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Sample records for airfoil leading edge

  1. 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

  2. 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.

  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. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. 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.

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

    NASA Astrophysics Data System (ADS)

    Kraljic, Matthew; Rusak, Zvi; Wang, Shixiao

    2015-11-01

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

  12. 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.

  13. 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.

  14. 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.

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

    NASA Astrophysics Data System (ADS)

    Phan, Nhan Huu

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

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  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. 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.

  4. 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.

  5. 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

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Supersonic Leading Edge Receptivity

    NASA Technical Reports Server (NTRS)

    Maslov, Anatoly A.

    1998-01-01

    This paper describes experimental studies of leading edge boundary layer receptivity for imposed stream disturbances. Studies were conducted in the supersonic T-325 facility at ITAM and include data for both sharp and blunt leading edges. The data are in agreement with existing theory and should provide guidance for the development of more complete theories and numerical computations of this phenomena.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Wing Leading Edge Joint Laminar Flow Tests

    NASA Technical Reports Server (NTRS)

    Drake, Aaron; Westphal, Russell V.; Zuniga, Fanny A.; Kennelly, Robert A., Jr.; Koga, Dennis J.

    1996-01-01

    An F-104G aircraft at NASA's Dryden Flight Research Center has been equipped with a specially designed and instrumented test fixture to simulate surface imperfections of the type likely to be present near the leading edge on the wings of some laminar flow aircraft. The simulated imperfections consisted of five combinations of spanwise steps and gaps of various sizes. The unswept fixture yielded a pressure distribution similar to that of some laminar flow airfoils. The experiment was conducted at cruise conditions typical for business-jets and light transports: Mach numbers were in the range 0.5-0.8, and unit Reynolds numbers were 1.5-2.5 million per foot. Skin friction measurements indicated that laminar flow was often maintained for some distance downstream of the surface imperfections. Further work is needed to more precisely define transition location and to extend the experiments to swept-wing conditions and a broader range of imperfection geometries.

  4. 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

  5. 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.

  6. 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.

  7. 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.

  8. 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

  9. Formation and Development of the Dynamic Stall Vortex on a Wing with Leading Edge Tubercles

    NASA Astrophysics Data System (ADS)

    Hrynuk, John; Bohl, Douglas

    2015-11-01

    Humpback whales are unique in that their flippers have leading edge ``bumps'' or tubercles. Past work on airfoils inspired by whale flippers has centered on the static aerodynamic characteristics of these airfoils. The current study uses Molecular Tagging Velocimetry (MTV) to investigate the effects of tubercles on dynamically pitching NACA 0012 airfoils. A baseline (i.e. straight leading edge) wing and one modified with leading edge tubercles are investigated. Tracking of the Dynamic Stall Vortex (DSV) is performed to quantitatively compare the DSV formation location, path, and convective velocity for tubercled and baseline wings. The results show that there is a spanwise variation in the initial formation location and motion of the DSV on the modified wing. Once formed, the DSV aligns into a more uniform spanwise structure. As the pitching motion progresses, the DSV on the modified wing convects away from the airfoil surface later and slower than is observed for the baseline airfoil. The results indicate that the tubercles may delay stall when compared to the baseline airfoil. This work was supported by NSF Grant # 0845882.

  10. Vertical axis wind turbine airfoil

    DOEpatents

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

    2012-12-18

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

  11. 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. 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.

  13. Analysis of a theoretically optimized transonic airfoil

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  14. 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.

  15. Investigation of a Laminar Flow Leading Edge

    NASA Technical Reports Server (NTRS)

    Drake, Aaron; Kennelly, Robert A., Jr.; Koga, Dennis J.; Westphal, Russell V.; Zuniga, Fanny

    1994-01-01

    The recent resurgence of interest in utilizing laminar flow on aircraft surfaces for reduction in skin friction drag has generated a considerable amount of research in natural laminar flow (NLF) and hybrid laminar flow control (HLFC) on transonic aircraft wings. This research has focused primarily on airfoil design and understanding transition behavior with little concern for the surface imperfections and manufacturing variations inherent to most production aircraft. In order for laminar flow to find wide-spread use on production aircraft, techniques for constructing the wings must be found such that the large surface imperfections present in the leading edge region of current aircraft do not occur. Toward this end, a modification to existing leading edge construction techniques was devised such that the resulting surface did not contain large gaps and steps as are common on current production aircraft of this class. A lowspeed experiment was first conducted on a simulation of the surface that would result from this construction technique. Preston tube measurements of the boundary layer downstream of the simulated joint and flow visualization using sublimation chemicals validated the literature on the effects of steps on a laminar boundary layer. These results also indicated that the construction technique was indeed compatible with laminar flow. In order to fully validate the compatibility of this construction technique with laminar flow, thus proving that it is possible to build wings that are smooth enough to be used on business jets and light transports in a manner compatible with laminar flow, a flight experiment is being conducted. In this experiment Mach number and Reynolds number will be matched in a real flight environment. The experiment is being conducted using the NASA Dryden F-104 Flight Test Fixture (FTF). The FTF is a low aspect ratio ventral fin mounted beneath an F-104G research aircraft. A new nose shape was designed and constructed for this

  16. 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.

  17. Single velocity-component modeling of leading edge turbulence interaction noise.

    PubMed

    Gill, J; Zhang, X; Joseph, P

    2015-06-01

    A computational aeroacoustics approach is used to predict leading edge turbulence interaction noise for real airfoils. One-component (transverse), two-component (transverse and streamwise), and three-component (transverse, streamwise, and spanwise) synthesized turbulence disturbances are modeled instead of harmonic transverse gusts, to which previous computational studies of leading edge noise have often been confined. The effects of the inclusion of streamwise and spanwise disturbances on the noise are assessed. It is shown that accurate noise predictions can be made by modeling only transverse disturbances which reduces the computational expense of simulations. The accuracy of using only transverse disturbances is assessed for symmetric and cambered airfoils, and also for airfoils at non-zero angle of attack.

  18. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

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

    1945-01-01

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

  19. 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.

  20. Effects of Nose Radius and Aerodynamic Loading on Leading Edge Receptivity

    NASA Technical Reports Server (NTRS)

    Hammerton, P. W.; Kerschen, E. J.

    1998-01-01

    An analysis is presented of the effects of airfoil thickness and mean aerodynamic loading on boundary-layer receptivity in the leading-edge region. The case of acoustic free-stream disturbances, incident on a thin cambered airfoil with a parabolic leading edge in a low Mach number flow, is considered. An asymptotic analysis based on large Reynolds number is developed, supplemented by numerical results. The airfoil thickness distribution enters the theory through a Strouhal number based on the nose radius of the airfoil, S = (omega)tau(sub n)/U, where omega is the frequency of the acoustic wave and U is the mean flow speed. The influence of mean aerodynamic loading enters through an effective angle-of-attack parameter ti, related to flow around the leading edge from the lower surface to the upper. The variation of the receptivity level is analyzed as a function of S, mu, and characteristics of the free-stream acoustic wave. For an unloaded leading edge, a finite nose radius dramatically reduces the receptivity level compared to that for a flat plate, the amplitude of the instability waves in the boundary layer being decreased by an order of magnitude when S = 0.3. Modest levels of aerodynamic loading are found to further decrease the receptivity level for the upper surface of the airfoil, while an increase in receptivity level occurs for the lower surface. For larger angles of attack close to the critical angle for boundary layer separation, a local rise in the receptivity level occurs for the upper surface, while for the lower surface the receptivity decreases. The effects of aerodynamic loading are more pronounced at larger values of S. Oblique acoustic waves produce much higher receptivity levels than acoustic waves propagating downstream parallel to the airfoil chord.

  1. Laminar flow control leading edge glove flight test article development

    NASA Technical Reports Server (NTRS)

    Pearce, W. E.; Mcnay, D. E.; Thelander, J. A.

    1984-01-01

    A laminar flow control (LFC) flight test article was designed and fabricated to fit into the right leading edge of a JetStar aircraft. The article was designed to attach to the front spar and fill in approx. 70 inches of the leading edge that are normally occupied by the large slipper fuel tank. The outer contour of the test article was constrained to align with an external fairing aft of the front spar which provided a surface pressure distribution over the test region representative of an LFC airfoil. LFC is achieved by applying suction through a finely perforated surface, which removes a small fraction of the boundary layer. The LFC test article has a retractable high lift shield to protect the laminar surface from contamination by airborne debris during takeoff and low altitude operation. The shield is designed to intercept insects and other particles that could otherwise impact the leading edge. Because the shield will intercept freezing rain and ice, a oozing glycol ice protection system is installed on the shield leading edge. In addition to the shield, a liquid freezing point depressant can be sprayed on the back of the shield.

  2. Moveable Leading Edge Device for a Wing

    NASA Technical Reports Server (NTRS)

    Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

    2013-01-01

    A method and apparatus for managing a flight control surface system. A leading edge section on a wing of an aircraft is extended into a deployed position. A deformable section connects the leading edge section to a trailing section. The deformable section changes from a deformed shape to an original shape when the leading edge section is moved into the deployed position. The leading edge section on the wing is moved from the deployed position to an undeployed position. The deformable section changes to the deformed shape inside of the wing.

  3. The aerodynamic design of an advanced rotor airfoil

    NASA Technical Reports Server (NTRS)

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

    1978-01-01

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

  4. 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.

  5. Experimental investigation of the leading edge vortex on vertical axis wind turbine blades

    NASA Astrophysics Data System (ADS)

    Dunne, Reeve; McKeon, Beverley

    2012-11-01

    A NACA 0018 airfoil is pitched about the leading edge over a large angle of attack range (+/- ~40°) at a chord Reynolds number of 110,000 to simulate the flow over a single blade in a vertical axis wind turbine (VAWT). Particle image velocimetry (PIV) measurements are made to investigate the effects of pitching on leading edge vortex (LEV) development and separation. Time resolved experiments are performed to track vortex formation and convection over the airfoil for sinusoidal pitching motions corresponding to a VAWT trajectory as well as impulsive pitch up and pitch down motions. These results are compared to the wake of steady, post stall, high angle of attack airfoils (α =20° -30°). The characteristics of the leading edge vortex development and subsequent separation from the airfoil are discussed, with a view to characterizing its effect on power generation with VAWTs and future flow control strategies for turbine performance improvement. Funding from the Gordon and Betty Moore Foundation is gratefully acknowledged.

  6. 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.

  7. Cavitation on hydrofoils with sinusoidal leading edge

    NASA Astrophysics Data System (ADS)

    Johari, H.

    2015-12-01

    Cavitation characteristics of hydrofoils with sinusoidal leading edge were examined experimentally at a Reynolds number of 7.2 × 105. The hydrofoils had an underlying NACA 634-021 profile and an aspect ratio of 4.3. The sinusoidal leading edge geometries included three amplitudes of 2.5%, 5%, and 12% and two wavelengths of 25% and 50% of the mean chord length. Results revealed that cavitation on the leading edge-modified hydrofoils existed in pockets behind the troughs whereas the baseline hydrofoil produced cavitation along its entire span. Moreover, cavitation on the modified hydrofoils appeared at consistently lower angles of attack than on the baseline hydrofoil.

  8. Imaging The Leading Edge Of A Weld

    NASA Technical Reports Server (NTRS)

    Mcgee, William F.; Rybicki, Daniel J.

    1994-01-01

    Proposed optical system integrated into plasma arc welding torch provides image of leading edge of weld pool and welding-arc-initiation point. Welding torch aligned better with joint. System includes coherent bundle of optical fibers and transparent cup.

  9. 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.

  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. PMID:27586762

  11. Computation of leading-edge vortex flows

    NASA Technical Reports Server (NTRS)

    Newsome, R. W.; Thomas, J. L.

    1986-01-01

    The simulation of the leading edge vortex flow about a series of conical delta wings through solution of the Navier-Stokes and Euler equations is studied. The occurrence, the validity, and the usefulness of separated flow solutions to the Euler equations of particular interest. Central and upwind difference solutions to the governing equations are compared for a series of cross sectional shapes, including both rounded and sharp tip geometries. For the rounded leading edge and the flight condition considered, viscous solutions obtained with either central or upwind difference methods predict the classic structure of vortical flow over a highly swept delta wing. Predicted features include the primary vortex due to leading edge separation and the secondary vortex due to crossflow separation. Central difference solutions to the Euler equations show a marked sensitivity to grid refinement. On a coarse grid, the flow separates due to numerical error and a primary vortex which resembles that of the viscous solution is predicted. In contrast, the upwind difference solutions to the Euler equations predict attached flow even for first-order solutions on coarse grids. On a sufficiently fine grid, both methods agree closely and correctly predict a shock-curvature-induced inviscid separation near the leeward plane of symmetry. Upwind difference solutions to the Navier-Stokes and Euler equations are presented for two sharp leading edge geometries. The viscous solutions are quite similar to the rounded leading edge results with vortices of similar shape and size. The upwind Euler solutions predict attached flow with no separation for both geometries. However, with sufficient grid refinement near the tip or through the use of more accurate spatial differencing, leading edge separation results. Once the leading edge separation is established, the upwind solution agrees with recently published central difference solutions to the Euler equations.

  12. 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.

  13. The effects of leading edge and downstream film cooling on turbine vane heat transfer

    NASA Astrophysics Data System (ADS)

    Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

    1988-11-01

    The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

  14. The effects of leading edge and downstream film cooling on turbine vane heat transfer

    NASA Technical Reports Server (NTRS)

    Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

    1988-01-01

    The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

  15. 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.

  16. Development of Columbia Leading Edge Reconstruction System

    NASA Technical Reports Server (NTRS)

    Trautwein, John; Wegerif, Dan

    2004-01-01

    After the loss of Columbia in 2003, the Columbia Accident Investigation Board and NASA KSC directed personnel at the Launch Equipment Test Facility (LETF) to design and build high fidelity mock-ups of Columbia's left wing leading edges. These leading edge segments, constructed of reinforced carbon-carbon, were a major point of inquiry by the investigation team. The LETF engineers developed a concept of building a clear Lexan panel with an aluminum support structure ten percent larger than the original panel. The leading edge debris are attached to the Lexan panels and both the front and back side of each panel are visible for inspection. The entire assembly can be rotated, to provide visual access to the entire panel. Six carts were fabricated to support the thirteen panels. These carts could be set up in order, next to each other, to provide the desired inspection access. The carts and attached debris are currently located in the Vehicle Assembly Building at KSC.

  17. Wing Leading Edge Concepts for Noise Reduction

    NASA Technical Reports Server (NTRS)

    Shmilovich, Arvin; Yadlin, Yoram; Pitera, David M.

    2010-01-01

    This study focuses on the development of wing leading edge concepts for noise reduction during high-lift operations, without compromising landing stall speeds, stall characteristics or cruise performance. High-lift geometries, which can be obtained by conventional mechanical systems or morphing structures have been considered. A systematic aerodynamic analysis procedure was used to arrive at several promising configurations. The aerodynamic design of new wing leading edge shapes is obtained from a robust Computational Fluid Dynamics procedure. Acoustic benefits are qualitatively established through the evaluation of the computed flow fields.

  18. 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.

  19. A mechanism for mitigation of blade-vortex interaction using leading edge blowing flow control

    NASA Astrophysics Data System (ADS)

    Weiland, Chris; Vlachos, Pavlos P.

    2009-09-01

    The interaction of a vortical unsteady flow with structures is often encountered in engineering applications. Such flow structure interactions (FSI) can be responsible for generating significant loads and can have many detrimental structural and acoustic side effects, such as structural fatigue, radiated noise and even catastrophic results. Amongst the different types of FSI, the parallel blade-vortex interaction (BVI) is the most common, often encountered in helicopters and propulsors. In this work, we report on the implementation of leading edge blowing (LEB) active flow control for successfully minimizing the parallel BVI. Our results show reduction of the airfoil vibrations up to 38% based on the root-mean-square of the vibration velocity amplitude. This technique is based on displacing an incident vortex using a jet issued from the leading edge of a sharp airfoil effectively increasing the stand-off distance of the vortex from the body. The effectiveness of the method was experimentally analyzed using time-resolved digital particle image velocimetry (TRDPIV) recorded at an 800 Hz rate, which is sufficient to resolve the spatio-temporal dynamics of the flow field and it was combined with simultaneous accelerometer measurements of the airfoil, which was free to oscillate in a direction perpendicular to the freestream. Analysis of the flow field spectra and a Proper Orthogonal Decomposition (POD) of the TRDPIV data of the temporally resolved planar flow fields indicate that the LEB effectively modified the flow field surrounding the airfoil and increased the convecting vortices stand-off distance for over half of the airfoil chord length. It is shown that LEB also causes a redistribution of the flow field spectral energy over a larger range of frequencies.

  20. 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

  1. Cavitation on Hydrofoils with Leading Edge Protuberances

    NASA Astrophysics Data System (ADS)

    Custodio, Derrick; Henoch, Charles; Johari, Hamid; Office of Naval Research Collaboration

    2012-11-01

    The effects of spanwise-uniform sinusoidal leading edge protuberances on the flow characteristics and forces of finite-span hydrofoils under vaporous cavitation conditions were examined experimentally over angles of attack ranging from -9° α <= 27°. Two planforms were studied, rectangular and swept, at a Reynolds number of ~ 720,000. Two protuberance wavelengths, λ = 0.25 c and 0.50 c, and three amplitudes, A = 0.025 c, 0.05 c, and 0.12 c, were examined as they resemble the humpback whale flipper morphology. All hydrofoils retain a mean NACA 634-021 profile. The forces and moments were measured at a freestream velocity of 7.2 m/s, and high-speed digital photography was used to capture flow field images at several angles of attack. The cavitation number corresponding to incipient leading edge cavitation was also calculated. As far as forces and cavitation number are concerned, results show that the baseline hydrofoil tends to have nearly equal or improved performance over the modified hydrofoils at most angles of attack tested. Flow images reveal that it is possible that the extent of sheet and tip vortex cavitation can be reduced with the introduction of leading edge protuberances. The forces and cavitation characteristics will be presented. Sponsored by the ONR-ULI program.

  2. Improved Method for Prediction of Attainable Wing Leading-Edge Thrust

    NASA Technical Reports Server (NTRS)

    Carlson, Harry W.; McElroy, Marcus O.; Lessard, Wendy B.; McCullers, L. Arnold

    1996-01-01

    Prediction of the loss of wing leading-edge thrust and the accompanying increase in drag due to lift, when flow is not completely attached, presents a difficult but commonly encountered problem. A method (called the previous method) for the prediction of attainable leading-edge thrust and the resultant effect on airplane aerodynamic performance has been in use for more than a decade. Recently, the method has been revised to enhance its applicability to current airplane design and evaluation problems. The improved method (called the present method) provides for a greater range of airfoil shapes from very sharp to very blunt leading edges. It is also based on a wider range of Reynolds numbers than was available for the previous method. The present method, when employed in computer codes for aerodynamic analysis, generally results in improved correlation with experimental wing-body axial-force data and provides reasonable estimates of the measured drag.

