chen, Xiaomin

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

  10. Characterization of dynamic stall on 9-15 % thick airfoils using experiment and computation

    NASA Astrophysics Data System (ADS)

    Davidson, Phillip B.

    In recent years, the blade geometry on wind turbines and helicopters has been optimized for a particular span location. Unsteady flow phenomena like dynamic stall limit these designs and need to be better understood and correctly simulated. Currently, empirical and computational fluid dynamics (CFD) methods are used to simulate rotating wind turbine or helicopter blades, but each of these methods has limitations in predicting unsteady separated flows. To address these needs, the present work investigated oscillating airfoils over a range of conditions with an approach that provided fast, low-cost unsteady pressure data combined with a highly resolved flow field to better understand the physics of dynamic stall. An additional objective was to show how such data may be used to assess CFD simulations. This research has yielded interesting results showing characteristics of thin airfoil stall, leading edge stall, and trailing edge stall that were sorted and classified. Classification of the oscillating airfoil behavior with or without dynamic stall was performed using previous definitions for stall regime, separation characteristics, and other qualitative differences in stall pattern. After classifying the unsteady flow for each of the cases, comparison of experimental results and results obtained using an unsteady Reynolds Averaged Navier-Stokes (URANS) solver was performed to assess the ability of the solver to produce the same unsteady effects. Although both experiment and computation produced similar flow features, the timing and magnitude of the features in the dynamic stall and re-attachment process of the pitching cycle exhibited some significant differences.

  11. 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.

  12. 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.

  13. 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.

  14. Wind tunnel results for a high-speed, natural laminar-flow airfoil designed for general aviation aircraft

    NASA Technical Reports Server (NTRS)

    Sewall, William G.; Mcghee, Robert J.; Viken, Jeffery K.; Waggoner, Edgar G.; Walker, Betty S.; Millard, Betty F.

    1985-01-01

    Two dimensional wind tunnel tests were conducted on a high speed natural laminar flow airfoil in both the Langley 6 x 28 inch Transonic Tunnel and the Langley Low Turbulence Pressure Tunnel. The test conditions consisted of Mach numbers ranging from 0.10 to 0.77 and Reynolds numbers ranging from 3 x 1 million to 11 x 1 million. The airfoil was designed for a lift coefficient of 0.20 at a Mach number of 0.70 and Reynolds number of 11 x 1 million. At these conditions, laminar flow would extend back to 50 percent chord of the upper surface and 70 percent chord of the lower surface. Low speed results were also obtained with a 0.20 chord trailing edge split flap deflected 60 deg.

  15. Experimental and numerical investigations of internal heat transfer in an innovative trailing edge blade cooling system: stationary and rotation effects, part 1—experimental results

    NASA Astrophysics Data System (ADS)

    Beniaiche, Ahmed; Ghenaiet, Adel; Facchini, Bruno

    2016-05-01

    The aero-thermal behavior of the flow field inside 30:1 scaled model reproducing an innovative smooth trailing edge of shaped wedge discharge duct with one row of enlarged pedestals have been investigated in order to determine the effect of rotation, inlet velocity and blowing conditions effects, for Re = 20,000 and 40,000 and Ro = 0-0.23. Two configurations are presented: with and without open tip configurations. Thermo-chromic liquid crystals technique is used to ensure a local measurement of the heat transfer coefficient on the blade suction side under stationary and rotation conditions. Results are reported in terms of detailed 2D HTC maps on the suction side surface as well as the averaged Nusselt number inside the pedestal ducts. Two correlations are proposed, for both closed and open tip configurations, based on the Re, Pr, Ro and a new non-dimensional parameter based on the position along the radial distance, to assess a reliable estimation of the averaged Nusselt number at the inter-pedestal region. A good agreement is found between prediction and experimental data with about ±10 to ±12 % of uncertainty, for the simple form correlation, and about ±16 % using a complex form. The obtained results help to predict the flow field visualization and the evaluation of the aero-thermal performance of the studied blade cooling system during the design step.

  16. Coupled Vortex-Lattice Flight Dynamic Model with Aeroelastic Finite-Element Model of Flexible Wing Transport Aircraft with Variable Camber Continuous Trailing Edge Flap for Drag Reduction

    NASA Technical Reports Server (NTRS)

    Nguyen, Nhan; Ting, Eric; Nguyen, Daniel; Dao, Tung; Trinh, Khanh

    2013-01-01

    This paper presents a coupled vortex-lattice flight dynamic model with an aeroelastic finite-element model to predict dynamic characteristics of a flexible wing transport aircraft. The aircraft model is based on NASA Generic Transport Model (GTM) with representative mass and stiffness properties to achieve a wing tip deflection about twice that of a conventional transport aircraft (10% versus 5%). This flexible wing transport aircraft is referred to as an Elastically Shaped Aircraft Concept (ESAC) which is equipped with a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for active wing shaping control for drag reduction. A vortex-lattice aerodynamic model of the ESAC is developed and is coupled with an aeroelastic finite-element model via an automated geometry modeler. This coupled model is used to compute static and dynamic aeroelastic solutions. The deflection information from the finite-element model and the vortex-lattice model is used to compute unsteady contributions to the aerodynamic force and moment coefficients. A coupled aeroelastic-longitudinal flight dynamic model is developed by coupling the finite-element model with the rigid-body flight dynamic model of the GTM.

  17. Experimental and numerical investigations of internal heat transfer in an innovative trailing edge blade cooling system: stationary and rotation effects, part 2: numerical results

    NASA Astrophysics Data System (ADS)

    Beniaiche, Ahmed; Ghenaiet, Adel; Carcasci, Carlo; Facchini, Bruno

    2016-05-01

    This paper presents a numerical validation of the aero-thermal study of a 30:1 scaled model reproducing an innovative trailing edge with one row of enlarged pedestals under stationary and rotating conditions. A CFD analysis was performed by means of commercial ANSYS-Fluent modeling the isothermal air flow and using k-ω SST turbulence model and an isothermal air flow for both static and rotating conditions (Ro up to 0.23). The used numerical model is validated first by comparing the numerical velocity profiles distribution results to those obtained experimentally by means of PIV technique for Re = 20,000 and Ro = 0-0.23. The second validation is based on the comparison of the numerical results of the 2D HTC maps over the heated plate to those of TLC experimental data, for a smooth surface for a Reynolds number = 20,000 and 40,000 and Ro = 0-0.23. Two-tip conditions were considered: open tip and closed tip conditions. Results of the average Nusselt number inside the pedestal ducts region are presented too. The obtained results help to predict the flow field visualization and the evaluation of the aero-thermal performance of the studied blade cooling system during the design step.

  18. Airfoil Tonal Noise Generation in Resonant Environments

    NASA Astrophysics Data System (ADS)

    Atobe, Takashi; Tuinstra, Marthijn; Takagi, Shohei

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

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

    NASA Technical Reports Server (NTRS)

    Wigley, D. A.

    1985-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

    2011-11-01

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

  1. The Determination of the Geometries of Multiple-Element Airfoils Optimized for Maximum Lift Coefficient. Ph.D. Thesis - Illinois Univ., Urbana

    NASA Technical Reports Server (NTRS)

    Chen, A. W.

    1971-01-01

    Optimum airfoils in the sense of maximum lift coefficient are obtained by a newly developed method. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution consists of a uniform stagnation pressure on the lower surface, a uniform minimum pressure on the upper surface immediately downstream of the front stagnation point followed by a Stratford zero skin friction pressure rise. When multiple-element airfoils are under consideration, this optimum pressure distribution appears on every element. The parameters used to specify the pressure distribution on each element are the Reynolds number and the normalized trailing edge velocity. The newly developed method of design computes the velocity distribution on a given airfoil and modifies the airfoil contour in a systematic manner until the desired velocity distribution is achieved. There are no limitations on how many elements the airfoil to be designed can have.

  2. 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.

  3. 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.

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

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  5. Development of high-rate large-deflection hingeless trailing-edge control surface for the Smart Wing wind tunnel model

    NASA Astrophysics Data System (ADS)

    Wang, Donny P.; Bartley-Cho, Jonathan D.; Martin, Christopher A.; Hallam, Brian J.