  3. Characteristics of surface roughness associated with leading edge ice accretion

    NASA Technical Reports Server (NTRS)

    Shin, Jaiwon

    1994-01-01

    Detailed size measurements of surface roughness associated with leading edge ice accretions are presented to provide information on characteristics of roughness and trends of roughness development with various icing parameters. Data was obtained from icing tests conducted in the Icing Research Tunnel (IRT) at NASA Lewis Research Center (LeRC) using a NACA 0012 airfoil. Measurements include diameters, heights, and spacing of roughness elements along with chordwise icing limits. Results confirm the existence of smooth and rough ice zones and that the boundary between the two zones (surface roughness transition region) moves upstream towards stagnation region with time. The height of roughness grows as the air temperature and the liquid water content increase, however, the airspeed has little effect on the roughness height. Results also show that the roughness in the surface roughness transition region grows during a very early stage of accretion but reaches a critical height and then remains fairly constant. Results also indicate that a uniformly distributed roughness model is only valid at a very initial stage of the ice accretion process.

  4. Investigation of Porous Gas-Heated Leading-Edge Section for Icing Protection of a Delta Wing

    NASA Technical Reports Server (NTRS)

    Bowden, Dean T.

    1955-01-01

    A tip section of a delta wing having an NACA 0004-65 airfoil section and a 600 leading-edge sweepback was equipped with a porous leading-edge section through which hot gas was 'bled for anti-icing. Heating rates for anti-icing were determined for a wide range of icing conditions. The effects of gas flow through the porous leading-edge section on airfoil pressure distribution and drag in dry air were investigated. The drag increase caused by an ice formation on the unheated airfoil was measured for several icing conditions. Experimental porous surface- to free-stream convective heat-transfer coefficients were obtained in dry air and compared with theory. Adequate icing protection was obtained at all icing conditions investigated. Savings in total gas-flow rate up to 42 percent may be obtained with no loss in anti-icing effectiveness by sealing half the upper-surface porous area. Gas flow through the leading-edge section had no appreciable effect on airfoil pressure distribution. The airfoil section drag increased slightly (5-percent average) with gas flow through the porous surface. A heavy glaze-ice formation produced after 10 minutes of icing caused an increase in section drag coefficient of 240 percent. Experimental convective heat-transfer coefficients obtained with hot-gas flow through the porous area in dry air and turbulent flow were 20 to 30 percent lower than the theoretical values for a solid surface under similar conditions. The transition region from laminar to turbulent flow moved forward as the ratio of gas velocity through the porous surface to air-stream velocity was increased.

  5. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Astrophysics Data System (ADS)

    Nirmalan, V.; Hylton, L. D.

    1989-06-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  6. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Technical Reports Server (NTRS)

    Nirmalan, V.; Hylton, L. D.

    1989-01-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  7. Design & fabrication of two seated aircraft with an advanced rotating leading edge wing

    NASA Astrophysics Data System (ADS)

    Al Ahmari, Saeed Abdullah Saeed

    The title of this thesis is "Design & Fabrication of two Seated Aircraft with an Advanced Rotating Leading Edge Wing", this gives almost a good description of the work has been done. In this research, the moving surface boundary-layer control (MSBC) concept was investigated and implemented. An experimental model was constructed and tested in wind tunnel to determine the aerodynamic characteristics using the leading edge moving surface of modified semi-symmetric airfoil NACA1214. The moving surface is provided by a high speed rotating cylinder, which replaces the leading edge of the airfoil. The angle of attack, the cylinder surfaces velocity ratio Uc/U, and the flap deflection angle effects on the lift and drag coefficients and the stall angle of attack were investigated. This new technology was applied to a 2-seat light-sport aircraft that is designed and built in the Aerospace Engineering Department at KFUPM. The project team is led by the aerospace department chairman Dr. Ahmed Z. AL-Garni and Dr. Wael G. Abdelrahman and includes graduate and under graduate student. The wing was modified to include a rotating cylinder along the leading edge of the flap portion. This produced very promising results such as the increase of the maximum lift coefficient at Uc/U=3 by 82% when flaps up and 111% when flaps down at 40° and stall was delayed by 8degrees in both cases. The laboratory results also showed that the effective range of the leading-edge rotating cylinder is at low angles of attack which reduce the need for higher angles of attack for STOL aircraft.

  8. 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.

  9. 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.

  10. 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.

  11. The adverse aerodynamic impact of very small leading-edge ice (roughness) buildups on wings and tails

    NASA Technical Reports Server (NTRS)

    Lynch, Frank T.; Valarezo, Walter O.; Mcghee, Robert J.

    1991-01-01

    Systematic experimental studies were performed to establish the aerodynamic impact of very small leading-edge simulated ice (roughness) formations on lifting surfaces. The geometries studied include single element configurations (airfoil and 3-D tail) as well as multi-element high-lift airfoil geometries. Emphasis in these studies was placed on obtaining results at high Reynolds numbers to insure the applicability of the findings to full-scale situations. It was found that the well-known Brumby correlation for the adverse lift impact of discrete roughness elements at the leading edge is not appropriate for cases representative of initial ice build up (i.e., distributed roughness). It was also found that allowing initial ice formations of a size required for removal by presently proposed deicing systems could lead to maximum lift losses of approximately 40 percent for single-element airfoils. Losses in angle-of-attack margin to stall are equally substantial - as high as 6 degrees. Percentage losses for multi-element airfoils are not as severe as for single-element configurations, but degradations of the angle-of-attack-to-stall margin are the same for both.

  12. 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.

  13. Multiple laminar-turbulent transition cycles around a swept leading edge

    NASA Astrophysics Data System (ADS)

    Mukund, R.; Narasimha, R.; Viswanath, P. R.; Crouch, J. D.

    2012-12-01

    Certain interesting flow features involving multiple transition/relaminarization cycles on the leading edge of a swept wing at low speeds are reported here. The wing geometry tested had a circular nose and a leading edge sweep of 60°. Tests were made at a chord Reynolds number of 1.3 × 106 with model incidence α varied in the range of 3°-18° in discrete steps. Measurements made included wing chord-wise surface pressure distributions and wall shear stress fluctuations (using hot-film gages) within about 10 % of the chord in the leading edge zone. Results at α = 16° and 18° showed that several (often incomplete) transition cycles between laminar-like and turbulent-like flows occurred. These rather surprising results are attributable chiefly to the fact that the Launder acceleration parameter K (appropriately modified for swept wings) can exceed a critical range more than once along the contour of the airfoil in the leading edge region. Each such crossing results in a relaminarization followed by direct retransition to turbulence as K drops to sufficiently low values. It is further shown that the extent of each observed transition zone (of either type) is consistent with earlier data acquired in more detailed studies of direct transition and relaminarization. Swept leading edge boundary layers therefore pose strong challenges to numerical modelling.

  14. 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.

  15. 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.

  16. Wind-Tunnel Tests on Various Types of Dive Brakes Mounted in Proximity of the Leading Edge of the Wing

    NASA Technical Reports Server (NTRS)

    Lattanzi, Bernardino; Bellante, Erno

    1949-01-01

    The present report is concerned with a series of tests on a model airplane fitted with four types of dive flaps of various shapes, positions, and incidence located near the leading edge of the wing (from 5 to 20 percent of the wing chord). Tests were also made on a stub airfoil fitted with a ventral dive (located at 8 percent of the wing chord). The hinge moments of the dive flaps were measured.

  17. Wind-tunnel studies of advanced cargo aircraft concepts. [leading edge vortex flaps for drag reduction

    NASA Technical Reports Server (NTRS)

    Rao, D. M.; Goglia, G. L.

    1981-01-01

    Accomplishments in vortex flap research are summarized. A singular feature of the vortex flap is that, throughout the range of angle of attack range, the flow type remains qualitatively unchanged. Accordingly, no large or sudden change in the aerodynamic characteristics, as happens when forcibly maintained attached flow suddenly reverts to separation, will occur with the vortex flap. Typical wind tunnel test data are presented which show the drag reduction potential of the vortex flap concept applied to a supersonic cruise airplane configuration. The new technology offers a means of aerodynamically augmenting roll-control effectiveness on slender wings at higher angles of attack by manipulating the vortex flow generated from leading edge separation. The proposed manipulator takes the form of a flap hinged at or close to the leading edge, normally retracted flush with the wing upper surface to conform to the airfoil shape.

  18. Textbook Multigrid Efficiency for Leading Edge Stagnation

    NASA Technical Reports Server (NTRS)

    Diskin, Boris; Thomas, James L.; Mineck, Raymond E.

    2004-01-01

    A multigrid solver is defined as having textbook multigrid efficiency (TME) if the solutions to the governing system of equations are attained in a computational work which is a small (less than 10) multiple of the operation count in evaluating the discrete residuals. TME in solving the incompressible inviscid fluid equations is demonstrated for leading- edge stagnation flows. The contributions of this paper include (1) a special formulation of the boundary conditions near stagnation allowing convergence of the Newton iterations on coarse grids, (2) the boundary relaxation technique to facilitate relaxation and residual restriction near the boundaries, (3) a modified relaxation scheme to prevent initial error amplification, and (4) new general analysis techniques for multigrid solvers. Convergence of algebraic errors below the level of discretization errors is attained by a full multigrid (FMG) solver with one full approximation scheme (F.4S) cycle per grid. Asymptotic convergence rates of the F.4S cycles for the full system of flow equations are very fast, approaching those for scalar elliptic equations.

  19. Textbook Multigrid Efficiency for Leading Edge Stagnation

    NASA Technical Reports Server (NTRS)

    Diskin, Boris; Thomas, James L.; Mineck, Raymond E.

    2004-01-01

    A multigrid solver is defined as having textbook multigrid efficiency (TME) if the solutions to the governing system of equations are attained in a computational work which is a small (less than 10) multiple of the operation count in evaluating the discrete residuals. TME in solving the incompressible inviscid fluid equations is demonstrated for leading-edge stagnation flows. The contributions of this paper include (1) a special formulation of the boundary conditions near stagnation allowing convergence of the Newton iterations on coarse grids, (2) the boundary relaxation technique to facilitate relaxation and residual restriction near the boundaries, (3) a modified relaxation scheme to prevent initial error amplification, and (4) new general analysis techniques for multigrid solvers. Convergence of algebraic errors below the level of discretization errors is attained by a full multigrid (FMG) solver with one full approximation scheme (FAS) cycle per grid. Asymptotic convergence rates of the FAS cycles for the full system of flow equations are very fast, approaching those for scalar elliptic equations.

  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. 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.

  2. 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.

  3. Laminar Flow Control Leading Edge Systems in Simulated Airline Service

    NASA Technical Reports Server (NTRS)

    Wagner, R. D.; Maddalon, D. V.; Fisher, D. F.

    1988-01-01

    Achieving laminar flow on the wings of a commercial transport involves difficult problems associated with the wing leading edge. The NASA Leading Edge Flight Test Program has made major progress toward the solution of these problems. The effectiveness and practicality of candidate laminar flow leading edge systems were proven under representative airline service conditions. This was accomplished in a series of simulated airline service flights by modifying a JetStar aircraft with laminar flow leading edge systems and operating it out of three commercial airports in the United States. The aircraft was operated as an airliner would under actual air traffic conditions, in bad weather, and in insect infested environments.

  4. Development of X-43A Mach 10 Leading Edges

    NASA Technical Reports Server (NTRS)

    Ohlhorst, Craig W.; Glass, David E.; Bruce, Walter E., III; Lindell, Michael C.; Vaughn, Wallace L.; Dirling, R. B., Jr.; Hogenson, P. A.; Nichols, J. M.; Risner, N. W.; Thompson, D. R.

    2005-01-01

    The nose leading edge of the Hyper-X Mach 10 vehicle was orginally anticipated to reach temperatures near 4000 F at the leading-edge stagnation line. A SiC coated carbon/carbon (C/C) leading-edge material will not survive that extreme temperature for even a short duration single flight. To identify a suitable leading edge for the Mach 10 vehicle, arc-jet testing was performed on thirteen leading-edge segments fabricated from different material systems to evaluate their performance in a simulated flight environment. Hf, Zr, Si, and Ir based materials, in most cases as a coating on C/C, were included in the evaluation. Afterwards, MER, Tucson, AZ was selected as the supplier of the flight vehicle leading edges. The nose and the vertical and horizontal tail leading edges were fabricated out of a 3:1 biased high thermal conductivity C/C. The leading edges were coated with a three layer coating comprised of a SiC conversion of the top surface of the C/C, followed by a chemical vapor deposited layer of SiC, followed by a thin chemical vapor deposited layer of HfC. This paper will describe the fabrication of the Mach 10 C/C leading edges and the testing performed to validate performance.

  5. The artificially blunted leading edge concept for aerothermodynamic performance enhancement

    NASA Astrophysics Data System (ADS)

    Gupta, Anurag

    An innovative aerothermodynamic performance enhancement concept for blunted geometries in hypervelocity flight is described. An Artificially Blunted Leading Edge (ABLE) is sought to be created by the use of a flow-through channel sized to choke at supersonic (in the normal direction) conditions. As a result, a normal shock stands off the channel but the high post-shock pressures have no wall to act on, leading to a reduction in wave drag. The effective blunt body flow structure can be effective at preventing the rise in heat transfer rates at channel entrance lips. In lifting flight, the flow in the channel creates suction at the lip, significantly enhancing lift for non-slender shapes. CFD studies using Reynolds Averaged Navier-Stokes simulations provide proof-of- concept for drag reduction for blunted slender geometries and L/D enhancements for sphere-cones. The ABLE flow mechanism's robustness and its effectiveness at off- design conditions is demonstrated. The computed sphere- cone L/D enhancements are also validated with experimental results from Aeroballistic Range tests. As opposed to straight channels, ABLE variants with curved channels that provide for better volumetric efficiency, reduced viscous drag penalties and better performance were designed and investigated. The channels curve outward and exhaust the flow close to the leading edge. Even while exhausting tangentially, the exhaust-mean flow interactions were shown to enhance or create lift. The force amplification due to such interactions can also be leveraged with the channel flow exhausting nearly normal to the surface. The potential of such thrust vectoring to reduce trim drag and augment directional control in the high-speed regime was demonstrated numerically. To evaluate the concept's effectiveness at improving cd or L/D values without paying any penalties in lift, enclosed volume and peak heating rates, Multidisciplinary Design Optimization techniques are used to characterize the design space

  6. Numerical study of delta wing leading edge blowing

    NASA Technical Reports Server (NTRS)

    Yeh, David; Tavella, Domingo; Roberts, Leonard

    1988-01-01

    Spanwise and tangential leading edge blowing as a means of controlling the position and strength of the leading edge vortices are studied by numerical solution of the three-dimensional Navier-Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of spanwise leading edge blowing at moderate angle of attack magnifies the size and strength of the leading edge vortices, and moves the vortex cores outboard and upward. The increase in lift primarily comes from the greater nonlinear vortex lift. However, spanwise blowing causes earlier vortex breakdown, thus decreasing the stall angle. The effects of tangential blowing at low to moderate angles of attack tend to reduce the pressure peaks associated with leading edge vortices and to increase the suction peak around the leading edge, so that the integrated value of the surface pressure remains about the same. Tangential leading edge blowing in post-stall conditions is shown to re-establish vortical flow and delay vortex bursting, thus increasing C sub L sub max and stall angle.

  7. Heat-Pipe-Cooled Leading Edges for Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    2006-01-01

    Heat pipes can be used to effectively cool wing leading edges of hypersonic vehicles. . Heat-pipe leading edge development. Design validation heat pipe testing confirmed design. Three heat pipes embedded and tested in C/C. Single J-tube heat pipe fabricated and testing initiated. HPCLE work is currently underway at several locations.

  8. 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.

  9. Timing discriminator using leading-edge extrapolation

    DOEpatents

    Gottschalk, B.

    1981-07-30

    A discriminator circuit to recover timing information from slow-rising pulses by means of an output trailing edge, a fixed time after the starting corner of the input pulse, which is nearly independent of risetime and threshold setting is described. This apparatus comprises means for comparing pulses with a threshold voltage; a capacitor to be charged at a certain rate when the input signal is one-third threshold voltage, and at a lower rate when the input signal is two-thirds threshold voltage; current-generating means for charging the capacitor; means for comparing voltage capacitor with a bias voltage; a flip-flop to be set when the input pulse reaches threshold voltage and reset when capacitor voltage reaches the bias voltage; and a clamping means for discharging the capacitor when the input signal returns below one-third threshold voltage.

  10. Timing discriminator using leading-edge extrapolation

    DOEpatents

    Gottschalk, Bernard

    1983-01-01

    A discriminator circuit to recover timing information from slow-rising pulses by means of an output trailing edge, a fixed time after the starting corner of the input pulse, which is nearly independent of risetime and threshold setting. This apparatus comprises means for comparing pulses with a threshold voltage; a capacitor to be charged at a certain rate when the input signal is one-third threshold voltage, and at a lower rate when the input signal is two-thirds threshold voltage; current-generating means for charging the capacitor; means for comparing voltage capacitor with a bias voltage; a flip-flop to be set when the input pulse reaches threshold voltage and reset when capacitor voltage reaches the bias voltage; and a clamping means for discharging the capacitor when the input signal returns below one-third threshold voltage.

  11. Application of a flush airdata sensing system to a wing leading edge (LE-FADS)

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Czerniejewski, Mark W.; Nichols, Douglas A.

    1993-01-01

    The feasibility of locating a flush airdata sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered is investigated. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the airdata calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush airdata systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 deg to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

  12. Vortex leading edge flap assembly for supersonic airplanes

    NASA Technical Reports Server (NTRS)

    Rudolph, Peter K. C. (Inventor)

    1997-01-01

    A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

  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. Structural Health Monitoring Analysis for the Orbiter Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Yap, Keng C.

    2010-01-01

    This viewgraph presentation reviews Structural Health Monitoring Analysis for the Orbiter Wing Leading Edge. The Wing Leading Edge Impact Detection System (WLE IDS) and the Impact Analysis Process are also described to monitor WLE debris threats. The contents include: 1) Risk Management via SHM; 2) Hardware Overview; 3) Instrumentation; 4) Sensor Configuration; 5) Debris Hazard Monitoring; 6) Ascent Response Summary; 7) Response Signal; 8) Distribution of Flight Indications; 9) Probabilistic Risk Analysis (PRA); 10) Model Correlation; 11) Impact Tests; 12) Wing Leading Edge Modeling; 13) Ascent Debris PRA Results; and 14) MM/OD PRA Results.

  15. Sharp Refractory Composite Leading Edges on Hypersonic Vehicles

    NASA Technical Reports Server (NTRS)

    Walker, Sandra P.; Sullivan, Brian J.

    2003-01-01

    On-going research of advanced sharp refractory composite leading edges for use on hypersonic air-breathing vehicles is presented in this paper. Intense magnitudes of heating and of heating gradients on the leading edge lead to thermal stresses that challenge the survivability of current material systems. A fundamental understanding of the problem is needed to further design development. Methodology for furthering the technology along with the use of advanced fiber architectures to improve the thermal-structural response is explored in the current work. Thermal and structural finite element analyses are conducted for several advanced fiber architectures of interest. A tailored thermal shock parameter for sharp orthotropic leading edges is identified for evaluating composite material systems. The use of the tailored thermal shock parameter has the potential to eliminate the need for detailed thermal-structural finite element analyses for initial screening of material systems being considered for a leading edge component.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Ultra-High Temperature Ceramic Composites for Leading Edges

    NASA Technical Reports Server (NTRS)

    Levine, Stanley R.

    2004-01-01

    Issues associated with the development and use of Ultra-High Temperature Ceramic Composites (UHTCC) for leading edges of hypersonic vehicles will be discussed. These include attachments, constituent selection, processing, oxidation, physical and mechanical properties, and attachments.

  1. Shock Interaction Control for Scramjet Cowl Leading Edges

    NASA Technical Reports Server (NTRS)

    Albertson, Cindy W.; Venkat, Venki, S.

    2005-01-01

    An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface.

  2. A Thermostructural Analysis of a Diboride Composite Leading Edge

    NASA Technical Reports Server (NTRS)

    Kowalski, Tom; Buesking, Kent; Kolodziej, Paul; Bull, Jeff

    1996-01-01

    In an effort to support the design of zirconium diboride composite leading edges for hypersonic vehicles, a finite element model (FEM) of a prototype leading edge was created and finite element analysis (FEA) was employed to assess its thermal and structural response to aerothermal boundary conditions. Unidirectional material properties for the structural components of the leading edge, a continuous fiber reinforced diboride composite, were computed with COSTAR. These properties agree well with those experimentally measured. To verify the analytical approach taken with COSMOS/M, an independent FEA of one of the leading edge assembly components was also done with COSTAR. Good agreement was obtained between the two codes. Both showed that a unidirectional lay-up had the best margin of safety for a simple loading case. Both located the maximum stress in the same region and ply. The magnitudes agreed within 4 percent. Trajectory based aerothermal heating was then applied to the leading edge assembly FEM created with COSMOS/M to determine steady state temperature response, displacement, stresses, and contact forces due to thermal expansion and thermal strains. Results show that the leading edge stagnation line temperature reached 4700 F. The maximum computed failure index for the laminated composite components peaks at 4.2, and is located at the bolt flange in layer 2 of the side bracket. The temperature gradient in the tip causes a compressive stress of 279 ksi along its width and substantial tensile stresses within its depth.

  3. 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.

  4. Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile development. Leading edge material development and testing

    NASA Astrophysics Data System (ADS)

    Hogenson, P. A.; Staszak, Paul; Hinkle, Karrie

    1995-05-01

    This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge configuration; and life cycle testing.

  5. Task 4 supporting technology. Part 1: Detailed test plan for leading edge tile development. Leading edge material development and testing

    NASA Technical Reports Server (NTRS)

    Hogenson, P. A.; Staszak, Paul; Hinkle, Karrie

    1995-01-01

    This task develops two alternative candidate tile materials for leading edge applications: coated alumina enhanced thermal barrier (AETB) tile and silicone impregnated reusable ceramic ablator (SIRCA) tile. Upon reentry of the X-33/RLV space vehicle, the leading edges experience the highest heating rates and temperatures. The wing leading edge and nose cap experience peak temperatures in the range 2000 to 2700 F. Replacing reinforced carbon-carbon (RCC) with tile-based thermal protection system (TPS) materials is the primary objective. Weight, complexity, coating impact damage, and repairability are among the problems that this tile technology development addresses. The following subtasks will be performed in this development effort: tile coating development; SIRCA tile development; robustness testing of tiles; tile repair development; tile operations/processing; tile leading edge configuration; and life cycle testing.

  6. 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.

  7. 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.

  8. 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.

  9. Low-speed cascade investigation of loaded leading-edge compressor blades

    NASA Technical Reports Server (NTRS)

    Emery, James C

    1956-01-01

    Six percent thick NACA 63-series compressor-blade sections having a loaded leading-edge A4K6 mean line have been investigated systematically in a two-dimensional porous-wall cascade over a range of Reynolds numbers from 160,000 to 385,000. Blades cambered to have isolated-airfoil lift coefficients of 0.6, 1.2, 1.8, and 2.4 were tested over the usable angle-of-attack range at inlet-air angles of 30 degrees, 45 degrees, and 60 degrees and solidities of 1.0 and 1.5. A comparison with data of NACA RM L51G31, shows that the angle-of-attack operating range is 2 degrees to 4 degrees less than the range for the uniformly loaded section; however, the wake losses near design angle of attack are slightly lower than those for the uniformly loaded section. Except for highly cambered blades at high inlet angles, the 63-(C s oA4K6)06 compressor-blade sections are capable of more efficient operation for moderate-speed subsonic compressors at design angle of attack than are the 65-(C s oa10)10 or the 65-(c s oA2I8b)10 compressor-blade sections. In contrast to the other sections, the loaded leading-edge sections are capable of operating efficiently at the lower Reynolds numbers.

  10. 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.

  11. 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.

  12. Fluid-thermal-structural study of aerodynamically heated leading edges

    NASA Technical Reports Server (NTRS)

    Deuchamphai, Pramote; Thornton, Earl A.; Wieting, Allan R.

    1988-01-01

    A finite element approach for integrated fluid-thermal-structural analysis of aerodynamically heated leading edges is presented. The Navier-Stokes equations for high speed compressible flow, the energy equation, and the quasi-static equilibrium equations for the leading edge are solved using a single finite element approach in one integrated, vectorized computer program called LIFTS. The fluid-thermal-structural coupling is studied for Mach 6.47 flow over a 3-in diam cylinder for which the flow behavior and the aerothermal loads are calibrated by experimental data. Issues of the thermal-structural response are studied for hydrogen-cooled, super thermal conducting leading edges subjected to intense aerodynamic heating.

  13. Design and Analysis of UHTC Leading Edge Attachment

    NASA Technical Reports Server (NTRS)

    Thomas, David J.; Nemeth, Noel N. (Technical Monitor)

    2002-01-01

    NASA Glenn Research Center was contacted to provide technical support to NASA Ames Research Center in the design and analysis of an ultra high temperature ceramic (UHTC) leading edge. UHTC materials are being considered for reusable launch vehicles because their high temperature capability may allow for un-cooled sharp leading edge designs. While ceramic materials have the design benefit of allowing subcomponents to run hot, they also provide a design challenge in that they invariably must be in contact with cooler subcomponents elsewhere in the structure. NASA Glenn Research Center proposed a modification to an existing attachment design. Thermal and structural analyses of the leading edge assembly were carried out using ABAQUS finite element software. Final results showed that the proposed modifications aided in thermally isolating hot and cold subcomponents and reducing bearing stresses at the attachment location.

  14. Effect of Leading Edge Tubercles on Marine Tidal Turbine Blades

    NASA Astrophysics Data System (ADS)

    Murray, Mark; Gruber, Timothy; Fredriksson, David

    2010-11-01

    This project investigated the impact that the addition of leading edge protuberances (tubercles) have on the effectiveness of marine tidal turbine blades, especially at lower flow speeds. The addition of leading edge tubercles to lifting foils has been shown, in previous research, to delay the onset of stall without significant hydrodynamic costs. The experimental results obtained utilizing three different blade designs (baseline and two tubercle modified) are compared. All blades were designed in SolidWorks and manufactured utilizing rapid prototype techniques. All tests were conducted in the 120 ft tow tank at the U.S. Naval Academy using a specifically designed experimental apparatus. Results for power coefficients are presented for a range of tip speed ratios. Cut-in velocity is also compared between the blade designs. For all test criteria, the tubercle modified blades significantly outperformed the smooth leading edge baseline design blades.

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

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

  16. Contact angle at the leading edge controls cell protrusion rate.

    PubMed

    Gabella, Chiara; Bertseva, Elena; Bottier, Céline; Piacentini, Niccolò; Bornert, Alicia; Jeney, Sylvia; Forró, László; Sbalzarini, Ivo F; Meister, Jean-Jacques; Verkhovsky, Alexander B

    2014-05-19

    Plasma membrane tension and the pressure generated by actin polymerization are two antagonistic forces believed to define the protrusion rate at the leading edge of migrating cells [1-5]. Quantitatively, resistance to actin protrusion is a product of membrane tension and mean local curvature (Laplace's law); thus, it depends on the local geometry of the membrane interface. However, the role of the geometry of the leading edge in protrusion control has not been yet investigated. Here, we manipulate both the cell shape and substrate topography in the model system of persistently migrating fish epidermal keratocytes. We find that the protrusion rate does not correlate with membrane tension, but, instead, strongly correlates with cell roundness, and that the leading edge of the cell exhibits pinning on substrate ridges-a phenomenon characteristic of spreading of liquid drops. These results indicate that the leading edge could be considered a triple interface between the substrate, membrane, and extracellular medium and that the contact angle between the membrane and the substrate determines the load on actin polymerization and, therefore, the protrusion rate. Our findings thus illuminate a novel relationship between the 3D shape of the cell and its dynamics, which may have implications for cell migration in 3D environments.

  17. Leading-edge receptivity for blunt-nose bodies

    NASA Technical Reports Server (NTRS)

    Kerschen, Edward J.

    1991-01-01

    This research program investigates boundary-layer receptivity in the leading-edge region for bodies with blunt leading edges. Receptivity theory provides the link between the unsteady distrubance environment in the free stream and the initial amplitudes of the instability waves in the boundary layer. This is a critical problem which must be addressed in order to develop more accurate prediction methods for boundary-layer transition. The first phase of this project examines the effects of leading-edge bluntness and aerodynamic loading for low Mach number flows. In the second phase of the project, the investigation is extended to supersonic Mach numbers. Singular perturbation techniques are utilized to develop an asymptotic theory for high Reynolds numbers. In the first year, the asymptotic theory was developed for leading-edge receptivity in low Mach number flows. The case of a parabolic nose is considered. Substantial progress was made on the Navier-Sotkes computations. Analytical solutions for the steady and unsteady potential flow fields were incorporated into the code, greatly expanding the types of free-stream disturbances that can be considered while also significantly reducing the the computational requirements. The time-stepping algorithm was modified so that the potential flow perturbations induced by the unsteady pressure field are directly introduced throughout the computational domain, avoiding an artificial 'numerical diffusion' of these from the outer boundary. In addition, the start-up process was modified by introducing the transient Stokes wave solution into the downstream boundary conditions.

  18. Analysis of Bird Impact on a Composite Tailplane Leading Edge

    NASA Astrophysics Data System (ADS)

    Guida, M.; Marulo, F.; Meo, M.; Riccio, M.

    2008-11-01

    One of the main structural requirements of a leading edge of a tailplane is to ensure that any significant damage caused by foreign object (i.e. birdstrike, etc...) would still allow the aircraft to land safely. In particular, leading edge must be certified for a proven level of bird impact resistance. Since the experimental tests are expensive and difficult to perform, numerical simulations can provide significant help in designing high-efficiency bird-proof structures. The aim of this research paper was to evaluate two different leading edge designs by reducing the testing costs by employing state-of-the-art numerical simulations. The material considered was a sandwich structure made up of aluminium skins and flexcore as core. Before each test was carried out, pre-test numerical analyses of birdstrike were performed adopting a lagrangian approach on a tailplane leading edge of a large scale aircraft using the MSC/Dytran solver code. The numerical and experimental correlation have shown good results both in terms of global behaviour of the test article and local evolution of some measurable parameters confirming the validity of the approach and possible guidelines for structural design including the bird impact requirements.

  19. Vortex interaction with a leading-edge of finite thickness

    NASA Technical Reports Server (NTRS)

    Sohn, D.; Rockwell, Donald

    1987-01-01

    Vortex interaction with a thick elliptical leading-edge at zero relative offset produces a pronounced secondary vortes of opposite sense that travels with the same phase speed as the primaty vortex along the lower surface of the edge. The edge thickness (scale) relative to the incident vorticity field has a strong effect on the distortion of the incident primary vortex during the impingement processs. When the thickness is sufficiently small, there is a definite severing of the incident vortex and the portion of the incident vortex that travels along the upper part of the elliptical surface has a considerably larger phase speed than that along the lower surface; this suggests that the integrated loading along the upper surface is more strongly correlated. When the thickness becomes too large, then most, if not all, of the incident vortex passes below the leading-edge. On the other hand, the relative tranverse offset of the edge with respect to the center of the incident vortex has a significant effect on the secondary vortex formation.

  20. Membrane insertion at the leading edge of motile fibroblasts.

    PubMed Central

    Bergmann, J E; Kupfer, A; Singer, S J

    1983-01-01

    We are concerned with the mechanisms involved in the directed migration of eukaryotic cells. Previously we found that, inside cells at the edge of an experimental wound, the Golgi apparatus and the microtubule-organizing center were rapidly repositioned forward of the nucleus in the direction of subsequent cell migration into the wound. This repositioning was proposed to serve the purpose of introducing new membrane mass at the leading edge of the cell, by directing Golgi apparatus-derived vesicles bound for the plasma membrane to that edge. We now provide evidence to support this proposal. Cultured fibroblastic cells at the edge of a wound were infected with a temperature-sensitive mutant (0-45) of vesicular stomatitis virus. It is known that the G-protein, an integral membrane protein of the virus, is synthesized and remains in the rough endoplasmic reticulum at the nonpermissive temperature, but when the infected cells are shifted to the permissive temperature, the G-protein moves through the Golgi apparatus to the plasma membrane. By immunofluorescence microscopy, we here show that the first appearance of the G-protein at the cell surface corresponds to the leading edge of the motile cell. These observations are incorporated into a coherent scheme for the mechanisms involved in cell migration. Images PMID:6298789

  1. Detail view of the leading and top edge of the ...

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

    Detail view of the leading and top edge of the vertical stabilizer of the Orbiter Discovery showing the thermal protection system components with the white Advanced Flexible Reusable Surface Insulation (AFRSI) blanket and the black High-temperature Reusable Surface Insulation (HRSI) tiles along the outer edges. The marks seen on the HRSI tiles are injection point marks and holes for the application of waterproofing material. This view was taken from a service platform in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  2. A numerical and experimental study of the effects of dynamic roughness on laminar leading edge separation

    NASA Astrophysics Data System (ADS)

    Gall, Peter D.

    The aircraft industry, as a whole, has been deeply concerned with improving the aerodynamic efficiency of current and future flight vehicles, particularly in the commercial and military markets. However, of particular interest to the field of aerodynamics is the elusive concept of a workable flow control mechanism. Effective flow control is a concept which if properly applied can increase aerodynamic efficiency. Various concepts and ideas to obtain successful flow control have been studied in an attempt to reap these rewards. Some examples include boundary layer blowing (steady and periodic), suction, and compliant walls for laminar flow control. The overall goal of flow control is to increase performance by increasing lift, reducing drag, and delaying or eliminating leading edge separation. The specific objectives of flow control are to (1) delay or eliminate flow separation, (2) delay boundary layer transition, and (3) and reduce skin friction drag. The purpose of this research is to investigate dynamic roughness as a novel method of flow control technology for external boundary layer flows. As opposed to standard surface roughness, dynamic roughness incorporates small time dependent perturbations to the surface of the airfoil. These surface perturbations are actual humps and/or ridges on the surface of the airfoil that are on the scale of the laminar boundary, and oscillate with an unsteady motion. Research has shown that this can provide a means to modify the instantaneous and mean velocity profile near the wall and favorably control the existing state of the boundary layer. Several flow control parameters were studied including dynamic roughness frequency, amplitude, and geometry. The results of this study have shown, both numerically and experimentally, that dynamic roughness can provide an effective means for eliminating both a short and long laminar separation bubble and possibly prove a viable alternative in effective flow control, hence reaping some of

  3. 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.

  4. 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.

  5. Turbine airfoil fabricated from tapered extrusions

    DOEpatents

    Marra, John J

    2013-07-16

    An airfoil (30) and fabrication process for turbine blades with cooling channels (26). Tapered tubes (32A-32D) are bonded together in a parallel sequence, forming a leading edge (21), a trailing edge (22), and pressure and suction side walls (23, 24) connected by internal ribs (25). The tapered tubes may be extruded without camber to simplify the extrusion process, then bonded along matching surfaces (34), forming a non-cambered airfoil (28), which may be cambered in a hot forming process and cut (48) to length. The tubes may have tapered walls that are thinner at the blade tip (T1) than at the base (T2), reducing mass. A cap (50) may be attached to the blade tip. A mounting lug (58) may be forged (60) on the airfoil base and then machined, completing the blade for mounting in a turbine rotor disk.

  6. The fish tail motion forms an attached leading edge vortex.

    PubMed

    Borazjani, Iman; Daghooghi, Mohsen

    2013-04-01

    The tail (caudal fin) is one of the most prominent characteristics of fishes, and the analysis of the flow pattern it creates is fundamental to understanding how its motion generates locomotor forces. A mechanism that is known to greatly enhance locomotor forces in insect and bird flight is the leading edge vortex (LEV) reattachment, i.e. a vortex (separation bubble) that stays attached at the leading edge of a wing. However, this mechanism has not been reported in fish-like swimming probably owing to the overemphasis on the trailing wake, and the fact that the flow does not separate along the body of undulating swimmers. We provide, to our knowledge, the first evidence of the vortex reattachment at the leading edge of the fish tail using three-dimensional high-resolution numerical simulations of self-propelled virtual swimmers with different tail shapes. We show that at Strouhal numbers (a measure of lateral velocity to the axial velocity) at which most fish swim in nature (approx. 0.25) an attached LEV is formed, whereas at a higher Strouhal number of approximately 0.6 the LEV does not reattach. We show that the evolution of the LEV drastically alters the pressure distribution on the tail and the force it generates. We also show that the tail's delta shape is not necessary for the LEV reattachment and fish-like kinematics is capable of stabilising the LEV. Our results suggest the need for a paradigm shift in fish-like swimming research to turn the focus from the trailing edge to the leading edge of the tail.