    2001-06-01

    A key objective of the Smart Wing Phase 2, Test 2 is to demonstrate high-rate actuation of hingeless control surfaces using smart material-based actuators. Actuation rates resulting in a minimum of 20 degree(s) flap deflection in 0.33 sec, producing a sweep rate of at least 60 degree(s)/sec, are desired. This sweep rate is similar to those specified for many of the existing military platforms with hinged control surfaces. The ability to deploy control surfaces without discrete hingeline would, however, enhance platform mission by reducing radar cross section and improving aerodynamic performance. Studies on numerous actuation concepts and flexible structures were executed during the early and mid phase of the program in an effort to satisfy these goals. In the first study, several actuation concepts with different transducers were modeled and analyzed. These concepts included distributed piezoelectric stack actuators with and without hydraulic amplifiers and pumps, antagonistic tendon actuation, and eccentuation. The transducers selected for the trade studies included piezoelectric ultrasonic motors, actively cooled SMA, ferromagnetic SMA, and stacks made from piezoelectric ceramic wafer, piezoelectric single crystal wafer, irradiated PVDF-TrFE film, and dielectric elastomer film. Although many of the technologies are not fully mature, they provide a glimpse of what improvements could be possible with their successful development. The studies showed that distributed polymer stacks provided the most elegant solution, but eccentuation was deemed the most realistic and lowest risk approach to attaining the program goals. A common issue to all the concepts was the structural stiffness that the actuators worked against. This was resolved in the second study by developing a flexcore- elastomeric skin trailing edge structure with eccentuation using high power ultrasonic motors. This paper describes the two studies and the final concept in detail.

  6. The impact of geometric non-linearities on the fatigue analysis of trailing edge bond lines in wind turbine rotor blades

    NASA Astrophysics Data System (ADS)

    Noever Castelos, Pablo; Balzani, Claudio

    2016-09-01

    The accurate prediction of stress histories for the fatigue analysis is of utmost importance for the design process of wind turbine rotor blades. As detailed, transient, and geometrically non-linear three-dimensional finite element analyses are computationally weigh too expensive, it is commonly regarded sufficient to calculate the stresses with a geometrically linear analysis and superimpose different stress states in order to obtain the complete stress histories. In order to quantify the error from geometrically linear simulations for the calculation of stress histories and to verify the practical applicability of the superposition principal in fatigue analyses, this paper studies the influence of geometric non-linearity in the example of a trailing edge bond line, as this subcomponent suffers from high strains in span-wise direction. The blade under consideration is that of the IWES IWT-7.5-164 reference wind turbine. From turbine simulations the highest edgewise loading scenario from the fatigue load cases is used as the reference. A 3D finite element model of the blade is created and the bond line fatigue assessment is performed according to the GL certification guidelines in its 2010 edition, and in comparison to the latest DNV GL standard from end of 2015. The results show a significant difference between the geometrically linear and non-linear stress analyses when the bending moments are approximated via a corresponding external loading, especially in case of the 2010 GL certification guidelines. This finding emphasizes the demand to reconsider the application of the superposition principal in fatigue analyses of modern flexible rotor blades, where geometrical nonlinearities become significant. In addition, a new load application methodology is introduced that reduces the geometrically non-linear behaviour of the blade in the finite element analysis.

  7. Initial Assessment of a Variable-Camber Continuous Trailing-Edge Flap System on a Rigid Wing for Drag Reduction in Subsonic Cruise

    NASA Technical Reports Server (NTRS)

    Ippolito, Corey; Nguyen, Nhan; Totah, Joe; Trinh, Khanh; Ting, Eric

    2013-01-01

    In this paper, we describe an initial optimization study of a Variable-Camber Continuous Trailing-Edge Flap (VCCTEF) system. The VCCTEF provides a light-weight control system for aircraft with long flexible wings, providing efficient high-lift capability for takeoff and landing, and greater efficiency with reduced drag at cruising flight by considering the effects of aeroelastic wing deformations in the control law. The VCCTEF system is comprised of a large number of distributed and individually-actuatable control surfaces that are constrained in movement relative to neighboring surfaces, and are non-trivially coupled through structural aeroelastic dynamics. Minimzation of drag results in a constrained, coupled, non-linear optimization over a high-dimension search space. In this paper, we describe the modeling, analysis, and optimization of the VCCTEF system control inputs for minimum drag in cruise. The purpose of this initial study is to quantify the expected benefits of the system concept. The scope of this analysis is limited to consideration of a rigid wing without structural flexibility in a steady-state cruise condition at various fuel weights. For analysis, we developed an optimization engine that couples geometric synthesis with vortex-lattice analysis to automate the optimization procedure. In this paper, we present and describe the VCCTEF system concept, optimization approach and tools, run-time performance, and results of the optimization at 20%, 50%, and 80% fuel load. This initial limited-scope study finds the VCCTEF system can potentially gain nearly 10% reduction in cruise drag, provides greater drag savings at lower operating weight, and efficiency is negatively impacted by the severity of relative constraints between control surfaces.

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  9. Dynamic stall simulation of a pitching airfoil under unsteady freestream velocity

    NASA Astrophysics Data System (ADS)

    Gharali, Kobra; Johnson, David A.

    2013-10-01

    Effects of horizontal oscillations of the freestream velocity superimposed on a pitch oscillating NACA0012 airfoil were investigated using Computational Fluid Dynamics (CFD). The SST k-ω model coupled with a low-Reynolds number correction was applied for Re ≈105, when the airfoil was undergoing dynamic stall. The main parameter Φ, the phase difference between the freestream oscillation and the airfoil oscillation, was varied from 0 to π. The Φ variation resulted in several times amplitude dynamic loads when Φ≤π/2 to several times damping dynamic loads for Φ>π/2 where some dynamic stall loads were damped even below static stall load values. It was found that Φ variation was divided into two main ranges based on the values of the unsteady freestream velocity at dynamic stall. The load variation also appeared with some differences including the shape of the trailing edge vortex sheet before stall, the circulation of the dynamic stall vortex pairs, the critical angles, vortex growth time and the secondary lift peak location that are discussed in detail.

  10. 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).

  11. 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.

  12. Effect of Compressibility on the Pressure and Forces Acting on a Modified NACA 65,3-019 Airfoil Having A 0.20-Chord Flap

    NASA Technical Reports Server (NTRS)

    Lindsey, W F

    1946-01-01

    An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value.

  13. Influence of airfoil thickness on convected gust interaction noise

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  14. 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.

  15. 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.

  16. New airfoils for small horizontal axis wind turbines

    SciTech Connect

    Giguere, P.; Selig, M.S.

    1997-12-31

    In a continuing effort to enhance the performance of small energy systems, one root airfoil and three primary airfoils were specifically designed for small horizontal axis wind turbines. These airfoils are intended primarily for 1-10 kW variable-speed wind turbines for both conventional (tapered/twisted) or pultruded blades. The four airfoils were wind-tunnel tested at Reynolds numbers between 100,000 and 500,000. Tests with simulated leading-edge roughness were also conducted. The results indicate that small variable-speed wind turbines should benefit from the use of the new airfoils which provide enhanced lift-to-drag ratio performance as compared with previously existing airfoils.

  17. 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.

  18. 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.

  19. 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.

  20. The construction and operation of a water tunnel in application to flow visualization studies of an oscillating airfoil

    NASA Technical Reports Server (NTRS)

    Olsen, J. H.; Liu, H. T.

    1973-01-01

    The water tunnel which was constructed at the NASA Ames Research Center is described along with the flow field adjacent to an oscillating airfoil. The design and operational procedures of the tunnel are described in detail. Hydrogen bubble and thymol blue techniques are used to visualize the flow field. Results of the flow visualizations are presented in a series of still pictures and a high speed movie. These results show that time stall is more complicated than simple shedding from the leading edge or the trailing edge, particularly at relatively low frequency oscillations comparable to those of a helicopter blade. Therefore, any successful theory for predicting the stall loads on the helicopter blades must treat an irregular separated region rather than a discrete vortex passing over each blade surface.

  1. Comparative Study of Airfoil Flow Separation Criteria

    NASA Astrophysics Data System (ADS)

    Laws, Nick; Kahouli, Waad; Epps, Brenden

    2015-11-01

    Airfoil flow separation impacts a multitude of applications including turbomachinery, wind turbines, and bio-inspired micro-aerial vehicles. In order to achieve maximum performance, some devices operate near the edge of flow separation, and others use dynamic flow separation advantageously. Numerous criteria exist for predicting the onset of airfoil flow separation. This talk presents a comparative study of a number of such criteria, with emphasis paid to speed and accuracy of the calculations. We evaluate the criteria using a two-dimensional unsteady vortex lattice method, which allows for rapid analysis (on the order of seconds instead of days for a full Navier-Stokes solution) and design of optimal airfoil geometry and kinematics. Furthermore, dynamic analyses permit evaluation of dynamic stall conditions for enhanced lift via leading edge vortex shedding, commonly present in small flapping-wing flyers such as the bumblebee and hummingbird.