  7. 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.

  8. Nondestructive Evaluation for the Space Shuttle's Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Madaras, Eric I.; Winfree, William P.; Prosser, William H.; Wincheski, Russell A.; Cramer, K. Elliot

    2005-01-01

    The loss of the Space Shuttle Columbia highlighted concerns about the integrity of the Shuttle's thermal protection system, which includes Reinforced Carbon-Carbon (RCC) on the leading edge. This led NASA to investigate nondestructive evaluation (NDE) methods for certifying the integrity of the Shuttle's wing leading edge. That investigation was performed simultaneously with a large study conducted to understand the impact damage caused by errant debris. Among the many advanced NDE methods investigated for applicability to the RCC material, advanced digital radiography, high resolution computed tomography, thermography, ultrasound, acoustic emission and eddy current systems have demonstrated the maturity and success for application to the Shuttle RCC panels. For the purposes of evaluating the RCC panels while they are installed on the orbiters, thermographic detection incorporating principal component analysis (PCA) and eddy current array scanning systems demonstrated the ability to measure the RCC panels from one side only and to detect several flaw types of concern. These systems were field tested at Kennedy Space Center (KSC) and at several locations where impact testing was being conducted. Another advanced method that NASA has been investigating is an automated acoustic based detection system. Such a system would be based in part on methods developed over the years for acoustic emission testing. Impact sensing has been demonstrated through numerous impact tests on both reinforced carbon-carbon (RCC) leading edge materials as well as Shuttle tile materials on representative aluminum wing structures. A variety of impact materials and conditions have been evaluated including foam, ice, and ablator materials at ascent velocities as well as simulated hypervelocity micrometeoroid and orbital debris impacts. These tests have successfully demonstrated the capability to detect and localize impact events on Shuttle's wing structures. A first generation impact sensing

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Probabilistic Structural Health Monitoring of the Orbiter Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Yap, Keng C.; Macias, Jesus; Kaouk, Mohamed; Gafka, Tammy L.; Kerr, Justin H.

    2011-01-01

    A structural health monitoring (SHM) system can contribute to the risk management of a structure operating under hazardous conditions. An example is the Wing Leading Edge Impact Detection System (WLEIDS) that monitors the debris hazards to the Space Shuttle Orbiter s Reinforced Carbon-Carbon (RCC) panels. Since Return-to-Flight (RTF) after the Columbia accident, WLEIDS was developed and subsequently deployed on board the Orbiter to detect ascent and on-orbit debris impacts, so as to support the assessment of wing leading edge structural integrity prior to Orbiter re-entry. As SHM is inherently an inverse problem, the analyses involved, including those performed for WLEIDS, tend to be associated with significant uncertainty. The use of probabilistic approaches to handle the uncertainty has resulted in the successful implementation of many development and application milestones.

  15. Heat pipes for wing leading edges of hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Boman, B. L.; Citrin, K. M.; Garner, E. C.; Stone, J. E.

    1990-01-01

    Wing leading edge heat pipes were conceptually designed for three types of vehicle: an entry research vehicle, aero-space plane, and advanced shuttle. A full scale, internally instrumented sodium/Hastelloy X heat pipe was successfully designed and fabricated for the advanced shuttle application. The 69.4 inch long heat pipe reduces peak leading edge temperatures from 3500 F to 1800 F. It is internally instrumented with thermocouples and pressure transducers to measure sodium vapor qualities. Large thermal gradients and consequently large thermal stresses, which have the potential of limiting heat pipe life, were predicted to occur during startup. A test stand and test plan were developed for subsequent testing of this heat pipe. Heat pipe manufacturing technology was advanced during this program, including the development of an innovative technique for wick installation.

  16. 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.

  17. 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.

  18. Leading-edge vortex solutions with large total pressure losses

    NASA Technical Reports Server (NTRS)

    Murman, Earll M.; Powell, Kenneth G.; Goodsell, Aga M.; Landahl, Marten T.

    1987-01-01

    Computations are presented for a Lambda = 75 deg delta wing in a supersonic freestream under two conditions which lead to leading-edge vortices. For one condition, analysis of the computed vortical flow reveals a closed streamline in the core. From varying computational parameters, it appears that this is due to truncation error of the convective derivatives. For the other condition, comparisons are made with wind-tunnel data, and good agreement is noted for pitot pressure distributions, flow angles on the symmetry plane, and the position of an embedded shock. Many of the aerodynamic parameters are shown to be insensitive to grid spacing.

  19. Optimization of the leading edge segment of a corrugated wing

    NASA Astrophysics Data System (ADS)

    Khurana, Manas; Chahl, Javaan

    2014-03-01

    Insect wings consist of flat plates of membranes stiffened by spars. The effect of this structure is that the wings appear as corrugated surfaces when considered on chordwise sections. We know that aerodynamically efficient insects such as a dragonfly engage in fixed wing flight modes for extended periods. The analysis in the literature has shown that the aerodynamic efficiency (cl/cd) of a corrugated aerofoil is sensitive to Reynolds number (Re) and angle-of-attack (AoA), yet the conclusions established are on the basis of flow analysis on a single baseline shape only. The sample size of the aerofoils must be extended further so that the influence and merits of corrugated shape features can be established. In this work, a design-of-experiments (DoE) approach is applied to induce systematic shape perturbations on a select, off-the-shelf baseline shape one feature at a time over a set number of increments. At each shape increment, the aerodynamic forces are established using a high fidelity CFD solver. The design space is modeled at a Re of 20,000 and 34,000 and at flow angle of 4.0° to represent a Micro Air Vehicle (MAV) in glide. The results confirmed the importance of the leading and trailing edge deflections on cl/cd. At Re = 20, 000, cl/cd of a corrugated aerofoil with deflection at the leading edge region only is 16% higher than the baseline shape, and 39% higher than the flat plate. At Re = 34, 000, cl/cd performance is sensitive to the trailing edge deflection. At the optimum deflection setting, cl/cd is 18% higher than the baseline shape and 23% higher than the flat plate. The results confirm that the leading and trailing edge deflections are critical to cl/cd for a MAV in glide.

  20. 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.

  1. Structure of leading-edge vortex flows including vortex breakdown

    SciTech Connect

    Payne, F.M.

    1987-01-01

    An experimental investigation of the structure of leading-edge vortex flows on thin sharp-edged delta wings was carried out at low Reynolds numbers. Flow-visualization techniques were used to study the topology of the vortex and the phenomenon of vortex breakdown. Seven-hole probe-wake surveys and laser-doppler-anemometer measurements were obtained and compared. Delta wings with sweep angles of 70, 75, 80, and 85/sup 0/ were tested at angles of attack of 10, 20, 30, and 40/sup 0/. The test were conducted in a Reynolds number range of 8.5 x 10/sup 4/ to 6.4 x 10/sup 5/. Smoke-flow visualization revealed the presence of small Kelvin-Helmholtz type vortical structures in the shear layer of a leading-edge vortex. These shear-layer vortices follow a helical path and grow in the streamwise direction as they wind into the vortex core where the individual shear layers merge. The phenomenon of vortex breakdown was studied using high-speed cinema photography. The bubble and spiral types of breakdown were observed and appear to represent the extremes in a continuum of breakdown forms.

  2. Insect Residue Contamination on Wing Leading Edge Surfaces: A Materials Investigation for Mitigation

    NASA Technical Reports Server (NTRS)

    Lorenzi, Tyler M.; Wohl, Christopher J.; Penner, Ronald K.; Smith, Joseph G.; Siochi, Emilie J.

    2011-01-01

    Flight tests have shown that residue from insect strikes on aircraft wing leading edge surfaces may induce localized transition of laminar to turbulent flow. The highest density of insect populations have been observed between ground level and 153 m during light winds (2.6 -- 5.1 m/s), high humidity, and temperatures from 21 -- 29 C. At a critical residue height, dependent on the airfoil and Reynolds number, boundary layer transition from laminar to turbulent results in increased drag and fuel consumption. Although this represents a minimal increase in fuel burn for conventional transport aircraft, future aircraft designs will rely on maintaining laminar flow across a larger portion of wing surfaces to reduce fuel burn during cruise. Thus, insect residue adhesion mitigation is most critical during takeoff and initial climb to maintain laminar flow in fuel-efficient aircraft configurations. Several exterior treatments investigated to mitigate insect residue buildup (e.g., paper, scrapers, surfactants, flexible surfaces) have shown potential; however, implementation has proven to be impractical. Current research is focused on evaluation of wing leading edge surface coatings that may reduce insect residue adhesion. Initial work under NASA's Environmentally Responsible Aviation Program focused on evaluation of several commercially available products (commercial off-the-shelf, COTS), polymers, and substituted alkoxy silanes that were applied to aluminum (Al) substrates. Surface energies of these coatings were determined from contact angle data and were correlated to residual insect excrescence on coated aluminum substrates using a custom-built "bug gun." Quantification of insect excrescence surface coverage was evaluated by a series of digital photographic image processing techniques.

  3. 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.

  4. An Aeroacoustic Study of a Leading Edge Slat Configuration

    NASA Technical Reports Server (NTRS)

    Mendoza, J. M.; Brooks, T. F.; Humphreys, W. M., Jr.

    2002-01-01

    Aeroacoustic evaluations of high-lift devices have been carried out in the Quiet Flow Facility of the NASA Langley Research Center. The present paper describes detailed flow and acoustic measurements that have been made in order to better understand the noise generated from airflow over a wing leading edge slat configuration, and to possibly predict and reduce this noise source. The acoustic database is obtained by a moveable Small Aperture Directional Array of microphones designed to electronically steer to different portions of models under study. The slat is shown to be a uniform distributed noise source. The data was processed such that spectra and directivity were determined with respect to a one-foot span of slat. The spectra are normalized in various fashions to demonstrate slat noise character. In order to equate portions of the spectra to different slat noise components, trailing edge noise predictions using measured slat boundary layer parameters as inputs are compared to the measured slat noise spectra.

  5. Flexible Plug Repair for Shuttle Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Camarda, Charles J.; Sikora, Joseph; Smith, Russel; Rivers, H.; Scotti, Stephen J.; Fuller, Alan M.; Klacka, Robert; Reinders, Martin; Schwind, Francis; Sullivan, Brian; Lester, Dean

    2012-01-01

    In response to the Columbia Accident Investigation Board report, a plug repair kit has been developed to enable astronauts to repair the space shuttle's wing leading edge (WLE) during orbit. The plug repair kit consists of several 17.78- cm-diameter carbon/silicon carbide (C/SiC) cover plates of various curvatures that can be attached to the refractory carbon-carbon WLE panels using a TZM refractory metal attach mechanism. The attach mechanism is inserted through the damage in the WLE panel and, as it is tightened, the cover plate flexes to conform to the curvature of the WLE panel within 0.050 mm. An astronaut installs the repair during an extravehicular activity (EVA). After installing the plug repair, edge gaps are checked and the perimeter of the repair is sealed using a proprietary material, developed to fill cracks and small holes in the WLE.

  6. 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.

  7. 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.

  8. The effect of leading edge tubercles on dynamic stall

    NASA Astrophysics Data System (ADS)

    Hrynuk, John

    The effect of the leading edge tubercles of humpback whales has been heavily studied for their static benefits. These studies have shown that tubercles inhibit flow separation, limit spanwise flow, and extend the operating angle of a wing beyond the static stall point while maintaining lift, all while having a comparatively low negative impact on drag. The current study extends the prior work to investigating the effect of tubercles on dynamic stall, a fundamental flow phenomenon that occurs when wings undergo dynamic pitching motions. Flow fields around the wing models tested were studied using Laser Induced Fluorescence (LIF) and Molecular Tagging Velocimetry (MTV).Resulting velocity fields show that the dynamics of the formation and separation of the leading edge vortex were fundamentally different between the straight wing and the tubercled wing. Tracking of the Dynamic Stall Vortex (DSV) and Shear Layer Vortices (SLVs), which may have a significant impact on the overall flow behavior, was done along with calculations of vortex circulation. Proximity to the wing surface and total circulation were used to evaluate potential dynamic lift increases provided by the tubercles. The effects of pitch rate on the formation process and benefits of the tubercles were also studied and were generally consistent with prior dynamic stall studies. However, tubercles were shown to affect the SLV formation and the circulation differently at higher pitch rates.

  9. Leading Edge Vortex Detection Using On-Body Pressure Sensing

    NASA Astrophysics Data System (ADS)

    Dusek, Jeff; Dahl, Jason; Triantafyllou, Michael

    2010-11-01

    Ongoing experiments within the Center for Environmental Sensing and Modeling (CENSAM) have shown that the low pressure region characteristic of a vortex allows for their detection and tracking using pressure sensors alone. While early experiments were conducted with wall mounted pressure sensors and externally generated vortices, a new series of experiments has succeeded in detecting separated flow generated by the sensing body. A combined pressure sensing and particle image velocimetry (PIV) approach was used to detect the leading edge vortex shed from a hydrofoil accelerated at a fixed angle of attack. A NACA 0018 foil was instrumented with four pressure sensors at discrete locations along the foil in the chord-wise direction. When accelerated from rest, the traces from each of the four pressure sensors displayed a distinctive, transient drop, consistent with results observed in previous experiments. From the pressure sensor results, it was theorized that a leading edge vortex was being created, and subsequently shed and convected along the foil chord. Two-dimensional PIV techniques were used to image the flow near the foil surface, allowing the anticipated vortex formation and shedding to be verified.

  10. Method for a Leading Edge Slat on a Wing of an Aircraft

    NASA Technical Reports Server (NTRS)

    Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

    2016-01-01

    A method for managing a flight control surface system. A leading edge device is moved on a leading edge from an undeployed position to a deployed position. The leading edge device has an outer surface, an inner surface, and a deformable fairing attached to the leading edge device such that the deformable fairing covers at least a portion of the inner surface. The deformable fairing changes from a deformed shape to an original shape when the leading edge device is moved to the deployed position. The leading edge device is then moved from the deployed position to the undeployed position, wherein the deformable fairing changes from the original shape to the deformed shape.

  11. Comparison of Theoretical and Experimental Unsteady Aerodynamics of Linear Oscillating Cascade With Supersonic Leading-Edge Locus

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2004-01-01

    An experimental influence coefficient technique was used to obtain unsteady aerodynamic influence coefficients and, consequently, unsteady pressures for a cascade of symmetric airfoils oscillating in pitch about mid-chord. Stagger angles of 0 deg and 10 deg were investigated for a cascade with a gap-to-chord ratio of 0.417 operating at an axial Mach number of 1.9, resulting in a supersonic leading-edge locus. Reduced frequencies ranged from 0.056 to 0.2. The influence coefficients obtained determine the unsteady pressures for any interblade phase angle. The unsteady pressures were compared with those predicted by several algorithms for interblade phase angles of 0 deg and 180 deg.

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

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.; Byrd, James E.; Mcgrath, Brian E.; Wesselmann, Gary F.

    1989-01-01

    An assessment of the influence of wing geometry on wing leading-edge vortex flows at supersonic speeds is discussed as well as the applicability of various aerodynamic codes for predicting these results. A series of delta-wing wind-tunnel models were tested in the NASA Langley Research Center Unitary Plan Wind Tunnel over a Mach number range from 1.6 to 4.6. The data show that wing airfoil has a significant impact on the localized loading on the wing. The experimental data for the flat wings were compared with results from full-potential, Euler, and Parabolized Navier-Stokes (PNS) computer codes. The theoretical evaluation showed that the full-potential analysis predicted accurate results for the attached-flow (alpha = 0 deg) conditions and that the Euler and PNS analyses made reasonable predictions for both attached and separated flow conditions.

  13. 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.

  14. Control of leading-edge vortices on a delta wing

    NASA Technical Reports Server (NTRS)

    Magness, C.; Robinson, O.; Rockwell, D.

    1992-01-01

    The unsteady flow structure of leading-edge vortices on a delta wing has been investigated using new types of experimental techniques, in order to provide insight into the consequences of various forms of active control. These investigations involve global control of the entire wing and local control applied at crucial locations on or adjacent to the wing. Transient control having long and short time-scales, relative to the convective time-scale C/U(sub infinity), allows substantial modification of the unsteady and time-mean flow structure. Global control at long time-scale involves pitching the wing at rates an order of magnitude lower than the convective time-scale C/U(sub infinity), but at large amplitudes. The functional form of the pitching maneuver exerts a predominant influence on the trajectory of the feeding sheet, the instantaneous vorticity distribution, and the instantaneous location of vortex breakdown. Global control at short time-scales of the order of the inherent frequency of the shear layer separating from the leading-edge and the natural frequency of vortex breakdown shows that 'resonant' response of the excited shear layer-vortex breakdown system is attainable. The spectral content of the induced disturbance is preserved not only across the entire core of the vortex, but also along the axis of the vortex into the region of vortex breakdown. This unsteady modification results in time-mean alteration of the axial and swirl velocity fields and the location of vortex breakdown. Localized control at long and short time-scales involves application of various transient forms of suction and blowing using small probes upstream and downstream of the location of vortex breakdown, as well as distributed suction and blowing along the leading-edge of the wing applied in a direction tangential to the feeding sheet. These local control techniques can result in substantial alteration of the location of vortex breakdown; in some cases, it is possible to

  15. Particle rebound characteristics of turbomachinery cascade leading edge geometry

    NASA Astrophysics Data System (ADS)

    Siravuri, Sastri

    The objective of this research work is to investigate and understand the complex phenomena associated with the mechanism of particle impacts on turbomachinery cascade leading edge geometry. At present, there is a need for experimental work in basic and applied research to find out the parameters that are relevant to particle rebound characteristics on turbomachinery blades. In the present work, experiments were conducted with air velocity at 15 m/s (˜50 ft/sec) and at 30 m/s (˜100 ft/sec) using high-speed photography and Laser Doppler Velocimetry (LDV). Silica sand particles of 1000--1500 micron size were used for this study. In the present investigation, particle rebound data was obtained for cylindrical targets with radius of curvature representative of leading edge geometry (cylinder diameter = 4.5mm & 6.5 mm) using LDV. The numerical simulations, which are based on non-linear dynamic analysis, were also performed using the finite element code DYNA3-D. Several different material models viz elastic-elastic, elastic-plastic, elastic-plastic with friction & isotropic-elastic-plastic with dynamic friction and particle rotation were used in the DYNA3-D numerical analysis. The computational results include a time history of the displacement, stress and strain profiles through the particle collision. Numerical results are presented for the rebound conditions of spherical silica sand particle for different pre-collision velocities. The computed particle restitution coefficients, after they reach steady rebound conditions, are compared with experimental results obtained from LDV. A probabilistic model was developed to incorporate the uncertainties in the impact velocity in the numerical model. Histograms and Cumulative Distribution Functions (CDFs) for impact velocity were obtained from experimental LDV data. Ten randomly selected probabilities for each impact angle were used to calculate the impact velocity from cumulative distribution function. This randomly selected

  16. Multiple shock-shock interference on a cylindrical leading edge

    NASA Technical Reports Server (NTRS)

    Wieting, Allan R.