  2. Low-speed investigation of effects of wing leading- and trailing-edge flap deflections and canard incidence on a fighter configuration equipped with a forward-swept wing

    NASA Technical Reports Server (NTRS)

    Gainer, T. G.; Mann, M. J.; Huffman, J. K.

    1984-01-01

    An advanced fighter configuration with a forward-swept wing of aspect ratio 3.28 is tested in the Langley 7 by 10 Foot High Speed Tunnel at a Mach number of 0.3. The wing has 29.5 degrees of forward sweep of the quarter chord line and is equipped with 15 percent chord leading edge and 30 percent chord trailing edge flaps. The canard is sweptback 45 degrees. Tests were made through a range of angle of attack from about -2 degrees to 22 degrees. Deflecting the flaps significantly improves the lift drag characteristics at the higher angles of attack. The canard is able to trim the configurations with different flap deflections over most of the range of angle of attack. The penalty in maximum lift coefficient due to trimming is about 0.10.

  3. Effects of Airfoil Thickness and Maximum Lift Coefficient on Roughness Sensitivity: 1997--1998

    SciTech Connect

    Somers, D. M.

    2005-01-01

    A matrix of airfoils has been developed to determine the effects of airfoil thickness and the maximum lift to leading-edge roughness. The matrix consists of three natural-laminar-flow airfoils, the S901, S902, and S903, for wind turbine applications. The airfoils have been designed and analyzed theoretically and verified experimentally in the Pennsylvania State University low-speed, low-turbulence wind tunnel. The effect of roughness on the maximum life increases with increasing airfoil thickness and decreases slightly with increasing maximum lift. Comparisons of the theoretical and experimental results generally show good agreement.

  4. Calculation of the rotor induced download on airfoils

    NASA Technical Reports Server (NTRS)

    Lee, C. S.

    1989-01-01

    Interactions between the rotors and wing of a rotary wing aircraft in hover have a significant detrimental effect on its payload performance. The reduction of payload results from the wake of lifting rotors impinging on the wing, which is at 90 deg angle of attack in hover. This vertical drag, often referred as download, can be as large as 15 percent of the total rotor thrust in hover. The rotor wake is a three-dimensional, unsteady flow with concentrated tip vortices. With the rotor tip vortices impinging on the upper surface of the wing, the flow over the wing is not only three-dimensional and unsteady, but also separated from the leading and trailing edges. A simplified two-dimensional model was developed to demonstrate the stability of the methodology. The flow model combines a panel method to represent the rotor and the wing, and a vortex method to track the wing wake. A parametric study of the download on a 20 percent thick elliptical airfoil below a rotor disk of uniform inflow was performed. Comparisons with experimental data are made where the data are available. This approach is now being extended to three-dimensional flows. Preliminary results on a wing at 90 deg angle of attack in free stream is presented.

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

    NASA Technical Reports Server (NTRS)

    Volino, Ralph J.; Hultgren, Lennart .

    2000-01-01

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

  6. Snail Trails

    ERIC Educational Resources Information Center

    Galus, Pamela

    2002-01-01

    The slime trails of snails lead the author's students to a better understanding of science as inquiry and the processes of science. During this five-day activity, students get up close and personal with one of her favorite creatures, the land snail. Students begin by observing the organism and recording their observations. After making initial…

  7. 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.

  8. Numerical simulations of iced airfoils and wings

    NASA Astrophysics Data System (ADS)

    Pan, Jianping

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

  9. Slat tonal noise mechanisms in a two-dimensional multi-element airfoil configuration

    NASA Astrophysics Data System (ADS)

    Olson, Susan Denise

    Airframe noise from commercial aircraft rivals engine noise in the landing phase of flight. Unsteady flow around the leading edge slat of the high-lift system contributes significantly to airframe noise. This research experimentally investigated tonal noise sources arising from the unsteady flow around the leading edge slat from a flow physics standpoint. Using this approach, efficient acoustic control strategies naturally follow. This study resolved two types of acoustic tones, trailing edge vortex shedding noise and acoustic resonance, originating from the slat component of a two-dimensional, three-element airfoil model. The latter noise mechanism had not been documented until the current study. Sharpening the slat trailing edge reduced the overall sound pressure level from the shedding noise by 12dB. The reduction in radiated acoustic pressure was directly proportional to the reduction in the strength of the vortex shedding. The acoustic power was conceptually related to the strength of the global instability in the wake. The onset of the slat aeroacoustic resonance corresponded to the formation of a separation bubble on the upper surface of the slat. The peak acoustic frequency matched the frequency of large-scale structures ejected from the separation bubble. A feedback loop between the near-wake region and the separation point sustained the resonance. A model based on a phase matching condition between the acoustic waves and the generation of the separated shear layer instability waves successfully predicted the acoustic frequencies. The acoustic tone followed the modal variation closest to the subharmonic of the natural instability frequency, which caused the discontinuous jumps in frequency with increasing velocity.

  10. Vortex flow structures and interactions for the optimum thrust efficiency of a heaving airfoil at different mean angles of attack

    SciTech Connect

    Martín-Alcántara, A.; Fernandez-Feria, R.

    2015-07-15

    The thrust efficiency of a two-dimensional heaving airfoil is studied computationally for a low Reynolds number using a vortex force decomposition. The auxiliary potentials that separate the total vortex force into lift and drag (or thrust) are obtained analytically by using an elliptic airfoil. With these auxiliary potentials, the added-mass components of the lift and drag (or thrust) coefficients are also obtained analytically for any heaving motion of the airfoil and for any value of the mean angle of attack α. The contributions of the leading- and trailing-edge vortices to the thrust during their down- and up-stroke evolutions are computed quantitatively with this formulation for different dimensionless frequencies and heave amplitudes (St{sub c} and St{sub a}) and for several values of α. Very different types of flows, periodic, quasi-periodic, and chaotic described as St{sub c}, St{sub a}, and α, are varied. The optimum values of these parameters for maximum thrust efficiency are obtained and explained in terms of the interactions between the vortices and the forces exerted by them on the airfoil. As in previous numerical and experimental studies on flapping flight at low Reynolds numbers, the optimum thrust efficiency is reached for intermediate frequencies (St{sub c} slightly smaller than one) and a heave amplitude corresponding to an advance ratio close to unity. The optimal mean angle of attack found is zero. The corresponding flow is periodic, but it becomes chaotic and with smaller average thrust efficiency as |α| becomes slightly different from zero.

  11. Transient phenomena of shock-induced turbulent separation for a spikebody and stalling airfoil at transonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Yoshikawa, K. K.

    1982-01-01

    The time-dependent, compressible, Reynolds-averaged, full Navier-Stokes equations are applied to solve an axisymmetric flow around a forward-facing stepbody (spikebody) at supersonic speeds and a stalling airfoil at transonic speeds. Important transient and unsteady phenomena, not yet well understood, are examined, and significant new findings of the present solution to the phenomena are discussed. The phenomena described in detail are as follows: The evolution of the shock wave pressure built up by the impact of the pressure waves, one from the trailing edge; the separation of the flow as influenced by the shock wave; the location of the reversed flow, the separation point, and the reattachment point; and the transient (or unsteady) phenomena of the flow pulsation, oscillation, and stalling of the body and airfoil wake flow. The numerical results show that the transient flow instability is caused by a supersonic jet induced in the separation bubble by the shock-bifurcation (lambda shock) mechanism between the separation shock and the reflected shock. Pulsation and stall phenomena are caused by a sudden increase in the leading-edge pressure due to the jet and the separation bubble interacting along the stagnation point flow.

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

    NASA Technical Reports Server (NTRS)

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

    1980-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Hui, W. H.; Tobak, M.

    1981-01-01

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

  14. Predictions of airfoil aerodynamic performance degradation due to icing

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Zaguli, R. J.

    1984-01-01

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

  16. Control of Flow Separation Using Adaptive Airfoils

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

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

  17. Experiments on a separation bubble over an Eppler 387 airfoil at low Reynolds numbers using thin-film arrays

    NASA Technical Reports Server (NTRS)

    Mousseux, Marc C.; Chokani, Ndaona; Stack, John P.; Mcghee, Robert J.