    1991-01-01

    The details of an experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet are presented. This Mach 8 study has provided the first detailed pressure and heat transfer rate distributions on a cylinder resulting from a two-dimensional shockwave interference pattern created by two incident oblique shock waves intersecting the cylinder bow shock wave. The peak heat transfer rate was 38 times the undisturbed flow stagnation point level and occurred when the two oblique shock waves coalesced prior to intersecting the cylinder bow shock wave. Development of pressure deflection diagrams identified a new interference pattern consisting of concomitant supersonic jets separated from each other by a shear layer and submerged in the subsonic region between the bow shock wave and body.

  17. Managed aquifer recharge: rediscovering nature as a leading edge technology.

    PubMed

    Dillon, P; Toze, S; Page, D; Vanderzalm, J; Bekele, E; Sidhu, J; Rinck-Pfeiffer, S

    2010-01-01

    Use of Managed Aquifer Recharge (MAR) has rapidly increased in Australia, USA, and Europe in recent years as an efficient means of recycling stormwater or treated sewage effluent for non-potable and indirect potable reuse in urban and rural areas. Yet aquifers have been relied on knowingly for water storage and unwittingly for water treatment for millennia. Hence if 'leading edge' is defined as 'the foremost part of a trend; a vanguard', it would be misleading to claim managed aquifer recharge as a leading edge technology. However it has taken a significant investment in scientific research in recent years to demonstrate the effectiveness of aquifers as sustainable treatment systems to enable managed aquifer recharge to be recognised along side engineered treatment systems in water recycling. It is a 'cross-over' technology that is applicable to water and wastewater treatment and makes use of passive low energy processes to spectacularly reduce the energy requirements for water supply. It is robust within limits, has low cost, is suitable from village to city scale supplies, and offers as yet almost untapped opportunities for producing safe drinking water supplies where they do not yet exist. It will have an increasingly valued role in securing water supplies to sustain cities affected by climate change and population growth. However it is not a universal panacea and relies on the presence of suitable aquifers and sources of water together with effective governance to ensure human health and environment protection and water resources planning and management. This paper describes managed aquifer recharge, illustrates its use in Australia, outlining economics, guidelines and policies, and presents some of the knowledge about aquifer treatment processes that are revealing the latent value of aquifers as urban water infrastructure and provide a driver to improving our understanding of urban hydrogeology. PMID:21076220

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Rotational accelerations stabilize leading edge vortices on revolving fly wings.

    PubMed

    Lentink, David; Dickinson, Michael H

    2009-08-01

    The aerodynamic performance of hovering insects is largely explained by the presence of a stably attached leading edge vortex (LEV) on top of their wings. Although LEVs have been visualized on real, physically modeled, and simulated insects, the physical mechanisms responsible for their stability are poorly understood. To gain fundamental insight into LEV stability on flapping fly wings we expressed the Navier-Stokes equations in a rotating frame of reference attached to the wing's surface. Using these equations we show that LEV dynamics on flapping wings are governed by three terms: angular, centripetal and Coriolis acceleration. Our analysis for hovering conditions shows that angular acceleration is proportional to the inverse of dimensionless stroke amplitude, whereas Coriolis and centripetal acceleration are proportional to the inverse of the Rossby number. Using a dynamically scaled robot model of a flapping fruit fly wing to systematically vary these dimensionless numbers, we determined which of the three accelerations mediate LEV stability. Our force measurements and flow visualizations indicate that the LEV is stabilized by the ;quasi-steady' centripetal and Coriolis accelerations that are present at low Rossby number and result from the propeller-like sweep of the wing. In contrast, the unsteady angular acceleration that results from the back and forth motion of a flapping wing does not appear to play a role in the stable attachment of the LEV. Angular acceleration is, however, critical for LEV integrity as we found it can mediate LEV spiral bursting, a high Reynolds number effect. Our analysis and experiments further suggest that the mechanism responsible for LEV stability is not dependent on Reynolds number, at least over the range most relevant for insect flight (100

  5. Rotational accelerations stabilize leading edge vortices on revolving fly wings.

    PubMed

    Lentink, David; Dickinson, Michael H

    2009-08-01

    The aerodynamic performance of hovering insects is largely explained by the presence of a stably attached leading edge vortex (LEV) on top of their wings. Although LEVs have been visualized on real, physically modeled, and simulated insects, the physical mechanisms responsible for their stability are poorly understood. To gain fundamental insight into LEV stability on flapping fly wings we expressed the Navier-Stokes equations in a rotating frame of reference attached to the wing's surface. Using these equations we show that LEV dynamics on flapping wings are governed by three terms: angular, centripetal and Coriolis acceleration. Our analysis for hovering conditions shows that angular acceleration is proportional to the inverse of dimensionless stroke amplitude, whereas Coriolis and centripetal acceleration are proportional to the inverse of the Rossby number. Using a dynamically scaled robot model of a flapping fruit fly wing to systematically vary these dimensionless numbers, we determined which of the three accelerations mediate LEV stability. Our force measurements and flow visualizations indicate that the LEV is stabilized by the ;quasi-steady' centripetal and Coriolis accelerations that are present at low Rossby number and result from the propeller-like sweep of the wing. In contrast, the unsteady angular acceleration that results from the back and forth motion of a flapping wing does not appear to play a role in the stable attachment of the LEV. Angular acceleration is, however, critical for LEV integrity as we found it can mediate LEV spiral bursting, a high Reynolds number effect. Our analysis and experiments further suggest that the mechanism responsible for LEV stability is not dependent on Reynolds number, at least over the range most relevant for insect flight (100

  6. 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.

  7. 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.

  8. Effects of leading-edge tubercles on wing flutter speeds.

    PubMed

    Ng, B F; New, T H; Palacios, R

    2016-06-01

    The dynamic aeroelastic effects on wings modified with bio-inspired leading-edge (LE) tubercles are examined in this study. We adopt a state-space aeroelastic model via the coupling of unsteady vortex-lattice method and a composite beam to evaluate stability margins as a result of LE tubercles on a generic wing. The unsteady aerodynamics and spanwise mass variations due to LE tubercles have counteracting effects on stability margins with the former having dominant influence. When coupled, flutter speed is observed to be 5% higher, and this is accompanied by close to 6% decrease in reduced frequencies as an indication of lower structural stiffness requirements for wings with LE tubercles. Both tubercle amplitude and wavelength have similar influences over the change in flutter speeds, and such modifications to the LE would have minimal effect on stability margins when concentrated inboard of the wing. Lastly, when used in sweptback wings, LE tubercles are observed to have smaller impacts on stability margins as the sweep angle is increased. PMID:27070824

  9. Leading edge vortex in a slow-flying passerine

    PubMed Central

    Muijres, Florian T.; Johansson, L. Christoffer; Hedenström, Anders

    2012-01-01

    Most hovering animals, such as insects and hummingbirds, enhance lift by producing leading edge vortices (LEVs) and by using both the downstroke and upstroke for lift production. By contrast, most hovering passerine birds primarily use the downstroke to generate lift. To compensate for the nearly inactive upstroke, weight support during the downstroke needs to be relatively higher in passerines when compared with, e.g. hummingbirds. Here we show, by capturing the airflow around the wing of a freely flying pied flycatcher, that passerines may use LEVs during the downstroke to increase lift. The LEV contributes up to 49 per cent to weight support, which is three times higher than in hummingbirds, suggesting that avian hoverers compensate for the nearly inactive upstroke by generating stronger LEVs. Contrary to other animals, the LEV strength in the flycatcher is lowest near the wing tip, instead of highest. This is correlated with a spanwise reduction of the wing's angle-of-attack, partly owing to upward bending of primary feathers. We suggest that this helps to delay bursting and shedding of the particularly strong LEV in passerines. PMID:22417792

  10. Mechanisms of leading edge protrusion in interstitial migration

    PubMed Central

    Wilson, Kerry; Lewalle, Alexandre; Fritzsche, Marco; Thorogate, Richard; Duke, Tom; Charras, Guillaume

    2013-01-01

    While the molecular and biophysical mechanisms underlying cell protrusion on two-dimensional substrates are well understood, our knowledge of the actin structures driving protrusion in three-dimensional environments is poor, despite relevance to inflammation, development and cancer. Here we report that, during chemotactic migration through microchannels with 5 μm × 5 μm cross-sections, HL60 neutrophil-like cells assemble an actin-rich slab filling the whole channel cross-section at their front. This leading edge comprises two distinct F-actin networks: an adherent network that polymerizes perpendicular to cell-wall interfaces and a ‘free’ network that grows from the free membrane at the cell front. Each network is polymerized by a distinct nucleator and, due to their geometrical arrangement, the networks interact mechanically. On the basis of our experimental data, we propose that, during interstitial migration, medial growth of the adherent network compresses the free network preventing its retrograde movement and enabling new polymerization to be converted into forward protrusion. PMID:24305616

  11. Unsteady flow phenomena associated with leading-edge vortices

    NASA Astrophysics Data System (ADS)

    Breitsamter, C.

    2008-01-01

    This paper presents selected results from extensive experimental investigations on turbulent flow fields and unsteady surface pressures caused by leading-edge vortices, in particular, for vortex breakdown flow. Such turbulent flows may cause severe dynamic aeroelastic problems like wing and/or fin buffeting on fighter-type aircraft. The wind tunnel models used include a generic delta wing as well as a detailed aircraft configuration of canard-delta wing type. The turbulent flow structures are analyzed by root-mean-square and spectral distributions of velocity and pressure fluctuations. Downstream of bursting local maxima of velocity fluctuations occur in a limited radial range around the vortex center. The corresponding spectra exhibit significant peaks indicating that turbulent kinetic energy is channeled into a narrow band. These quasi-periodic velocity oscillations arise from a helical mode instability of the breakdown flow. Due to vortex bursting there is a characteristic increase in surface pressure fluctuations with increasing angle of attack, especially when the burst location moves closer to the apex. The pressure fluctuations also show dominant frequencies corresponding to those of the velocity fluctuations. Using the measured flow field data, scaling parameters are derived for design purposes. It is shown that a frequency parameter based on the local semi-span and the sinus of angle of attack can be used to estimate the frequencies of dynamic loads evoked by vortex bursting.

  12. Mask rule check for inspection of leading-edge photomask

    NASA Astrophysics Data System (ADS)

    Sakata, Wakahiko; Yamasaki, Kiyoshi; Narukawa, Shogo; Hayashi, Naoya

    2005-11-01

    Leading-edge photomask, to which optical proximity correction (OPC) and dummy pattern are applied, almost always has complex patterns. Complex patterns such as "Narrow Space", "Thin Pattern", "Dummy Pattern", "Closely Face-to-Face Heads" of Posi Serifs, "Narrow Waisted Pattern" formed by a Nega Serif, "Jogs", etc. are a factor to complicate photomask manufacturing. Some the problems caused by complex patterns are increase in EB writing time, and decrease in performance of etching and cleaning process caused by Cr peeling and, above all, increase in the inspection time. Patterns whose complexity is beyond the resolution limit of inspection tool are detected as false defects. Therefore, it will greatly take time for the data investigation and re-inspection, etc. for assurance, and this causes congestion of half-finished products. To improve the process efficiency, it is necessary to locate false defects, so that the Do-Not-Inspection-Area(DNIR) or replaced with simpler patterns. In order to locate false defects, it is proposed to apply Mask Rule Check (MRC) to mask data for EB-writing.

  13. 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.

  14. 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.

  15. Airfoil structure

    DOEpatents

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

    1998-01-13

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

  16. 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.

  17. Hypersonic aerospace vehicle leading edge cooling using heat pipe, transpiration and film cooling techniques

    NASA Astrophysics Data System (ADS)

    Modlin, James Michael

    An investigation was conducted to study the feasibility of cooling hypersonic vehicle leading edge structures exposed to severe aerodynamic surface heat fluxes using a combination of liquid metal heat pipes and surface mass transfer cooling techniques. A generalized, transient, finite difference based hypersonic leading edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading edge section. The hypersonic leading edge cooling model was developed using an existing, experimentally verified heat pipe model. Two applications of the hypersonic leading edge cooling model were examined. An assumed aerospace plane-type wing leading edge section exposed to a severe laminar, hypersonic aerodynamic surface heat flux was studied. A second application of the hypersonic leading edge cooling model was conducted on an assumed one-quarter inch nose diameter SCRAMJET engine inlet leading edge section exposed to both a transient laminar, hypersonic aerodynamic surface heat flux and a type 4 shock interference surface heat flux. The investigation led to the conclusion that cooling leading edge structures exposed to severe hypersonic flight environments using a combination of liquid metal heat pipe, surface transpiration, and film cooling methods appeared feasible.

  18. 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.

  19. The Columbia River--on the Leading Edge

    NASA Astrophysics Data System (ADS)

    O'Connor, J. E.

    2005-05-01

    On the leading edge of the North American plate, the Columbia River is the largest of the world's 40 or so rivers with drainage areas greater than 500,000 square kilometers to drain toward a convergent plate boundary. This unique setting results in a unique continental river basin; marked by episodic and cataclysmic geologic disturbance, but also famously fecund with perhaps 10 to 16 million salmon historically spawning in its waters each year. Now transformed by dams, transportation infrastructure, dikes and diversions, the Columbia River presents an expensive conundrum for management of its many values. Inclusion of river ecology and geomorphology in discussions of river management is generally limited to observations of the last 200 years-a time period of little natural disturbance and low sediment transport. However, consideration of longer timescales provides additional perspective of historical ecologic and geomorphic conditions. Only 230 km from its mouth, the Columbia River bisects the volcanic arc of the Cascade Range, forming the Columbia River Gorge. Cenozoic lava flows have blocked the river, forcing diversions and new canyon cutting. Holocene eruptions of Mount Mazama (Crater Lake), Mount Hood, Mount St. Helens, and Mount Rainier have shed immense quantities of sediment into the lower Columbia River, forming a large percentage of the Holocene sediment transported through the lower river. Quaternary landslides, perhaps triggered by great earthquakes, have descended from the 1000-m-high gorge walls, also blocking and diverting the river, one as recently as 550 years ago. These geologic disturbances, mostly outside the realm of historical observation and operating at timescales of 100s to 1000s of years in the gorge and elsewhere, have clearly affected basin geomorphology, riverine ecology, and past and present cultural utilization of river resources. The historic productivity of the river, however, hints at extraordinary resilience (and perhaps

  20. Morphological Variations of Leading-Edge Serrations in Owls (Strigiformes)

    PubMed Central

    Weger, Matthias; Wagner, Hermann

    2016-01-01

    Background Owls have developed serrations, comb-like structures, along the leading edge of their wings. Serrations were investigated from a morphological and a mechanical point of view, but were not yet quantitatively compared for different species. Such a comparative investigation of serrations from species of different sizes and activity patterns may provide new information about the function of the serrations. Results Serrations on complete wings and on tenth primary remiges of seven owl species were investigated. Small, middle-sized, and large owl species were investigated as well as species being more active during the day and owls being more active during the night. Serrations occurred at the outer parts of the wings, predominantly at tenth primary remiges, but also on further wing feathers in most species. Serration tips were oriented away from the feather rachis so that they faced into the air stream during flight. The serrations of nocturnal owl species were higher developed as demonstrated by a larger inclination angle (the angle between the base of the barb and the rachis), a larger tip displacement angle (the angle between the tip of the serration and the base of the serration) and a longer length. Putting the measured data into a clustering algorithm yielded dendrograms that suggested a strong influence of activity pattern, but only a weak influence of size on the development of the serrations. Conclusions Serrations are supposed to be involved in noise reduction during flight and also depend on the aerodynamic properties that in turn depend on body size. Since especially nocturnal owls have to rely on hearing during prey capture, the more pronounced serrations of nocturnal species lend further support to the notion that serrations have an important function in noise reduction. The differences in shape of the serrations investigated indicate that a silent flight requires well-developed serrations. PMID:26934104

  1. 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

  2. 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.

  3. 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.

  4. Flatback airfoil wind tunnel experiment.

    SciTech Connect

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

    2008-04-01

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

  5. Method and Apparatus for a Leading Edge Slat on a Wing of an Aircraft

    NASA Technical Reports Server (NTRS)

    Pitt, Dale M. (Inventor); Eckstein, Nicholas Stephen (Inventor)

    2013-01-01

    A method and apparatus for managing a flight control surface system. A leading edge device is moved on a leading edge from an undeployed position to a deployed position. The leading edge device has an outer surface, an inner surface, and a deformable fairing attached to the leading edge device such that the deformable fairing covers at least a portion of the inner surface. The deformable fairing changes from a deformed shape to an original shape when the leading edge device is moved to the deployed position. The leading edge device is then moved from the deployed position to the undeployed position, wherein the deformable fairing changes from the original shape to the deformed shape.

  6. Control of Flow Separation Using Adaptive Airfoils

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

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

  7. Noise Radiation From a Leading-Edge Slat

    NASA Technical Reports Server (NTRS)

    Lockhard, David P.; Choudhari, Meelan M.

    2009-01-01

    This paper extends our previous computations of unsteady flow within the slat cove region of a multi-element high-lift airfoil configuration, which showed that both statistical and structural aspects of the experimentally observed unsteady flow behavior can be captured via 3D simulations over a computational domain of narrow spanwise extent. Although such narrow domain simulation can account for the spanwise decorrelation of the slat cove fluctuations, the resulting database cannot be applied towards acoustic predictions of the slat without invoking additional approximations to synthesize the fluctuation field over the rest of the span. This deficiency is partially alleviated in the present work by increasing the spanwise extent of the computational domain from 37.3% of the slat chord to nearly 226% (i.e., 15% of the model span). The simulation database is used to verify consistency with previous computational results and, then, to develop predictions of the far-field noise radiation in conjunction with a frequency-domain Ffowcs-Williams Hawkings solver.

  8. Numerical and experimental study on the ability of dynamic roughness to alter the development of a leading edge vortex

    NASA Astrophysics Data System (ADS)

    Griffin, Christopher D.

    Dynamic stall is an unsteady aerodynamic phenomenon garnering much research interest because it occurs in a variety of applications. For example, dynamic stall is known to occur on helicopter rotor blades, wind turbines, high maneuvering military aircraft, and flapping wings. Dynamic stall occurs when an aerodynamic lifting device, such as an airfoil, wing, or turbomachine blade, undergoes a rapid pitching motion. It also occurs on lifting devices that are impulsively started at high angles of attack. Dynamic stall can "delay" aerodynamic stall to angles of attack that are significantly beyond the static stall angle of attack. During dynamic stall a large leading edge vortex (LEV) is formed, which creates greater fluid acceleration over the wing or airfoil, thus sustaining lift. As this vortex is shed downstream stall eventually occurs and there is an abrupt increase in drag and a large shift in pitching moment. Research has been performed to better understand the mechanisms occurring during dynamic stall in an effort to find ways to best take advantage of the increased lift associated with dynamic stall, but avoid the downfalls that occur once stall is initiated. Few attempts have been made to alter the LEV, and these attempts have used methods associated with laminar boundary layer separation control. Although these methods have shown promise, they suffer from the drawback that they exhaust more energy than is gained by flow control, while also only being effective at certain flight regimes. The research described herein documents the first study on the ability of dynamic roughness to alter the LEV encountered on a rapidly pitching airfoil. Both numerical and experimental studies were performed, including two-dimensional and three-dimensional computational fluid dynamics (CFD) simulations as well as stereo and planar particle image velocimetry (PIV) experiments. Evidence for the ability of small scale dynamic roughness to alter the development of the LEV was

  9. A numerical model for the platelet heat-pipe-cooled leading edge of hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Liu, Hongpeng; Liu, Weiqiang

    2016-01-01

    A new design, the platelet heat-pipe-cooled leading edge, is discussed for the thermal management to prevent damage to hypersonic vehicle leading edge component. For calculating the steady state behavior of platelet heat-pipe-cooled leading edge, a numerical model based on the principles of evaporation, convection, and condensation of a working fluid is presented. And then its effectiveness is validated by comparing the wall and vapor temperature against experimental data for a conventional heat pipe. Further investigations indicate that alloy IN718, with sodium as the working fluid is a feasible combination for Mach 8 flight with a 15 mm leading edge radius.

  10. 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.

  11. Combined overlay, focus and CD metrology for leading edge lithography

    NASA Astrophysics Data System (ADS)

    Ebert, Martin; Cramer, Hugo; Tel, Wim; Kubis, Michael; Megens, Henry

    2011-04-01

    As leading edge lithography moves to 22-nm design rules, low k1 technologies like double patterning are the new resolution enablers, and system control and setup are the new drivers to meet remarkably tight process requirements. The way of thinking and executing setup and control of lithography scanners is changing in four ways. First, unusually tight process tolerances call for very dense sampling [1], which in effect means measurements at high throughput combined with high order modeling and corrections to compensate for wafer spatial fingerprint. Second, complex interactions between scanner and process no longer allow separation of error sources through traditional metrology approaches, which are based on using one set of metrology tools and methods for setup and another for scanner performance control. Moreover, setup and control of overlay is done independently from CD uniformity, which in effect leads to independent and conflicting adjustments for the scanner. Third, traditional CD setup and control is based on the focus and dose calculated from their CD response and not from measurement of their effect on pattern profile, which allows a clean and orthogonal de-convolution of focus and dose variations across the wafer. Fourth, scanner setup and control has to take into consideration the final goal of lithography, which is the accurate printing of a complex pattern describing a real device layout. To this end we introduce a new setup and control metrology step: measuring-to-match scanner 1D and 2D proximity. In this paper we will describe the strategy for setup and control of overlay, focus, CD and proximity based on the YieldStarTM metrology tool and present the resulting performance. YieldStar-200 is a new, high throughput metrology tool based on a high numerical aperture scatterometer concept. The tool can be used stand-alone as well as integrated in a processing track. It is suitable for determining process offsets in X,Y and Z directions through Overlay

  12. 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.

  13. 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.

  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 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.

  16. 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.

  17. 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

  18. Preparation and Support of a Tap Test on the Leading Edge Surfaces of the Space Shuttle

    NASA Technical Reports Server (NTRS)

    Bohr, Jerry

    2009-01-01

    This slide presentation reports on a Tap test for the leading edge surfaces of the Space Shuttle. A description of the Wing Leading Edge Impact Detection System (WLEIDS) flight system is given, and the rationale and approach for improving the WLEIDS system. The three phases of the strategy of the test project amd the results of the tests are reviewed.

  19. Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 4. Boston Arts Academy

    ERIC Educational Resources Information Center

    Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

    2008-01-01

    This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high schools across the…

  20. Leading-edge deflection optimization for a highly swept arrow wing configuration

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Huffman, J. K.; Fenbert, J. W.

    1980-01-01

    Tests were also conducted to determine the sensitivity of the lateral stability derivative C sub l sub beta to geometric anhedral. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to the spanwise variation of upwash, the resulting optimized leading edge was a smooth, continuously warped surface. For the particular configuration studied, levels of leading edge suction on the order of 90 percent were achieved with the smooth, continuously warped leading edge contour. The results of tests conducted to determine the sensitivity of C sub l sub beta to geometric anhedral indicate values of delta C sub l sub beta/delta T which are in reasonable agreement with estimates provided by simple vortex lattice theories.

  1. 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.

  2. 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.

  3. Air-Cooled Turbine Blades with Tip Cap For Improved Leading-Edge Cooling

    NASA Technical Reports Server (NTRS)

    Calvert, Howard F.; Meyer, Andre J., Jr.; Morgan, William C.

    1959-01-01

    An investigation was conducted in a modified turbojet engine to determine the cooling characteristics of the semistrut corrugated air- cooled turbine blade and to compare and evaluate a leading-edge tip cap as a means for improving the leading-edge cooling characteristics of cooled turbine blades. Temperature data were obtained from uncapped air-cooled blades (blade A), cooled blades with the leading-edge tip area capped (blade B), and blades with slanted corrugations in addition to leading-edge tip caps (blade C). All data are for rated engine speed and turbine-inlet temperature (1660 F). A comparison of temperature data from blades A and B showed a leading-edge temperature reduction of about 130 F that could be attributed to the use of tip caps. Even better leading-edge cooling was obtained with blade C. Blade C also operated with the smallest chordwise temperature gradients of the blades tested, but tip-capped blade B operated with the lowest average chordwise temperature. According to a correlation of the experimental data, all three blade types 0 could operate satisfactorily with a turbine-inlet temperature of 2000 F and a coolant flow of 3 percent of engine mass flow or less, with an average chordwise temperature limit of 1400 F. Within the range of coolant flows investigated, however, only blade C could maintain a leading-edge temperature of 1400 F for a turbine-inlet temperature of 2000 F.

  4. Leading-edge vortex burst on a low-aspect-ratio rotating flat plate

    NASA Astrophysics Data System (ADS)

    Medina, Albert; Jones, Anya R.

    2016-08-01

    This study experimentally investigates the phenomenon of leading-edge-vortex burst on rotating flat plate wings. An aspect-ratio-2 wing was driven in pure rotation at a Reynolds number of Re=2500 . Of primary interest is the evolution of the leading-edge vortex along the wing span over a single-revolution wing stroke. Direct force measurements of the lift produced by the wing revealed a single global lift maximum relatively early in the wing stroke. Stereoscopic particle image velocimetry was applied to several chordwise planes to quantify the structure and strength of the leading-edge vortex and its effect on lift production. This analysis revealed opposite-sign vorticity entrainment into the core of the leading-edge vortex, originating from a layer of secondary vorticity along the wing surface. Coincident with the lift peak, there emerged both a concentration of opposite vorticity in the leading-edge-vortex core, as well as axial flow stagnation within the leading-edge-vortex core. Planar control volume analysis was performed at the midspan to quantify the contributions of vorticity transport mechanisms to the leading-edge-vortex circulation. The rate of circulation annihilation by opposite-signed vorticity entrainment was found to be minimal during peak lift production, where convection balanced the flux of vorticity resulting in stagnation and eventually reversal of axial flow. Finally, vortex burst was found to be correlated with swirl number, where bursting occurs at a swirl threshold of Sw<0.6 .

  5. The influence of leading-edge load alleviation on supersonic wing design

    NASA Technical Reports Server (NTRS)

    Darden, C. M.

    1984-01-01

    A theoretical and experimental program to assess the effect of leading-edge load constraints on wing design and performance was conducted. For a planform characterized by a highly swept leading edge on the inboard region, linear theory was used to design camber surfaces which produced minimum drag-due-to-lift at the design lift coefficient of 0.08 and a design Mach number of 2.4. In an effort to delay the formation of leading edge vortices which often occur on highly swept wings, two approaches were used in the design criteria to limit the loadings on the leading edge. One wing was constrained to have the normal Mach number less than one everywhere along the leading edge and the second wing was constrained to have a pressure coefficient of zero on the leading edge. Force tests were run on the two constrained wings, on a flat reference wing and on an optimized wing with no leading edge constraints. All wings had identical planforms and thicknesses and were tested over a range of Mach numbers from 1.8 to 2.8 and a range in angles of attack from -5 deg to 8 deg. A comparison of the experimental performance of these four models is shown. Correlations of these results with theoretical predictions and flow visualization photographs are also included.

  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. Aerodynamic Characteristics of a Large-Scale Unswept Wing-Body-Tail Configuration with Blowing Applied Over the Flap and Wind Leading Edge

    NASA Technical Reports Server (NTRS)

    McLemore, H. Clyde; Peterson, John B., Jr.

    1960-01-01

    An investigation has been conducted in the Langley full-scale tunnel to determine the effects of a blowing boundary-layer-control lift-augmentation system on the aerodynamic characteristics of a large-scale model of a fighter-type airplane. The wing was unswept at the 70-percent- chord station, had an aspect ratio of 2.86, a taper ratio of 0.40, and 4-percent-thick biconvex airfoil sections parallel to the plane of symmetry. The tests were conducted over a range of angles of attack from approximately -4 deg to 23 deg for a Reynolds number of approximately 5.2 x 10(exp 6) which corresponds to a Mach number of 0.08. Blowing rates were normally restricted to values just sufficient to control air-flow separation. The results of this investigation showed that wing leading-edge blowing in combination with large values of wing leading-edge-flap deflection was a very effective leading-edge flow-control device for wings having highly loaded trailing-edge flaps. With leading-edge blowing there was no hysteresis of the lift, drag, and pitching-moment characteristics upon recovery from stall. End plates were found to improve the lift and drag characteristics of the test configuration in the moderate angle-of-attack range, and blockage to one-quarter of the blowing-slot area was not detrimental to the aerodynamic characteristics. Blowing boundary-layer control resulted in a considerably reduced landing speed and reduced landing and take-off distances. The ailerons were very effective lateral-control devices when used with blowing flaps.

  8. 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.

  9. Effect of leading-edge load constraints on the design and performance of supersonic wings

    NASA Technical Reports Server (NTRS)

    Darden, C. M.

    1985-01-01

    A theoretical and experimental investigation was conducted to assess the effect of leading-edge load constraints on supersonic wing design and performance. In the effort to delay flow separation and the formation of leading-edge vortices, two constrained, linear-theory optimization approaches were used to limit the loadings on the leading edge of a variable-sweep planform design. Experimental force and moment tests were made on two constrained camber wings, a flat uncambered wing, and an optimum design with no constraints. Results indicate that vortex strength and separation regions were mildest on the severely and moderately constrained wings.

  10. Simulated airline service experience with laminar-flow control leading-edge systems

    NASA Technical Reports Server (NTRS)

    Maddalon, Dal V.; Fisher, David F.; Jennett, Lisa A.; Fischer, Michael C.

    1987-01-01

    The first JetStar leading edge flight test was made November 30, 1983. The JetStar was flown for more than 3 years. The titanium leading edge test articles today remain in virtually the same condition as they were in on that first flight. No degradation of laminar flow performance has occurred as a result of service. The JetStar simulated airline service flights have demonstrated that effective, practical leading edge systems are available for future commercial transports. Specific conclusions based on the results of the simulated airline service test program are summarized.

  11. 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.

  12. 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.

  13. Evolution and Control of the Leading Edge Vortex on an Unsteady Wing

    NASA Astrophysics Data System (ADS)

    Akkala, James; Buchholz, James

    2015-11-01

    The development of the leading-edge vortex is investigated on a periodically plunging plate within a uniform free stream. Vortex circulation is governed primarily by the strength of the leading edge shear layer, which provides the primary source of circulation, and a substantial opposite-sign contribution due to the pressure-gradient-driven diffusive flux of vorticity from the suction surface of the plate. The latter has been shown to produce a substantial reduction in leading-edge vortex strength, and leads to the development of a secondary vortex whose evolution influences the interaction between the leading edge vortex and the surface, and thus alters the surface pressure gradients. Suction is applied in the vicinity of the secondary vortex in an attempt to regulate the aerodynamic loads in the presence of the leading-edge vortex. The effect on vorticity transport, leading-edge vortex dynamics, and the resulting aerodynamic loads is discussed. This work was supported, in part, by the Air Force Office of Scientific Research under Grant number FA9550-11-1-0019 and the National Science Foundation under EPSCoR grant EPS1101284.

  14. Reynolds Number Effects on Leading Edge Radius Variations of a Supersonic Transport at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Rivers, S. M. B.; Wahls, R. A.; Owens, L. R.

    2001-01-01

    A computational study focused on leading-edge radius effects and associated Reynolds number sensitivity for a High Speed Civil Transport configuration at transonic conditions was conducted as part of NASA's High Speed Research Program. The primary purposes were to assess the capabilities of computational fluid dynamics to predict Reynolds number effects for a range of leading-edge radius distributions on a second-generation supersonic transport configuration, and to evaluate the potential performance benefits of each at the transonic cruise condition. Five leading-edge radius distributions are described, and the potential performance benefit including the Reynolds number sensitivity for each is presented. Computational results for two leading-edge radius distributions are compared with experimental results acquired in the National Transonic Facility over a broad Reynolds number range.

  15. User's manual for interfacing a leading edge, vortex rollup program with two linear panel methods

    NASA Technical Reports Server (NTRS)

    Desilva, B. M. E.; Medan, R. T.

    1979-01-01

    Sufficient instructions are provided for interfacing the Mangler-Smith, leading edge vortex rollup program with a vortex lattice (POTFAN) method and an advanced higher order, singularity linear analysis for computing the vortex effects for simple canard wing combinations.

  16. Simulation of brush insert for leading-edge-passage convective heat transfer

    NASA Technical Reports Server (NTRS)

    Hendricks, R. C.; Braun, M. J.; Canacci, V.; Mullen, R. L.

    1991-01-01

    Current and proposed high speed aircraft have high leading edge heat transfer (to 160 MW/sq m, 100 Btu/sq in/sec) and surface temperatures to 1370 K (2000 F). Without cooling, these surfaces could not survive. In one proposal the coolant hydrogen is circulated to the leading edge through a passage and returned to be consumed by the propulsion system. Simulated flow studies and visualizations have shown flow separation within the passage with a stagnation locus that isolates a zone of recirculation at the most critical portion of the passage, namely the leading edge itself. A novel method is described for mitigating the flow separation and the isolated recirculation zones by using a brush insert in the flow passage near the leading edge zone, thus providing a significant increase in heat transfer.

  17. Asymmetric distribution of Echinoid defines the epidermal leading edge during Drosophila dorsal closure

    PubMed Central

    Laplante, Caroline

    2011-01-01

    During Drosophila melanogaster dorsal closure, lateral sheets of embryonic epidermis assemble an actomyosin cable at their leading edge and migrate dorsally over the amnioserosa, converging at the dorsal midline. We show that disappearance of the homophilic cell adhesion molecule Echinoid (Ed) from the amnioserosa just before dorsal closure eliminates homophilic interactions with the adjacent dorsal-most epidermal (DME) cells, which comprise the leading edge. The resulting planar polarized distribution of Ed in the DME cells is essential for the localized accumulation of actin regulators and for actomyosin cable formation at the leading edge and for the polarized localization of the scaffolding protein Bazooka/PAR-3. DME cells with uniform Ed fail to assemble a cable and protrude dorsally, suggesting that the cable restricts dorsal migration. The planar polarized distribution of Ed in the DME cells thus provides a spatial cue that polarizes the DME cell actin cytoskeleton, defining the epidermal leading edge and establishing its contractile properties. PMID:21263031

  18. Manipulation of upstream rotor leading edge vortex and its effects on counter rotating propeller noise

    NASA Technical Reports Server (NTRS)

    Squires, Becky

    1993-01-01

    The leading edge vortex of a counter rotating propeller (CRP) model was altered by using shrouds and by turning the upstream rotors to a forward sweep configuration. Performance, flow, and acoustic data were used to determine the effect of vortex impingement on the noise signature of the CRP system. Forward sweep was found to eliminate the leading edge vortex of the upstream blades. Removal of the vortex had little effect on the tone noise at the forward and rear blade passing frequencies (BPF's) but significantly altered both the sound pressure level and directivity of the interaction tone which occurs at the sum of the two BPF's. A separate manipulation of the leading edge vortex performed by installing shrouds of various inlet length on the CRP verified that diverting the vortex path increases the noise level of the interaction tone. An unexpected link has been established between the interaction tone and the leading edge vortex-blade interaction phenomenon.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Space shuttle orbiter leading-edge flight performance compared to design goals

    NASA Technical Reports Server (NTRS)

    Curry, D. M.; Johnson, D. W.; Kelly, R. E.

    1983-01-01

    Thermo-structural performance of the Space Shuttle orbiter Columbia's leading-edge structural subsystem for the first five (5) flights is compared with the design goals. Lessons learned from thse initial flights of the first reusable manned spacecraft are discussed in order to assess design maturity, deficiencies, and modifications required to rectify the design deficiencies. Flight data and post-flight inspections support the conclusion that the leading-edge structural subsystem hardware performance was outstanding for the initial five (5) flights.

  4. Design and fabrication of a high temperature leading edge heating array, phase 1

    NASA Technical Reports Server (NTRS)

    1972-01-01

    Progress during a Phase 1 program to design a high temperature heating array is reported for environmentally testing full-scale shuttle leading edges (30 inch span, 6 to 15 inch radius) at flight heating rates and pressures. Heat transfer analyses of the heating array, individual modules, and the shuttle leading edge were performed, which influenced the array design, and the design, fabrication, and testing of a prototype heater module.