    1990-01-01

    Experiments carried out at the NASA Langley Low-Turbulence Pressure Tunnel on an Eppler 387 airfoil at several angles of attack for Reynolds numbers of 100,000 to 300,000 are discussed. Cross-correlations and phase diagrams are used in calculating velocities, while spectra are used in checking for amplified frequencies, and coherence plots for causality. Hot-film measurements, oil-flow visualization, pressure measurements, and Drela code are compared for finding the leading and trailing edges of the bubble. It is observed that all methods agree in showing the beginning and end of the separation bubble, the velocity measurements agree in magnitude for the lower velocities, and wave packets appear in the cross-correlation.

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

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Craig, Anthony P.; Hansman, R. John

    1987-01-01

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

  1. 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.

  2. The SNL100-03 Blade: Design Studies with Flatback Airfoils for the Sandia 100-meter Blade.

    SciTech Connect

    Griffith, Daniel; Richards, Phillip William

    2014-09-01

    A series of design studies were performed to inv estigate the effects of flatback airfoils on blade performance and weight for large blades using the Sandi a 100-meter blade designs as a starting point. As part of the study, the effects of varying the blade slenderness on blade structural performance was investigated. The advantages and disadvantages of blad e slenderness with respect to tip deflection, flap- wise & edge-wise fatigue resistance, panel buckling capacity, flutter speed, manufacturing labor content, blade total weight, and aerodynamic design load magn itude are quantified. Following these design studies, a final blade design (SNL100-03) was prod uced, which was based on a highly slender design using flatback airfoils. The SNL100-03 design with flatback airfoils has weight of 49 tons, which is about 16% decrease from its SNL100-02 predecessor that used conventional sharp trailing edge airfoils. Although not systematically optimized, the SNL100 -03 design study provides an assessment of and insight into the benefits of flatback airfoils for la rge blades as well as insights into the limits or negative consequences of high blade slenderness resulting from a highly slender SNL100-03 planform as was chosen in the final design definition. This docum ent also provides a description of the final SNL100-03 design definition and is intended to be a companion document to the distribution of the NuMAD blade model files for SNL100-03, which are made publicly available. A summary of the major findings of the Sandia 100-meter blade development program, from the initial SNL100-00 baseline blade through the fourth SNL100-03 blade study, is provided. This summary includes the major findings and outcomes of blade d esign studies, pathways to mitigate the identified large blade design drivers, and tool development that were produced over the course of this five-year research program. A summary of large blade tec hnology needs and research opportunities is also presented.

  3. Closed loop steam cooled airfoil

    DOEpatents

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

    2006-04-18

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

  4. Transonic Airfoil Development

    NASA Technical Reports Server (NTRS)

    Whitcomb, R. T.

    1983-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Hajian, Rozhin; Jaworski, Justin W.

    2015-11-01

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

  6. Active Control of Flow Separation Over an Airfoil

    NASA Technical Reports Server (NTRS)

    Ravindran, S. S.

    1999-01-01

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

  7. Tracking of raindrops in flow over an airfoil

    NASA Technical Reports Server (NTRS)

    Valentine, James R.; Decker, Rand

    1993-01-01

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

  8. Shift happens: trailing edge contraction associated with recent warming trends threatens a distinct genetic lineage in the marine macroalga Fucus vesiculosus

    PubMed Central

    2013-01-01

    Background Significant effects of recent global climate change have already been observed in a variety of ecosystems, with evidence for shifts in species ranges, but rarely have such consequences been related to the changes in the species genetic pool. The stretch of Atlantic coast between North Africa and North Iberia is ideal for studying the relationship between species distribution and climate change as it includes the distributional limits of a considerable number of both cold- and warm-water species. We compared temporal changes in distribution of the canopy-forming alga Fucus vesiculosus with historical sea surface temperature (SST) patterns to draw links between range shifts and contemporary climate change. Moreover, we genetically characterized with microsatellite markers previously sampled extinct and extant populations in order to estimate resulting cryptic genetic erosion. Results Over the past 30 years, a geographic contraction of the southern range edge of this species has occurred, with a northward latitudinal shift of approximately 1,250 km. Additionally, a more restricted distributional decline was recorded in the Bay of Biscay. Coastal SST warming data over the last three decades revealed a significant increase in temperature along most of the studied coastline, averaging 0.214°C/decade. Importantly, the analysis of existing and extinct population samples clearly distinguished two genetically different groups, a northern and a southern clade. Because of the range contraction, the southern group is currently represented by very few extant populations. This southern edge range shift is thus causing the loss of a distinct component of the species genetic background. Conclusions We reveal a climate-correlated diversity loss below the species level, a process that could render the species more vulnerable to future environmental changes and affect its evolutionary potential. This is a remarkable case of genetic uniqueness of a vanishing cryptic genetic

  9. Steady inviscid transonic flows over planar airfoils: A search for a simplified procedure

    NASA Technical Reports Server (NTRS)

    Magnus, R.; Yoshihara, H.

    1973-01-01

    A finite difference procedure based upon a system of unsteady equations in proper conservation form with either exact or small disturbance steady terms is used to calculate the steady flows over several classes of airfoils. The airfoil condition is fulfilled on a slab whose upstream extremity is a semi-circle overlaying the airfoil leading edge circle. The limitations of the small disturbance equations are demonstrated in an extreme example of a blunt-nosed, aft-cambered airfoil. The necessity of using the equations in proper conservation form to capture the shock properly is stressed. Ability of the steady relaxation procedures to capture the shock is briefly examined.

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

    NASA Technical Reports Server (NTRS)

    Harris, Charles D.

    1990-01-01

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

  11. 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.

  12. 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.

  13. 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.

  14. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2014-12-01

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

  16. Ice Accretions on a Swept GLC-305 Airfoil

    NASA Technical Reports Server (NTRS)

    Vargas, Mario; Papadakis, Michael; Potapczuk, Mark; Addy, Harold; Sheldon, David; Giriunas, Julius

    2002-01-01

    An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to obtain castings of ice accretions formed on a 28 deg. swept GLC-305 airfoil that is representative of a modern business aircraft wing. Because of the complexity of the casting process, the airfoil was designed with three removable leading edges covering the whole span. Ice accretions were obtained at six icing conditions. After the ice was accreted, the leading edges were detached from the airfoil and moved to a cold room. Molds of the ice accretions were obtained, and from them, urethane castings were fabricated. This experiment is the icing test of a two-part experiment to study the aerodynamic effects of ice accretions.

  17. Control of Flow Separation Using Adaptive Airfoils

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

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

  18. 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.

  19. Experimental Measurements of Two-dimensional Planar Propagating Edge Flames

    NASA Technical Reports Server (NTRS)

    Villa-Gonzalez, Marcos; Marchese, Anthony J.; Easton, John W.; Miller, Fletcher J.

    2007-01-01

    The study of edge flames has received increased attention in recent years. This work reports the results of a recent study into two-dimensional, planar, propagating edge flames that are remote from solid surfaces (called here, free-layer flames, as opposed to layered flames along floors or ceilings). They represent an ideal case of a flame propagating down a flammable plume, or through a flammable layer in microgravity. The results were generated using a new apparatus in which a thin stream of gaseous fuel is injected into a low-speed laminar wind tunnel thereby forming a flammable layer along the centerline. An airfoil-shaped fuel dispenser downstream of the duct inlet issues ethane from a slot in the trailing edge. The air and ethane mix due to mass diffusion while flowing up towards the duct exit, forming a flammable layer with a steep lateral fuel concentration gradient and smaller axial fuel concentration gradient. We characterized the flow and fuel concentration fields in the duct using hot wire anemometer scans, flow visualization using smoke traces, and non-reacting, numerical modeling using COSMOSFloWorks. In the experiment, a hot wire near the exit ignites the ethane air layer, with the flame propagating downwards towards the fuel source. Reported here are tests with the air inlet velocity of 25 cm/s and ethane flows of 967-1299 sccm, which gave conditions ranging from lean to rich along the centerline. In these conditions the flame spreads at a constant rate faster than the laminar burning rate for a premixed ethane air mixture. The flame spread rate increases with increasing transverse fuel gradient (obtained by increasing the fuel flow rate), but appears to reach a maximum. The flow field shows little effect due to the flame approach near the igniter, but shows significant effect, including flow reversal, well ahead of the flame as it approaches the airfoil fuel source.