  5. Reduction of wing rock amplitudes using leading-edge vortex manipulations

    NASA Technical Reports Server (NTRS)

    Walton, James; Katz, Joseph

    1992-01-01

    A mechanically operated leading edge flap system was used to perturb leading edge vortex position on a free-to-roll double-delta wing. The motion of the flaps was synchronized with the wing rolling oscillations and the effect of the phase shift between the oscillations of the wing and the flaps was investigated. Experimental results indicated that this simple approach was effective in reducing the amplitude of the unintended rolling motion and its implementation to actual airplane configurations is rather simple.

  6. 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.

  7. Analytical model and stability analysis of the leading edge spar of a passively morphing ornithopter wing.

    PubMed

    Wissa, Aimy; Calogero, Joseph; Wereley, Norman; Hubbard, James E; Frecker, Mary

    2015-12-01

    This paper presents the stability analysis of the leading edge spar of a flapping wing unmanned air vehicle with a compliant spine inserted in it. The compliant spine is a mechanism that was designed to be flexible during the upstroke and stiff during the downstroke. Inserting a variable stiffness mechanism into the leading edge spar affects its structural stability. The model for the spar-spine system was formulated in terms of the well-known Mathieu's equation, in which the compliant spine was modeled as a torsional spring with a sinusoidal stiffness function. Experimental data was used to validate the model and results show agreement within 11%. The structural stability of the leading edge spar-spine system was determined analytically and graphically using a phase plane plot and Strutt diagrams. Lastly, a torsional viscous damper was added to the leading edge spar-spine model to investigate the effect of damping on stability. Results show that for the un-damped case, the leading edge spar-spine response was stable and bounded; however, there were areas of instability that appear for a range of spine upstroke and downstroke stiffnesses. Results also show that there exist a damping ratio between 0.2 and 0.5, for which the leading edge spar-spine system was stable for all values of spine upstroke and downstroke stiffnesses. PMID:26502210

  8. Analytical model and stability analysis of the leading edge spar of a passively morphing ornithopter wing.

    PubMed

    Wissa, Aimy; Calogero, Joseph; Wereley, Norman; Hubbard, James E; Frecker, Mary

    2015-10-26

    This paper presents the stability analysis of the leading edge spar of a flapping wing unmanned air vehicle with a compliant spine inserted in it. The compliant spine is a mechanism that was designed to be flexible during the upstroke and stiff during the downstroke. Inserting a variable stiffness mechanism into the leading edge spar affects its structural stability. The model for the spar-spine system was formulated in terms of the well-known Mathieu's equation, in which the compliant spine was modeled as a torsional spring with a sinusoidal stiffness function. Experimental data was used to validate the model and results show agreement within 11%. The structural stability of the leading edge spar-spine system was determined analytically and graphically using a phase plane plot and Strutt diagrams. Lastly, a torsional viscous damper was added to the leading edge spar-spine model to investigate the effect of damping on stability. Results show that for the un-damped case, the leading edge spar-spine response was stable and bounded; however, there were areas of instability that appear for a range of spine upstroke and downstroke stiffnesses. Results also show that there exist a damping ratio between 0.2 and 0.5, for which the leading edge spar-spine system was stable for all values of spine upstroke and downstroke stiffnesses.

  9. Hypersonic Engine Leading Edge Experiments in a High Heat Flux, Supersonic Flow Environment

    NASA Technical Reports Server (NTRS)

    Gladden, Herbert J.; Melis, Matthew E.

    1994-01-01

    A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Three aerothermal load related concerns are the boundary layer transition from laminar to turbulent flow, articulating panel seals in high temperature environments, and strut (or cowl) leading edges with shock-on-shock interactions. A multidisciplinary approach is required to address these technical concerns. A hydrogen/oxygen rocket engine heat source has been developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to experimentally evaluate the heat transfer and structural response of the strut (or cowl) leading edge. A recent experimental program conducted in this facility is discussed and related to cooling technology capability. The specific objective of the experiment discussed is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Heat transfer analyses of a similar leading edge concept cooled with gaseous hydrogen is included to demonstrate the complexity of the problem resulting from plastic deformation of the structures. Macro-photographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight.

  10. An experimental investigation of leading-edge vortex augmentation by blowing

    NASA Technical Reports Server (NTRS)

    Bradley, R. G.; Wray, W. O.; Smith, C. W.

    1974-01-01

    A wind tunnel test was conducted to determine the effects of over-the-wing blowing as a means of augmenting the leading-edge vortex flow of several pointed-tip, sharp-edged planforms. Arrow, delta, and diamond wings with leading-edge sweeps of 30 and 45 degrees were mounted on a body-of-revolution fuselage and tested in a low-speed wind tunnel at a Mach number of 0.2. Nozzle location data, pitch data, and flow-visualization pictures were obtained for a range of blowing rates. Results show pronounced increases in vortex lift due to the blowing.

  11. 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.

  12. Shuttle Wing Leading Edge Root Cause NDE Team Findings and Implementation of Quantitative Flash Infrared Thermography

    NASA Technical Reports Server (NTRS)

    Burke, Eric R.

    2009-01-01

    Comparison metrics can be established to reliably and repeatedly establish the health of the joggle region of the Orbiter Wing Leading Edge reinforced carbon carbon (RCC) panels. Using these metrics can greatly reduced the man hours needed to perform, wing leading edge scanning for service induced damage. These time savings have allowed for more thorough inspections to be preformed in the necessary areas with out affecting orbiter flow schedule. Using specialized local inspections allows for a larger margin of safety by allowing for more complete characterizations of panel defects. The presence of the t-seal during thermographic inspection can have adverse masking affects on ability properly characterize defects that exist in the joggle region of the RCC panels. This masking affect dictates the final specialized inspection should be preformed with the t-seal removed. Removal of the t-seal and use of the higher magnification optics has lead to the most effective and repeatable inspection method for characterizing and tracking defects in the wing leading edge. Through this study some inadequacies in the main health monitoring system for the orbiter wing leading edge have been identified and corrected. The use of metrics and local specialized inspection have lead to a greatly increased reliability and repeatable inspection of the shuttle wing leading edge.

  13. Evaluation of cloud detection instruments and performance of laminar-flow leading-edge test articles during NASA Leading-Edge Flight-Test Program

    NASA Technical Reports Server (NTRS)

    Davis, Richard E.; Maddalon, Dal V.; Wagner, Richard D.; Fisher, David F.; Young, Ronald

    1989-01-01

    Summary evaluations of the performance of laminar-flow control (LFC) leading edge test articles on a NASA JetStar aircraft are presented. Statistics, presented for the test articles' performance in haze and cloud situations, as well as in clear air, show a significant effect of cloud particle concentrations on the extent of laminar flow. The cloud particle environment was monitored by two instruments, a cloud particle spectrometer (Knollenberg probe) and a charging patch. Both instruments are evaluated as diagnostic aids for avoiding laminar-flow detrimental particle concentrations in future LFC aircraft operations. The data base covers 19 flights in the simulated airline service phase of the NASA Leading-Edge Flight-Test (LEFT) Program.

  14. Preferential attachment of membrane glycoproteins to the cytoskeleton at the leading edge of lamella

    PubMed Central

    1991-01-01

    The active forward movement of cells is often associated with the rearward transport of particles over the surfaces of their lamellae. Unlike the rest of the lamella, we found that the leading edge (within 0.5 microns of the cell boundary) is specialized for rearward transport of membrane-bound particles, such as Con A-coated latex microspheres. Using a single-beam optical gradient trap (optical tweezers) to apply restraining forces to particles, we can capture, move and release particles at will. When first bound on the central lamellar surface, Con A-coated particles would diffuse randomly; when such bound particles were brought to the leading edge of the lamella with the optical tweezers, they were often transported rearward. As in our previous studies, particle transport occurred with a concurrent decrease in apparent diffusion coefficient, consistent with attachment to the cytoskeleton. For particles at the leading edge of the lamella, weak attachment to the cytoskeleton and transport occurred with a half- time of 3 s; equivalent particles elsewhere on the lamella showed no detectable attachment when monitored for several minutes. Particles held on the cell surface by the laser trap attached more strongly to the cytoskeleton with time. These particles could escape a trapping force of 0.7 X 10(-6) dyne after 18 +/- 14 (sd) s at the leading edge, and after 64 +/- 34 (SD) s elsewhere on the lamella. Fluorescent succinylated Con A staining showed no corresponding concentration of general glycoproteins at the leading edge, but cytochalasin D-resistant filamentous actin was found at the leading edge. Our results have implications for cell motility: if the forces used for rearward particle transport were applied to a rigid substratum, cells would move forward. Such a mechanism would be most efficient if the leading edge of the cell contained preferential sites for attachment and transport. PMID:1874785

  15. 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

  16. Studies on wake-affected heat transfer around the circular leading edge of blunt body

    SciTech Connect

    Funazaki, K.

    1996-07-01

    Detailed measurements are performed about time-averaged heat transfer distributions around the leading edge of a blunt body, which is affected by incoming periodic wakes from the upstream moving bars. The blunt body is a test model of a front portion of a turbine blade in gas turbines and consists of a semicircular cylindrical leading edge and a flat plate afterbody. A wide range of the steady and unsteady flow conditions are adopted as for the Reynolds number based on the diameter of the leading edge and the bar-passing Strouhal number. The measured heat transfer distributions indicate that the wakes passing over the leading edge cause a significant increase in heat transfer before the separation and the higher Strouhal number results in higher heat transfer. From this experiment, a correlation for the heat transfer enhancement around the leading edge due to the periodic wakes is deduced as a function of the Stanton number and it is reviewed by comparison with the other experimental works.

  17. The three-dimensional leading-edge vortex of a 'hovering' model hawkmoth

    PubMed Central

    Berg, C. van den; Ellington, C.P.

    1997-01-01

    Recent flow visualisation experiments with the hawkmoth, Manduca sexta, revealed small but clear leading-edge vortex and a pronounced three-dimensional flow. Details of this flow pattern were studied with a scaled-up, robotic insect ('the flapper') that accurately mimicked the wing movements of a hovering hawkmoth. Smoke released from the leading edge of the flapper wing confirmed the existence of a small, strong and stable leading-edge vortex, increasing in size from wingbase to wingtip. Between 25 and 75 per cent of the wing length, its diameter increased approximately from 10 to 50 per cent of the wing chord. The leading-edge vortex had a strong axial flow veolocity, which stabilized it and reduced its diamater. The vortex separated from the wing at approximately 75 per cent of the wing length and thus fed vorticity into a large, tangled tip vortex. If the circulation of the leading-edge vortex were fully used for lift generation, it could support up to two-thirds of the hawkmoth's weight during the downstroke. The growth of this circulation with time and spanwise position clearly identify dynamic stall as the unsteady aerodynamic mechanism responsible for high lift production by hovering hawkmoths and possibly also by many other insect species.

  18. Acoustic Receptivity of Mach 4.5 Boundary Layer with Leading- Edge Bluntness

    NASA Technical Reports Server (NTRS)

    Malik, Mujeeb R.; Balakumar, Ponnampalam

    2007-01-01

    Boundary layer receptivity to two-dimensional slow and fast acoustic waves is investigated by solving Navier-Stokes equations for Mach 4.5 flow over a flat plate with a finite-thickness leading edge. Higher order spatial and temporal schemes are employed to obtain the solution whereby the flat-plate leading edge region is resolved by providing a sufficiently refined grid. The results show that the instability waves are generated in the leading edge region and that the boundary-layer is much more receptive to slow acoustic waves (by almost a factor of 20) as compared to the fast waves. Hence, this leading-edge receptivity mechanism is expected to be more relevant in the transition process for high Mach number flows where second mode instability is dominant. Computations are performed to investigate the effect of leading-edge thickness and it is found that bluntness tends to stabilize the boundary layer. Furthermore, the relative significance of fast acoustic waves is enhanced in the presence of bluntness. The effect of acoustic wave incidence angle is also studied and it is found that the receptivity of the boundary layer on the windward side (with respect to the acoustic forcing) decreases by more than a factor of 4 when the incidence angle is increased from 0 to 45 deg. However, the receptivity coefficient for the leeward side is found to vary relatively weakly with the incidence angle.

  19. The Three-Dimensional Leading-Edge Vortex of a "Hovering" Model Hawkmoth

    NASA Astrophysics Data System (ADS)

    van den Berg, Coen; Ellington, Charles P.

    1997-03-01

    Recent flow visualization experiments with the hawkmoth, Manduca sexta, revealed a small but clear leading-edge vortex and a pronounced three-dimensional flow. Details of this flow pattern were studied with a scaled-up, robotic insect ('the flapper') that accurately mimicked the wing movements of a hovering hawkmoth. Smoke released from the leading edge of the flapper wing confirmed the existence of a small, strong and stable leading-edge vortex, increasing in size from wingbase to wingtip. Between 25 and 75% of the wing length, its diameter increased approximately from 10 to 50% of the wing chord. The leading-edge vortex had a strong axial flow velocity, which stabilized it and reduced its diameter. The vortex separated from the wing at approximately 75% of the wing length and thus fed vorticity into a large, tangled tip vortex. If the circulation of the leading-edge vortex were fully used for lift generation, it could support up to two-thirds of the hawkmoth's weight during the downstroke. The growth of this circulation with time and spanwise position clearly identify dynamic stall as the unsteady aerodynamic mechanism responsible for high lift production by hovering hawkmoths and possibly also by many other insect species.

  20. Subsonic balance and pressure investigation of a 60 deg delta wing with leading edge devices

    NASA Technical Reports Server (NTRS)

    Tingas, S. A.; Rao, D. M.

    1982-01-01

    Low supersonic wave drag makes the thin highly swept delta wing the logical choice for use on aircraft designed for supersonic cruise. However, the high-lift maneuver capability of the aircraft is limited by severe induced-drag penalties attributed to loss of potential flow leading-edge suction. This drag increase may be alleviated through leading-edge flow control to recover lost aerodynamic thrust through either retention of attached leading-edge flow to higher angles of attack or exploitation of the increased suction potential of separation-induced vortex flow. A low-speed wind-tunnel investigation was undertaken to examine the high-lift devices such as fences, chordwise slots, pylon vortex generators, leading-edge vortex flaps, and sharp leading-edge extensions. The devices were tested individually and in combinations in an attempt to improve high-alpha drag performance with a minimum of low-alpha drag penalty. This report presents an analysis of the force, moment, and static pressure data obtained in angles of attack up to 23 deg, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter, respectively. The results indicate that all the devices produced drag and longitudinal/lateral stability improvements at high lift with, in most cases, minor drag penalties at low angles of attack.

  1. Transonic flight test of a laminar flow leading edge with surface excrescences

    NASA Technical Reports Server (NTRS)

    Zuniga, Fanny A.; Drake, Aaron; Kennelly, Robert A., Jr.; Koga, Dennis J.; Westphal, Russell V.

    1994-01-01

    A flight experiment, conducted at NASA Dryden Flight Research Center, investigated the effects of surface excrescences, specifically gaps and steps, on boundary-layer transition in the vicinity of a leading edge at transonic flight conditions. A natural laminar flow leading-edge model was designed for this experiment with a spanwise slot manufactured into the leading-edge model to simulate gaps and steps like those present at skin joints of small transonic aircraft wings. The leading-edge model was flown with the flight test fixture, a low-aspect ratio fin mounted beneath an F-104G aircraft. Test points were obtained over a unit Reynolds number range of 1.5 to 2.5 million/ft and a Mach number range of 0.5 to 0.8. Results for a smooth surface showed that laminar flow extended to approximately 12 in. behind the leading edge at Mach number 0.7 over a unit Reynolds number range of 1.5 to 2.0 million/ft. The maximum size of the gap-and-step configuration over which laminar flow was maintained consisted of two 0.06-in. gaps with a 0.02-in. step at a unit Reynolds number of 1.5 million/ft.

  2. 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.

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

    NASA Technical Reports Server (NTRS)

    Purser, Paul E.; Mckinney, Elizabeth G.

    1945-01-01

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

  4. 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

  5. Fabrication and Testing of a Leading-Edge-Shaped Heat Pipe

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Merrigan, Michael A.; Sena, J. Tom; Reid, Robert S.

    1998-01-01

    The development of a refractory-composite/heat-pipe-cooled leading edge has evolved from the design stage to the fabrication and testing of a full size, leading-edge-shaped heat pipe. The heat pipe had a 'D-shaped' cross section and was fabricated from arc cast Mo-4lRe. An artery was included in the wick. Several issues were resolved with the fabrication of the sharp leading edge radius heat pipe. The heat pipe was tested in a vacuum chamber at Los Alamos National Laboratory using induction heating and was started up from the frozen state several times. However, design temperatures and heat fluxes were not obtained due to premature failure of the heat pipe resulting from electrical discharge between the induction heating apparatus and the heat pipe. Though a testing anomaly caused premature failure of the heat pipe, successful startup and operation of the heat pipe was demonstrated.

  6. A flight test of laminar flow control leading-edge systems

    NASA Technical Reports Server (NTRS)

    Fischer, M. C.; Wright, A. S., Jr.; Wagner, R. D.

    1983-01-01

    NASA's program for development of a laminar flow technology base for application to commercial transports has made significant progress since its inception in 1976. Current efforts are focused on development of practical reliable systems for the leading-edge region where the most difficult problems in applying laminar flow exist. Practical solutions to these problems will remove many concerns about the ultimate practicality of laminar flow. To address these issues, two contractors performed studies, conducted development tests, and designed and fabricated fully functional leading-edge test articles for installation on the NASA JetStar aircraft. Systems evaluation and performance testing will be conducted to thoroughly evaluate all system capabilities and characteristics. A simulated airline service flight test program will be performed to obtain the operational sensitivity, maintenance, and reliability data needed to establish that practical solutions exist for the difficult leading-edge area of a future commercial transport employing laminar flow control.