  20. Into the Woods: A 6th-Grade Nature Trail.

    ERIC Educational Resources Information Center

    Wilburn, Douglas

    1983-01-01

    Describes an ecology project in which sixth-grade students built and operated a nature trail on the edge of school property. Classes toured the trail and participated in grade-appropriate follow-up activities (e.g., art lessons and soil analysis activities). (RH)

  1. Numerical simulation and reduced-order modeling of a flapping airfoil

    NASA Astrophysics Data System (ADS)

    Lewin, Gregory Carl

    Recent advances in many fields have made the design of micro-aerial vehicles that implement flapping wings a possibility. However, there are many outstanding problems that must be solved before flapping flight can be implemented as a practical means of propulsion. This dissertation focuses on two important aspects of flapping flight: the physics of the flow of a fluid around a heaving airfoil and the development of a reduced-order model for the control of a flapping airfoil. To study the physics of the flow, a numerical model for two-dimensional flow around an airfoil undergoing prescribed oscillatory motions in a viscous flow is developed. The model is used to examine the flow characteristics and power coefficients of a symmetric airfoil heaving sinusoidally over a range of frequencies and amplitudes. Both periodic and aperiodic solutions are found. Additionally, some flows are asymmetric in that the up-stroke is not a mirror image of the down-stroke. For a given Strouhal number---defined as the product of dimensionless frequency and heave amplitude---the maximum efficiency occurs at an intermediate heaving frequency. This is in contrast to ideal flow models, in which efficiency increases monotonically as frequency decreases. Below a threshold frequency, the separation of the leading edge vortices early in each stroke reduces the force on the airfoil and leads to diminished thrust and efficiency. Above the optimum frequency, the efficiency decreases similarly to inviscid theory. For most cases, the efficiency can be correlated to interactions between leading and trailing edge vortices, with positive reinforcement leading to relatively high efficiency, and negative reinforcement leading to relatively low efficiency. Additionally, the efficiency is related to the proximity of the heaving frequency to the frequency of the most spatially unstable mode of the average velocity profile of the wake; the greatest efficiency occurs when the two frequencies are nearly

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

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1929-01-01

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

  3. The Pimlico Chemistry Trail.

    ERIC Educational Resources Information Center

    Borrows, Peter

    1984-01-01

    Describes a chemistry "trail" (similar to a nature trail) which focuses on chemical phenomena in the environment. The trail includes 20 stops in and around a local school. Types of phenomena examined include building materials, air pollution, corrosion of metals, swimming pools, and others. Additional activities are also suggested. (DH)

  4. 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.

  5. Planform curvature effects on flutter characteristics of a wing with 56 deg leading-edge sweep and panel aspect ratio of 1.14

    NASA Technical Reports Server (NTRS)

    Keller, Donald F.; Sandford, Maynard C.; Pinkerton, Theresa L.

    1991-01-01

    An experimental and analytical investigation was initiated to determine the effects of planform curvature (curving the leading and trailing edges of a wing in the X-Y plane) on the transonic flutter characteristics of a series of three moderately swept wing models. Experimental flutter results were obtained in the Langley Transonic Dynamics Tunnel for Mach numbers from 0.60-1.00, with air as the test medium. The models were semispan cantilevered wings with a 3 percent biconvex airfoil and a panel aspect ratio of 1.14. The baseline model had straight leading and trailing edges (i.e., no planform curvature). The radii of curvature of the leading edges for these two models were 200 and 80 inches. The radii of curvature of the leading edges of the other two models were determined so that the root and tip chords were identical for all three models. Experimental results showed that flutter-speed index and flutter frequency ratio increased as planform curvature increase (radius of curvature of the leading edge was decreased) over the test range of Mach numbers. Analytical flutter results were calculated with a subsonic flutter-prediction program, and they agreed well with the experimental results.

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

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  8. Time-resolved stereo PIV measurements of shock-boundary layer interaction on a supercritical airfoil

    NASA Astrophysics Data System (ADS)

    Hartmann, Axel; Klaas, Michael; Schröder, Wolfgang

    2012-03-01

    Time-resolved stereo particle-image velocimetry (TR-SPIV) and unsteady pressure measurements are used to analyze the unsteady flow over a supercritical DRA-2303 airfoil in transonic flow. The dynamic shock wave-boundary layer interaction is one of the most essential features of this unsteady flow causing a distinct oscillation of the flow field. Results from wind-tunnel experiments with a variation of the freestream Mach number at Reynolds numbers ranging from 2.55 to 2.79 × 106 are analyzed regarding the origin and nature of the unsteady shock-boundary layer interaction. Therefore, the TR-SPIV results are analyzed for three buffet flows. One flow exhibits a sinusoidal streamwise oscillation of the shock wave only due to an acoustic feedback loop formed by the shock wave and the trailing-edge noise. The other two buffet flows have been intentionally influenced by an artificial acoustic source installed downstream of the test section to investigate the behavior of the interaction to upstream-propagating disturbances generated by a defined source of noise. The results show that such upstream-propagating disturbances could be identified to be responsible for the upstream displacement of the shock wave and that the feedback loop is formed by a pulsating separation of the boundary layer dependent on the shock position and the sound pressure level at the shock position. Thereby, the pulsation of the separation could be determined to be a reaction to the shock motion and not vice versa.

  9. 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.

  10. 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.

  11. Influence of airfoil camber on convected gust interaction noise

    NASA Astrophysics Data System (ADS)

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

    1986-07-01

    This paper investigates the effect of airfoil steady loading on the sound generated by the interaction of an airfoil with a convected disturbance. A previous theory, which included only the incidence angle contribution to the mean loading, is extended to include camber. The theory is based on a linearization of the Euler equations about a nonuniform, 0(1) Mach number subsonic mean flow. The discussion concentrates on the case of a slightly cambered airfoil at small incidence angle, interacting with a gust whose wavelength is short compared to the airfoil chord. The small parameter representing the amount of camber and incidence, and the large parameter representing the ratio of airfoil chord to disturbance wavelength, are utilized in a singular perturbation solution to the governing equations. Acoustic power calculations reveal that the amount of sound generated increases significantly with increased loading. More importantly, it is shown that the radiated acoustic power correlates very well with the strength of the mean flow around the leading edge.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Wilder, M. C.

    1994-01-01

    A unique active flow-control device is proposed for the control of unsteady separated flow associated with the dynamic stall of airfoils. The device is an adaptive-geometry leading-edge which will allow controlled, dynamic modification of the leading-edge profile of an airfoil while the airfoil is executing an angle-of-attack pitch-up maneuver. A carbon-fiber composite skin has been bench tested, and a wind tunnel model is under construction. A baseline parameter study of compressible dynamic stall was performed for flow over an NACA 0012 airfoil. Parameters included Mach number, pitch rate, pitch history, and boundary layer tripping. Dynamic stall data were recorded via point-diffraction interferometry and the interferograms were analyzed with in-house developed image processing software. A new high-speed phase-locked photographic image recording system was developed for real-time documentation of dynamic stall.

  14. Turbine airfoil to shround attachment

    DOEpatents

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

    2014-05-06

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

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

    NASA Astrophysics Data System (ADS)

    Ramesh, Kiran

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

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

    NASA Technical Reports Server (NTRS)

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

    1951-01-01

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

  17. Edge of polar cap patches

    NASA Astrophysics Data System (ADS)

    Hosokawa, K.; Taguchi, S.; Ogawa, Y.

    2016-04-01

    On the night of 4 December 2013, a sequence of polar cap patches was captured by an all-sky airglow imager (ASI) in Longyearbyen, Norway (78.1°N, 15.5°E). The 630.0 nm airglow images from the ASI of 4 second exposure time, oversampled the emission of natural lifetime (with quenching) of at least ˜30 sec, introduce no observational blurring effects. By using such high-quality ASI images, we succeeded in visualizing an asymmetry in the gradients between the leading/trailing edges of the patches in a 2-D fashion. The gradient in the leading edge was found to be 2-3 times steeper than that in the trailing edge. We also identified fingerlike structures, appearing only along the trailing edge of the patches, whose horizontal scale size ranged from 55 to 210 km. These fingers are considered to be manifestations of plasma structuring through the gradient-drift instability (GDI), which is known to occur only along the trailing edge of patches. That is, the current 2-D observations visualized, for the first time, how GDI stirs the patch plasma and such a mixing process makes the trailing edge more gradual. This result strongly implies a close connection between the GDI-driven plasma stirring and the asymmetry in the large-scale shape of patches and then suggests that the fingerlike structures can be used as markers to estimate the fine-scale structure in the plasma flow within patches.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. Oregon Trail Teacher's Guide.

    ERIC Educational Resources Information Center

    National Park Service (Dept. of Interior), Washington, DC.

    The road to the U.S. West, known as the Oregon Trail, had its first real traffic in 1843 when a group of about 1000 people left Independence, Missouri and traveled west. This teacher's guide contains short descriptions of the main landmarks and stopping points that were significant along the northwest portion of the Oregon Trail. The guide is…

  3. 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.

  4. 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)

  5. Nozzle airfoil having movable nozzle ribs

    DOEpatents

    Yu, Yufeng Phillip; Itzel, Gary Michael

    2002-01-01

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

  6. Argentine ant trail pheromone disruption is mediated by trail concentration.

    PubMed

    Suckling, David Maxwell; Stringer, Lloyd D; Corn, Joshua E

    2011-10-01

    Argentine ant trail pheromone disruption, using continuous release of the trail pheromone compound (Z)-9-hexadecanal, reduces the incidence of trails and foraging rates of field populations. However, little is known about the concentrations of pheromone required for successful disruption. We hypothesized that higher pheromone quantities would be necessary to disrupt larger ant populations. To test this, we laid a 30-cm long base trail of (Z)-9-hexadecanal on a glass surface at low and high rates (1 and 100 pg/cm) (Trail 1), and laid a second, shorter trail (Trail 2, 10 cm long, located 1.5 cm upwind) near the middle of Trail 1 at six rates (1, 10, 100, 1,000, 10,000, and 100,000 pg/cm). We then recorded and digitized movements of individual ants following Trail 1, and derived a regression statistic, r (2), as an index of trail integrity, and also recorded arrival success at the other end of the trail (30 cm) near a food supply. Disruption of trails required 100 fold more pheromone upwind, independent of base-trail concentration. This implies that in the field, trail disruption is likely to be less successful against high ant-trail densities (greater concentration of trail pheromone), and more successful against newly formed or weak trails, as could be expected along invasion fronts. PMID:21964852

  7. 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.

  8. Effect of an extendable slat on the stall behavior of a VR-12 airfoil

    NASA Technical Reports Server (NTRS)

    Dehugues, P. Plantin; Mcalister, K. W.; Tung, C.

    1993-01-01

    Experimental and computational tests were performed on a VR-12 airfoil to determine if the dynamic-stall behavior that normally accompanies high-angle pitch oscillations could be modified by segmenting the forward portion of the airfoil and extending it ahead of the main element. In the extended position the configuration would appear as an airfoil with a leading-edge slat, and in the retracted position it would appear as a conventional VR-12 airfoil. The calculations were obtained from a numerical code that models the vorticity transport equation for an incompressible fluid. These results were compared with test data from the water tunnel facility of the Aeroflightdynamics Directorate at Ames Research Center. Steady and unsteady flows around both airfoils were examined at angles of attack between 0 and 30 deg. The Reynolds number was fixed at 200,000 and the unsteady pitch oscillations followed a sinusoidal motion described by alpha = alpha(sub m) + 10 deg sin(omega t). The mean angle (alpha(sub m)) was varied from 10 to 20 deg and the reduced frequency from 0.05 to 0.20. The results from the experiment and the calculations show that the extended-slat VR-12 airfoil experiences a delay in both static and dynamic stall not experienced by the basic VR-12 airfoil.

  9. A Rapid Distortion Theory modified turbulence spectra for semi-analytical airfoil noise prediction

    NASA Astrophysics Data System (ADS)

    Santana, Leandro D.; Christophe, Julien; Schram, Christophe; Desmet, Wim

    2016-11-01

    This paper proposes an implementation of the Rapid Distortion Theory, for the prediction of the noise resulting from the interaction of an airfoil with incoming turbulence. In the framework of the semi-analytical modeling strategy known as Amiet's theory, this interaction mechanism is treated in a linearized form where the airfoil thickness, camber and angle of attack are assumed negligible, leading to a frozen turbulence description of the incident gust. Important semi-analytical developments have been proposed in the literature to improve the modeling of the gust-airfoil interaction accounting for parallel and skewed gusts, non-rectangular linearized airfoil shapes or blade tip effects. This work is rather focused on the investigation of the distortion of turbulence that occurs in the vicinity of the airfoil leading edge, compared with Rapid Distortion Theory, where main results are briefly reminded in this paper. The main contribution of this work is a detailed experimental investigation of the evolution of turbulent quantities relevant to noise production, performed in the close vicinity of the airfoil leading edge subjected to grid turbulence, by means of stereoscopic Particle Image Velocimetry measurements. The results indicate that the distortion effects are concentrated in a narrow region close to the stagnation point of the leading edge, with dimension of the order of its radius of curvature. Additionally, it is shown that the turbulence intensity grows significantly as the flow approaches the airfoil leading-edge. Based on those results, a modified turbulence spectrum is proposed to describe the incoming turbulence in Amiet's theory. The sound predictions show a significantly better match with acoustic measurements than using the original turbulence model.

  10. DIRBE Comet Trails

    NASA Technical Reports Server (NTRS)

    Arendt, Richard G.

    2015-01-01

    Re-examination of the COBE DIRBE data reveals the thermal emission of several comet dust trails.The dust trails of 1P/Halley, 169P/NEAT, and 3200 Phaethon have not been previously reported.The known trails of 2P/Encke, and 73P/Schwassmann-Wachmann 3 are also seen. The dust trails have 12 and 25 microns surface brightnesses of <0.1 and <0.15 MJy/sr, respectively, which is <1% of the zodiacal light intensity. The trails are very difficult to see in any single daily image of the sky, but are evident as rapidly moving linear features in movies of the DIRBE data. Some trails are clearest when crossing through the orbital plane of the parent comet, but others are best seen at high ecliptic latitudes as the Earth passes over or under the dust trail. All these comets have known associations with meteor showers. This re-examination also reveals one additional comet and 13 additional asteroids that had not previously been recognized in the DIRBE data.

  11. DRBE comet trails

    SciTech Connect

    Arendt, Richard G.

    2014-12-01

    Re-examination of the Cosmic Background Explorer Diffuse Infrared Background Experiment (DIRBE) data reveals the thermal emission of several comet dust trails. The dust trails of 1P/Halley, 169P/NEAT, and 3200 Phaethon have not been previously reported. The known trails of 2P/Encke and 73P/Schwassmann–Wachmann 3 are also seen. The dust trails have 12 and 25 μm surface brightnesses of <0.1 and <0.15 MJy sr{sup −1}, respectively, which is <1% of the zodiacal light intensity. The trails are very difficult to see in any single daily image of the sky, but are evident as rapidly moving linear features in movies of the DIRBE data. Some trails are clearest when crossing through the orbital plane of the parent comet, but others are best seen at high ecliptic latitudes as the Earth passes over or under the dust trail. All these comets have known associations with meteor showers. This re-examination also reveals 1 additional comet and 13 additional asteroids that had not previously been recognized in the DIRBE data.

  12. DIRBE Comet Trails

    NASA Astrophysics Data System (ADS)

    Arendt, Richard G.

    2014-12-01

    Re-examination of the Cosmic Background Explorer Diffuse Infrared Background Experiment (DIRBE) data reveals the thermal emission of several comet dust trails. The dust trails of 1P/Halley, 169P/NEAT, and 3200 Phaethon have not been previously reported. The known trails of 2P/Encke and 73P/Schwassmann-Wachmann 3 are also seen. The dust trails have 12 and 25 μm surface brightnesses of \\lt 0.1 and \\lt 0.15 MJy sr-1, respectively, which is \\lt 1% of the zodiacal light intensity. The trails are very difficult to see in any single daily image of the sky, but are evident as rapidly moving linear features in movies of the DIRBE data. Some trails are clearest when crossing through the orbital plane of the parent comet, but others are best seen at high ecliptic latitudes as the Earth passes over or under the dust trail. All these comets have known associations with meteor showers. This re-examination also reveals 1 additional comet and 13 additional asteroids that had not previously been recognized in the DIRBE data.

  13. Flow characteristics over NACA4412 airfoil at low Reynolds number

    NASA Astrophysics Data System (ADS)

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

    2016-03-01

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

  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. A Method for the Constrained Design of Natural Laminar Flow Airfoils

    NASA Technical Reports Server (NTRS)

    Green, Bradford E.; Whitesides, John L.; Campbell, Richard L.; Mineck, Raymond E.

    1996-01-01

    A fully automated iterative design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. Drag reductions have been realized using the design method over a range of Mach numbers, Reynolds numbers and airfoil thicknesses. The thrusts of the method are its ability to calculate a target N-Factor distribution that forces the flow to undergo transition at the desired location; the target-pressure-N-Factor relationship that is used to reduce the N-Factors in order to prolong transition; and its ability to design airfoils to meet lift, pitching moment, thickness and leading-edge radius constraints while also being able to meet the natural laminar flow constraint. The method uses several existing CFD codes and can design a new airfoil in only a few days using a Silicon Graphics IRIS workstation.

  16. Tables of properties of airfoil polynomials

    NASA Technical Reports Server (NTRS)

    Desmarais, Robert N.; Bland, Samuel R.

    1995-01-01

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

  17. Make a Nature Trail

    ERIC Educational Resources Information Center

    Johnson, Janice K.

    1973-01-01

    Discusses the planning, construction, use, and maintenance of a nature trail. Ideal for demonstrating interrelationships between plants and animals, conservation practices, wildlife management, plant succession, forestry, geologic features and other scientific phenomena. (JR)

  18. 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.

  19. Airfoil deposition model

    NASA Technical Reports Server (NTRS)

    Kohl, F. J.

    1982-01-01

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

  20. Airfoil Vibration Dampers program

    NASA Technical Reports Server (NTRS)

    Cook, Robert M.

    1991-01-01

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

  1. 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.

  2. 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.

  3. 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.

  4. Low-Reynolds number compressible flow around a triangular airfoil

    NASA Astrophysics Data System (ADS)

    Munday, Phillip; Taira, Kunihiko; Suwa, Tetsuya; Numata, Daiju; Asai, Keisuke

    2013-11-01

    We report on the combined numerical and experimental effort to analyze the nonlinear aerodynamics of a triangular airfoil in low-Reynolds number compressible flow that is representative of wings on future Martian air vehicles. The flow field around this airfoil is examined for a wide range of angles of attack and Mach numbers with three-dimensional direct numerical simulations at Re = 3000 . Companion experiments are conducted in a unique Martian wind tunnel that is placed in a vacuum chamber to simulate the Martian atmosphere. Computational findings are compared with pressure sensitive paint and direct force measurements and are found to be in agreement. The separated flow from the leading edge is found to form a large leading-edge vortex that sits directly above the apex of the airfoil and provides enhanced lift at post stall angles of attack. For higher subsonic flows, the vortical structures elongate in the streamwise direction resulting in reduced lift enhancement. We also observe that the onset of spanwise instability for higher angles of attack is delayed at lower Mach numbers. Currently at Mitsubishi Heavy Industries, Ltd., Nagasaki.

  5. 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

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

    NASA Technical Reports Server (NTRS)

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

    1974-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1988-01-01

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

  8. Experimental studies of the boundary layer on an airfoil at low Reynolds numbers

    NASA Technical Reports Server (NTRS)

    Jansen, B. J., Jr.; Mueller, T. J.

    1983-01-01

    An experimental wind tunnel investigation was carried out to study the effect of laminar separation bubbles on a NACA 66(3)-018 airfoil for Reynolds numbers less than 4.0 x 10 to the 5th. Leading edge laminar separation bubbles formed for angles of attack of approximately 7 to 12 deg. To study the leading edge separation bubble more closely, hotwire anemometer measurements were made in the airfoil a Reynolds number of 8.0 x 10 to the 4th. Velocity and turbulence intensity profiles were obtained and boundary layer parameters were calculated. Frequency spectra were also calculated at key points in the airfoil boundary layer for this case. Correlation of the anemometry data with static pressure distributions, and flow visualization data provided insight into laminar separation bubble behavior at low Reynolds numbers.

  9. Certification trails for data structures

    NASA Technical Reports Server (NTRS)

    Sullivan, Gregory F.; Masson, Gerald M.

    1993-01-01

    Certification trails are a recently introduced and promising approach to fault detection and fault tolerance. The applicability of the certification trail technique is significantly generalized. Previously, certification trails had to be customized to each algorithm application; trails appropriate to wide classes of algorithms were developed. These certification trails are based on common data-structure operations such as those carried out using these sets of operations such as those carried out using balanced binary trees and heaps. Any algorithms using these sets of operations can therefore employ the certification trail method to achieve software fault tolerance. To exemplify the scope of the generalization of the certification trail technique provided, constructions of trails for abstract data types such as priority queues and union-find structures are given. These trails are applicable to any data-structure implementation of the abstract data type. It is also shown that these ideals lead naturally to monitors for data-structure operations.

  10. Hook nozzle arrangement for supporting airfoil vanes

    SciTech Connect

    Shaffer, James E.; Norton, Paul F.

    1996-01-01

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

  11. Hook nozzle arrangement for supporting airfoil vanes

    DOEpatents

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

    1996-02-20

    A gas turbine engine`s nozzle structure includes a nozzle support ring, a plurality of shroud segments, and a plurality of airfoil vanes. The plurality of shroud segments are distributed around the nozzle support ring. Each airfoil vane is connected to a corresponding shroud segment so that the airfoil vanes are also distributed around the nozzle support ring. Each shroud segment has a hook engaging the nozzle support ring so that the shroud segments and corresponding airfoil vanes are supported by the nozzle support ring. The nozzle support ring, the shroud segments, and the airfoil vanes may be ceramic. 8 figs.

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

    NASA Astrophysics Data System (ADS)

    Saeed, Farooq

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

  13. 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.

  14. Leading edge protection for composite blades

    NASA Technical Reports Server (NTRS)

    Brantley, J. W.; Irwin, T. P. (Inventor)

    1977-01-01

    A laminated filament composite structure, such as an airfoil for use in an environment in which it is subjected to both foreign object impact and bending is provided with improved leading edge protection. At least one fine wire mesh layer is partially bonded within the composite structure along its neutral bending axis. A portion of the wire mesh layer extends beyond the neutral bending axis and partially around the leading edge where it is bonded to the outer periphery of the primary composite structure. The wire mesh is clad with a metal such as nickel to provide an improved leading edge protective device which is firmly anchored within the composite structure. Also described is a novel method of constructing a composite airfoil so as to further minimize the possibility of losing the leading edge protective device due to delamination caused by impact and bending.

  15. Comprehensive Trail Making Test

    ERIC Educational Resources Information Center

    Gray, Rebecca

    2006-01-01

    The Comprehensive Trail Making Test (CTMT) is designed to be used in neuropsychological assessment for the purposes of detecting effects of brain defects and deficits and in tracking progress in rehabilitation. More specific purposes include the detection of frontal lobe deficits, problems with psychomotor speed, visual search and sequencing,…

  16. Effect of Intercycle Ice Accretions on Airfoil Performance

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Kearney, John; Glezer, Ari

    2013-11-01

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

  18. 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.

  19. Aerodynamic characteristics of airfoils with ice accretions

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

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

  20. Evaluation of a stalled airfoil analysis program

    NASA Technical Reports Server (NTRS)

    Rumsey, C. L.

    1985-01-01

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

  1. 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.

  2. 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.

  3. Airfoil seal system for gas turbine engine

    SciTech Connect

    Diakunchak, Ihor S.

    2013-06-25

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

  4. Computation of airfoil buffet boundaries

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

  5. 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.

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

    NASA Technical Reports Server (NTRS)

    Mathews, Charles W; Thompson, Jim Rogers

    1950-01-01

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

  7. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Cebeci, T.; Carr, L. W.

    1983-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  11. Results of the AVATAR project for the validation of 2D aerodynamic models with experimental data of the DU95W180 airfoil with unsteady flap

    NASA Astrophysics Data System (ADS)

    Ferreira, C.; Gonzalez, A.; Baldacchino, D.; Aparicio, M.; Gómez, S.; Munduate, X.; Garcia, N. R.; Sørensen, J. N.; Jost, E.; Knecht, S.; Lutz, T.; Chassapogiannis, P.; Diakakis, K.; Papadakis, G.; Voutsinas, S.; Prospathopoulos, J.; Gillebaart, T.; van Zuijlen, A.

    2016-09-01

    The FP7 AdVanced Aerodynamic Tools for lArge Rotors - Avatar project aims to develop and validate advanced aerodynamic models, to be used in integral design codes for the next generation of large scale wind turbines (10-20MW). One of the approaches towards reaching rotors for 10-20MW size is the application of flow control devices, such as flaps. In Task 3.2: Development of aerodynamic codes for modelling of flow devices on aerofoils and, rotors of the Avatar project, aerodynamic codes are benchmarked and validated against the experimental data of a DU95W180 airfoil in steady and unsteady flow, for different angle of attack and flap settings, including unsteady oscillatory trailing-edge-flap motion, carried out within the framework of WP3: Models for Flow Devices and Flow Control, Task 3.1: CFD and Experimental Database. The aerodynamics codes are: AdaptFoil2D, Foil2W, FLOWer, MaPFlow, OpenFOAM, Q3UIC, ATEFlap. The codes include unsteady Eulerian CFD simulations with grid deformation, panel models and indicial engineering models. The validation cases correspond to 18 steady flow cases, and 42 unsteady flow cases, for varying angle of attack, flap deflection and reduced frequency, with free and forced transition. The validation of the models show varying degrees of agreement, varying between models and flow cases.

  12. Application of Particle Image Velocimetry to a Study of Flow About a Multi-Element Airfoil

    NASA Technical Reports Server (NTRS)

    Walker, Stephen M.; Baganoff, Donald

    1996-01-01

    An experiment was performed on the flap tip vortex shed from a half span Fowler flap. This flap was mounted on a 5 foot span NACA 63(2)-215 Mod B airfoil in the 7 by 10 foot wind tunnel at NASA Ames Research Center. Several noise reduction studies were performed with this model, and the addition of the Particle Image Velocimetry (PIV) research discussed here served as a proof case of large scale PIV. The measurement plane investigated here was a cross plane region. This is cross plane relative to the freestream flow direction. The measurement plane was located at a position 18 inches downstream of the flap trailing edge. This served to prove that measurements could also be made in the more difficult cross plane direction rather than in the downstream flow direction. Lastly the PIV data was used as a practical research tool that yielded important results that could not otherwise be obtained. The flow field area measured was 40 cm by 40 cm square, and served to characterize the downstream flow characteristics of the flap tip vortex under three configurations: the baseline configuration which was the flap and the wing only; the baseline with the addition of a 3/4 span slat; and the baseline with a Flap Edge Device which was designed to reduce the noise generated at the flap. All configurations were tested at a freestream velocity of 64.84 m/s. The test resulted in average velocity fields for the three configurations tested. The velocity fields aided in verifying other testing methods on this particular experiment, and also yielded further insight into the characteristics of the flap tip vortex under the three configurations considered. The velocity data was reduced, and we were able to calculate the vorticity of the flow field. From the position of minimum vorticity the location of the center of the vortex was determined. The circulation was also calculated and aided in comparing the effects of the three configurations on the lifting characteristics of the flap.

  13. Crescentic ramp turbine stage

    NASA Technical Reports Server (NTRS)

    Lee, Ching-Pang (Inventor); Tam, Anna (Inventor); Kirtley, Kevin Richard (Inventor); Lamson, Scott Henry (Inventor)

    2007-01-01

    A turbine stage includes a row of airfoils joined to corresponding platforms to define flow passages therebetween. Each airfoil includes opposite pressure and suction sides and extends in chord between opposite leading and trailing edges. Each platform includes a crescentic ramp increasing in height from the leading and trailing edges toward the midchord of the airfoil along the pressure side thereof.

  14. Differences in the impacts of formal and informal recreational trails on urban forest loss and tree structure.

    PubMed

    Ballantyne, Mark; Pickering, Catherine Marina

    2015-08-15

    Recreational trails are one of the most common types of infrastructure used for nature-based activities such as hiking and mountain biking worldwide. Depending on their design, location, construction, maintenance and use, these trails differ in their environmental impacts. There are few studies, however, comparing the impacts of different trail types including between formal management-created trails and informal visitor-created trails. Although both types of trails can be found in remote natural areas, dense networks of them often occur in forests close to cities where they experience intense visitor use. To assess the relative impacts of different recreational trails in urban forests, we compared the condition of the trail surface, loss of forest strata and changes in tree structure caused by seven types of trails (total network 46.1 km) traversing 17 remnants of an endangered urban forest in Australia. After mapping and classifying all trails, we assessed their impact on the forest condition at 125 sites (15 sites per trail type, plus 15 control sites within undisturbed forest). On the trail sites, the condition of the trail surface, distance from the trail edge to four forest strata (litter, understory, midstorey and tree cover) and structure of the tree-line were assessed. Informal trails generally had poorer surface conditions and were poorly-designed and located. Per site, formal and informal trails resulted in similar loss of forest strata, with wider trails resulting in greater loss of forest. Because there were more informal trails, however, they accounted for the greatest cumulative forest loss. Structural impacts varied, with the widest informal trails and all formal hardened trails resulting in similar reductions in canopy cover and tree density but an increase in saplings. These structural impacts are likely a function of the unregulated and intense use of large informal trails, and disturbance from the construction and maintenance of formal trails

  15. Differences in the impacts of formal and informal recreational trails on urban forest loss and tree structure.

    PubMed

    Ballantyne, Mark; Pickering, Catherine Marina

    2015-08-15

    Recreational trails are one of the most common types of infrastructure used for nature-based activities such as hiking and mountain biking worldwide. Depending on their design, location, construction, maintenance and use, these trails differ in their environmental impacts. There are few studies, however, comparing the impacts of different trail types including between formal management-created trails and informal visitor-created trails. Although both types of trails can be found in remote natural areas, dense networks of them often occur in forests close to cities where they experience intense visitor use. To assess the relative impacts of different recreational trails in urban forests, we compared the condition of the trail surface, loss of forest strata and changes in tree structure caused by seven types of trails (total network 46.1 km) traversing 17 remnants of an endangered urban forest in Australia. After mapping and classifying all trails, we assessed their impact on the forest condition at 125 sites (15 sites per trail type, plus 15 control sites within undisturbed forest). On the trail sites, the condition of the trail surface, distance from the trail edge to four forest strata (litter, understory, midstorey and tree cover) and structure of the tree-line were assessed. Informal trails generally had poorer surface conditions and were poorly-designed and located. Per site, formal and informal trails resulted in similar loss of forest strata, with wider trails resulting in greater loss of forest. Because there were more informal trails, however, they accounted for the greatest cumulative forest loss. Structural impacts varied, with the widest informal trails and all formal hardened trails resulting in similar reductions in canopy cover and tree density but an increase in saplings. These structural impacts are likely a function of the unregulated and intense use of large informal trails, and disturbance from the construction and maintenance of formal trails

  16. Numerical simulation of the interaction of a vortex with stationary airfoil in transonic flow

    NASA Technical Reports Server (NTRS)

    Srinivasan, G. R.; Mccroskey, W. J.; Kutler, P.

    1984-01-01

    A perturbation form of an implicit conservative, noniterative numerical algorithm for the two-dimensional thin layer Navier-Stokes and Euler equations is used to compute the interaction flow-field of a vortex with stationary airfoil. A Lamb-like analytical vortex having a finite core is chosen to interact with a thick (NACA 0012) and a thin (NACA 64A006) airfoil independently in transonic flow. Two different configurations of vortex interaction are studied, viz., (1) when the vortex is fixed at one location in the flowfield, and (2) when the vortex is convecting past the airfoil at freestream velocity. Parallel computations of this interacting flowfield are also done using a version of the Transonic Small Disturbance Code (ATRAN2). A special treatment of the leading edge region for thin airfoils is included in this code. With this, the three methods gave qualitatively similar results for the weaker interactions considered in this study. However, the strongest interactions considered proved to be beyond the capabilities of the small disturbance code. The results also show a far greater influence of the vortex on the airfoil flowfield when the vortex is stationary than when it is convecting with the flow.

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

    NASA Astrophysics Data System (ADS)

    Boutilier, Michael S. H.; Yarusevych, Serhiy

    2012-08-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

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

  19. Detail of double pole utility tower at the edge of ...

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

    Detail of double pole utility tower at the edge of the mesa. National Old Trails Road right-of-way crosses in foreground over a dry laid basaltic rock retaining wall. View south. - La Bajada Historic Trails and Roads, Approximately 1 mile East/Northeast of intersection of State Highway 16 and Indian Service Road 841, La Bajada, Santa Fe County, NM

  20. Airfoil profile in a nonuniform flow

    NASA Technical Reports Server (NTRS)

    Polasek, J.

    1978-01-01

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