  7. Actin filament turnover drives leading edge growth during myelin sheath formation in the central nervous system.

    PubMed

    Nawaz, Schanila; Sánchez, Paula; Schmitt, Sebastian; Snaidero, Nicolas; Mitkovski, Mišo; Velte, Caroline; Brückner, Bastian R; Alexopoulos, Ioannis; Czopka, Tim; Jung, Sang Y; Rhee, Jeong S; Janshoff, Andreas; Witke, Walter; Schaap, Iwan A T; Lyons, David A; Simons, Mikael

    2015-07-27

    During CNS development, oligodendrocytes wrap their plasma membrane around axons to generate multilamellar myelin sheaths. To drive growth at the leading edge of myelin at the interface with the axon, mechanical forces are necessary, but the underlying mechanisms are not known. Using an interdisciplinary approach that combines morphological, genetic, and biophysical analyses, we identified a key role for actin filament network turnover in myelin growth. At the onset of myelin biogenesis, F-actin is redistributed to the leading edge, where its polymerization-based forces push out non-adhesive and motile protrusions. F-actin disassembly converts protrusions into sheets by reducing surface tension and in turn inducing membrane spreading and adhesion. We identified the actin depolymerizing factor ADF/cofilin1, which mediates high F-actin turnover rates, as an essential factor in this process. We propose that F-actin turnover is the driving force in myelin wrapping by regulating repetitive cycles of leading edge protrusion and spreading.

  8. Closed Form Equations for the Preliminary Design of a Heat-Pipe-Cooled Leading Edge

    NASA Technical Reports Server (NTRS)

    Glass, David E.

    1998-01-01

    A set of closed form equations for the preliminary evaluation and design of a heat-pipe-cooled leading edge is presented. The set of equations can provide a leading-edge designer with a quick evaluation of the feasibility of using heat-pipe cooling. The heat pipes can be embedded in a metallic or composite structure. The maximum heat flux, total integrated heat load, and thermal properties of the structure and heat-pipe container are required input. The heat-pipe operating temperature, maximum surface temperature, heat-pipe length, and heat pipe-spacing can be estimated. Results using the design equations compared well with those from a 3-D finite element analysis for both a large and small radius leading edge.

  9. Test and Analysis Correlation of Form Impact onto Space Shuttle Wing Leading Edge RCC Panel 8

    NASA Technical Reports Server (NTRS)

    Fasanella, Edwin L.; Lyle, Karen H.; Gabrys, Jonathan; Melis, Matthew; Carney, Kelly

    2004-01-01

    Soon after the Columbia Accident Investigation Board (CAIB) began their study of the space shuttle Columbia accident, "physics-based" analyses using LS-DYNA were applied to characterize the expected damage to the Reinforced Carbon-Carbon (RCC) leading edge from high-speed foam impacts. Forensic evidence quickly led CAIB investigators to concentrate on the left wing leading edge RCC panels. This paper will concentrate on the test of the left-wing RCC panel 8 conducted at Southwest Research Institute (SwRI) and the correlation with an LS-DYNA analysis. The successful correlation of the LS-DYNA model has resulted in the use of LS-DYNA as a predictive tool for characterizing the threshold of damage for impacts of various debris such as foam, ice, and ablators onto the RCC leading edge for shuttle return-to-flight.

  10. An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

    NASA Technical Reports Server (NTRS)

    Ghaffari, F.; Chaturvedi, S. K.

    1984-01-01

    An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

  11. Summary of past experience in natural laminar flow and experimental program for resilient leading edge

    NASA Technical Reports Server (NTRS)

    Carmichael, B. H.

    1979-01-01

    The potential of natural laminar flow for significant drag reduction and improved efficiency for aircraft is assessed. Past experience with natural laminar flow as reported in published and unpublished data and personal observations of various researchers is summarized. Aspects discussed include surface contour, waviness, and smoothness requirements; noise and vibration effects on boundary layer transition, boundary layer stability criteria; flight experience with natural laminar flow and suction stabilized boundary layers; and propeller slipstream, rain, frost, ice and insect contamination effects on boundary layer transition. The resilient leading edge appears to be a very promising method to prevent leading edge insect contamination.

  12. Image analysis tools to quantify cell shape and protein dynamics near the leading edge.

    PubMed

    Ryan, Gillian L; Watanabe, Naoki; Vavylonis, Dimitrios

    2013-01-01

    We present a set of flexible image analysis tools to analyze dynamics of cell shape and protein concentrations near the leading edge of cells adhered to glass coverslips. Plugins for ImageJ streamline common analyses of microscopic images of cells, including the calculation of leading edge speeds, total and average intensities of fluorescent markers, and retrograde flow rate measurements of fluorescent single-molecule speckles. We also provide automated calculations of auto- and cross-correlation functions between velocity and intensity measurements. The application of the methods is illustrated on images of XTC cells.

  13. Method and equipment for induction surface hardening of the leading edges of turbine blades

    SciTech Connect

    Sorokina, T.M.; Dymchenko, V.V.

    1988-01-01

    Methodology and equipment for hardening the leading edges of blades for large nuclear reactor steam turbines was investigated using blades made of 15Kh11MF hardened and tempered steel. A machine was designed and built for hardening the blade leading edges with a vacuum-tube oscillator and 66,000 Hz frequency. The electrical parameters of the induction heating were recorded. Hardening of the actual blades made it possible to obtain a hardened case with a depth of 1-3 mm and up to 5 mm in the lower portion of the blade and increased erosion resistance.

  14. Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2002-01-01

    A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  15. Reynolds Number and Leading-Edge Bluntness Effects on a 65 Deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2002-01-01

    A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  16. Reynolds Number, Compressibility, and Leading-Edge Bluntness Effects on Delta-Wing Aerodynamics

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    2004-01-01

    An overview of Reynolds number, compressibility, and leading edge bluntness effects is presented for a 65 degree delta wing. The results of this study address both attached and vortex-flow aerodynamics and are based upon a unique data set obtained in the NASA-Langley National Transonic Facility (NTF) for i) Reynolds numbers ranging from conventional wind-tunnel to flight values, ii) Mach numbers ranging from subsonic to transonic speeds, and iii) leading-edge bluntness values that span practical slender wing applications. The data were obtained so as to isolate the subject effects and they present many challenges for Computational Fluid Dynamics (CFD) studies.

  17. Space Shuttle Orbiter nose cap and wing leading edge certification test program

    NASA Technical Reports Server (NTRS)

    Suppanz, M. J.; Grimaud, J. E.

    1980-01-01

    A reinforced carbon-carbon thermal protection system is used on the space shuttle orbiter vehicle's nose cap and wing leading edge regions where temperatures reach 1538 C (2800 F). To verify the analyses used to certify reinforced carbon-carbon for the first flight and operational missions, a multi-environment incremental test program was developed and implemented through the combined efforts of Rockwell International and the National Aeronautics and Space Administration. Three separate facilities at the Johnson Space Center were used to subject full-scale nose cap and wing leading edge test articles to simulated critical launch, on-orbit, and atmospheric entry environments.

  18. Experimental study of delta wing leading-edge devices for drag reduction at high lift

    NASA Technical Reports Server (NTRS)

    Johnson, T. D., Jr.; Rao, D. M.

    1982-01-01

    The drag reduction devices selected for evaluation were the fence, slot, pylon-type vortex generator, and sharp leading-edge extension. These devices were tested on a 60 degree flatplate delta (with blunt leading edges) in the Langley Research Center 7- by 10-foot high-speed tunnel at low speed and to angles of attack of 28 degrees. Balance and static pressure measurements were taken. The results indicate that all the devices had significant drag reduction capability and improved longitudinal stability while a slight loss of lift and increased cruise drag occurred.

  19. 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.

  20. 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.

  1. Composite airfoil assembly

    SciTech Connect

    Garcia-Crespo, Andres Jose

    2015-03-03

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

  2. Airfoils for wind turbine

    DOEpatents

    Tangler, James L.; Somers, Dan M.

    1996-01-01

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

  3. Airfoils for wind turbine

    DOEpatents

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

    1996-10-08

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

  4. The Leading Edge: A Career Development Workshop Series for Young Adults. Participant Workbook.

    ERIC Educational Resources Information Center

    Canadian Career Development Foundation, Ottawa (Ontario).

    This booklet is designed for participants in "The Leading Edge: A Career Development Workshop Series for Young Adults." It provides the 27 participant handouts for the six workshops in the series. The first in the series, "Setting the Stage: The Changing World of Work," is a workshop to clarify what is occurring in the world of work and apply that…

  5. The Leading Edge: Competencies for Community College Leadership in the New Millennium.

    ERIC Educational Resources Information Center

    Desjardins, Carolyn

    Published in collaboration with the National Institute for Leadership Development (NILD), this monograph reports on a study of leading-edge community college presidents. The authors cluster the 22 competencies identified through the study into four categories. Each category is presented as a chapter which identifies and provides examples of the…

  6. Project 2000-3 Leading Edge Enterprise: Insights into Employment and Training Practices. Working Paper.

    ERIC Educational Resources Information Center

    Long, Michael; Fischer, John

    Leading-edge firms (LEFs)--at the forefront of their industry in terms of growth or market share--may influence skill development through diffusion of technology, products, or practices and use of market power to set standards or change customer businesses. Study of LEFs can identify the type and mix of skills needed in the industry. LEFs are…

  7. A leading edge heating array and a flat surface heating array - operation, maintenance and repair manual

    NASA Technical Reports Server (NTRS)

    1975-01-01

    A general description of the leading edge/flat surface heating array is presented along with its components, assembly instructions, installation instructions, operation procedures, maintenance instructions, repair procedures, schematics, spare parts lists, engineering drawings of the array, and functional acceptance test log sheets. The proper replacement of components, correct torque values, step-by-step maintenance instructions, and pretest checkouts are described.

  8. Spanwise gradients in flow speed help stabilize leading-edge vortices on revolving wings.

    PubMed

    Jardin, T; David, L

    2014-07-01

    While a leading-edge vortex on an infinite translating wing is shed after a short distance of travel, its counterpart on a finite span revolving insect wing or maple seed membrane exhibits robust attachment. The latter explains the aerodynamic lift generated by such biological species. Here we analyze the mechanisms responsible for leading-edge vortex attachment. We compute the Navier-Stokes solution of the flow past a finite span wing (i) embedded in a uniform oncoming flow, (ii) embedded in a spanwise varying oncoming flow, and (iii) revolving about its root. We show that over flapping amplitudes typical of insect flight (ϕ = 120°), the spanwise gradient of the local wing speed may suffice in maintaining leading-edge vortex attachment. We correlate this result with the development of spanwise flow, driven by the spanwise gradient of pressure, and we evaluate the sensitivity of such a mechanism to the Reynolds number. It is noted, however, that leading-edge vortex attachment through the spanwise gradient of the local wing speed does not promote large lift, which ultimately arises from centrifugal and Coriolis effects. PMID:25122373

  9. Direct numerical simulations of leading-edge receptivity for freestream sound

    NASA Technical Reports Server (NTRS)

    Fuciarelli, David A.; Reed, Helen L.

    1994-01-01

    The objective of this paper is to review our efforts in spatial direct numerical simulations for modeling leading-edge receptivity to freestream sound and vorticity. These results begin to provide the link between the freestream and the initial boundary-layer response and can provide the upstream conditions for further simulations marching through the transition process toward turbulence.

  10. Effects of Fin Leading Edge Sweep on Shock-Shock Interaction at Mach 6

    NASA Technical Reports Server (NTRS)

    Berry, Scott A.; Nowak, Robert J.

    1996-01-01

    The effects of fin leading edge sweep on peak heating rates due to shock-shock interaction have been experimentally examined in the Langley 20-Inch Mach 6 Tunnel. The shock interaction was produced by the intersection of a planar incident shock (16.8 deg shock angle relative to the freestream, generated by a 9 deg wedge) with the bow shock formed around a O.5-inch diameter cylindrical leading edge fin. Heating distributions along the leading edge stagnation line have been obtained using densely spaced thin film resistive-type sensors. Schlieren images were obtained to illustrate the very complex shock-shock interactions. The fin leading edge sweep angle was varied from 15-degrees swept back to 45-degrees swept forward for a freestream unit Reynolds number of 2 x 10(exp 6)/ft. Two models were utilized during the study, one with 0.025-inch spacing between gage centers, and the other 0.015-inch spacing. Gage spatial resolution on the order of 0.015-in appeared to accurately capture the narrow spike in heating. Peak heating due to shock interaction was maximized when the fin was swept forward 15 deg and 25 deg, both promoting augmentations about 7 times the baseline value. The schlieren images for these cases revealed Type 4 and Type 3 interactions, respectively.

  11. The Leading Edge: A Career Development Workshop Series for Young Adults. Facilitator Guide.

    ERIC Educational Resources Information Center

    Canadian Career Development Foundation, Ottawa (Ontario).

    This booklet is designed to be used by facilitators of the Canadian Career Development Foundation's "The Leading Edge: A Career Development Workshop Series for Young Adults." The guide provides information, including objectives of the workshops and lists of required materials, needed in order to facilitate an introductory session as well as the…

  12. New American High Schools: Profiles of the Nation's Leading Edge Schools.

    ERIC Educational Resources Information Center

    Department of Education, Washington, DC.

    This booklet profiles "leading edge" schools committed to ensuring that all students meet challenging academic standards and are prepared for college and careers. In 1996, these 10 New American High Schools were chosen by the U.S. Department of Education for their innovation and commitment to academic excellence. As these award-winning,…

  13. Spanwise gradients in flow speed help stabilize leading-edge vortices on revolving wings

    NASA Astrophysics Data System (ADS)

    Jardin, T.; David, L.

    2014-07-01

    While a leading-edge vortex on an infinite translating wing is shed after a short distance of travel, its counterpart on a finite span revolving insect wing or maple seed membrane exhibits robust attachment. The latter explains the aerodynamic lift generated by such biological species. Here we analyze the mechanisms responsible for leading-edge vortex attachment. We compute the Navier-Stokes solution of the flow past a finite span wing (i) embedded in a uniform oncoming flow, (ii) embedded in a spanwise varying oncoming flow, and (iii) revolving about its root. We show that over flapping amplitudes typical of insect flight (ϕ =120∘), the spanwise gradient of the local wing speed may suffice in maintaining leading-edge vortex attachment. We correlate this result with the development of spanwise flow, driven by the spanwise gradient of pressure, and we evaluate the sensitivity of such a mechanism to the Reynolds number. It is noted, however, that leading-edge vortex attachment through the spanwise gradient of the local wing speed does not promote large lift, which ultimately arises from centrifugal and Coriolis effects.

  14. A role for actin arcs in the leading edge advance of migrating cells

    PubMed Central

    Burnette, Dylan T.; Manley, Suliana; Sengupta, Prabuddha; Sougrat, Rachid; Davidson, Michael W.; Kachar, Bechara; Lippincott-Schwartz, Jennifer

    2013-01-01

    The migration of epithelial cells requires coordination of two actin modules at the leading edge: one in the lamellipodium and one in the lamella. How the two modules connect mechanistically to regulate directed edge motion is not understood. Using a combination of live-cell imaging and photoactivation approaches, we demonstrate that the actin network of the lamellipodium evolves spatio-temporally into the lamella. This occurs during the retraction phase of edge motion when myosin II redistributes to the cell edge and condenses the lamellipodial-actin into an arc-like bundle (i.e., actin arc) parallel to the edge. The newly formed actin arc moves rearward and couples to focal adhesions as it enters the lamella. We propose net edge extension occurs by nascent focal adhesions advancing the site at which new actin arcs slow down and form the base of the next protrusion event. The actin arc thus serves as a structural element underlying the temporal and spatial connection between the lamellipodium and lamella to drive directed cell motion. PMID:21423177

  15. 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.

  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. 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.

  18. 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.

  19. 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.

  20. 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

  1. 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.

  2. 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.

  3. Experimental And Numerical Study Of CMC Leading Edges In Hypersonic Flows

    NASA Astrophysics Data System (ADS)

    Kuhn, Markus; Esser, Burkard; Gulhan, Ali; Dalenbring, Mats; Cavagna, Luca

    2011-05-01

    Future transportation concepts aim at high supersonic or hypersonic speeds, where the formerly sharp boundaries between aeronautic and aerospace applications become blurred. One of the major issues involved to high speed flight are extremely high aerothermal loads, which especially appear at the leading edges of the plane’s wings and at sharp edged air intake components of the propulsion system. As classical materials like metals or simple ceramics would thermally and structurally fail here, new materials have to be applied. In this context, lightweight ceramic matrix composites (CMC) seem to be prospective candidates as they are high-temperature resistant and offer low thermal expansion along with high specific strength at elevated temperature levels. A generic leading edge model with a ceramic wing assembly with a sweep back angle of 53° was designed, which allowed for easy leading edge sample integration of different CMC materials. The samples consisted of the materials C/C-SiC (non-oxide), OXIPOL and WHIPOX (both oxide) with a nose radius of 2 mm. In addition, a sharp edged C/C-SiC sample was prepared to investigate the nose radius influence. Overall, 13 thermocouples were installed inside the entire model to measure the temperature evolution at specific locations, whereby 5 thermocouples were placed inside the leading edge sample itself. In addition, non-intrusive techniques were applied for surface temperature measurements: An infrared camera was used to measure the surface temperature distribution and at specific spots, the surface temperature was also measured by pyrometers. Following, the model was investigated in DLR’s arc-heated facility L3K at a total enthalpy of 8.5 MJ/kg, Mach number of 7.8, different angles of attack and varying wing inclination angles. These experiments provide a sound basis for the simulation of aerothermally loaded CMC leading edge structures. Such fluid-structure coupled approaches have been performed by FOI, basing on a

  4. 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.

  5. A theoretical investigation of the aerodynamics of low-aspect-ratio wings with partial leading-edge separation

    NASA Technical Reports Server (NTRS)

    Mehrotra, S. C.; Lan, C. E.

    1978-01-01

    A numerical method is developed to predict distributed and total aerodynamic characteristics for low aspect-ratio wings with partial leading-edge separation. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the quasi-vortex-lattice method. The leading-edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at mid-points to satisfy the force free condition. The wake behind the trailing-edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading- and trailing-edges. Comparison of the predicted results with complete leading-edge separation has shown reasonably good agreement. For cases with partial leading-edge separation, the lift is found to be highly nonlinear with angle of attack.

  6. 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.

  7. 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.

  8. 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.

  9. Aerothermal Performance Constraints for Hypervelocity Small Radius Unswept Leading Edges and Nosetips

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul

    1997-01-01

    Small radius leading edges and nosetips were utilized to minimize wave drag in early hypervelocity vehicle concepts until further analysis demonstrated that extreme aerothermodynamic heating would cause severe ablation or blunting of the available thermal protection system materials. Recent studies indicate that ultrahigh temperature ceramic (UHTC) materials are shape stable at temperatures approaching 3033 K and will be available for use as sharp UHTC leading edge components in the near future. Aerothermal performance constraints for sharp components made from these materials are presented in this work to demonstrate the effects of convective blocking, surface catalycity, surface emissivity, and rarefied flow effects on steady state operation at altitudes from sea level to 90 km. These components are capable of steady state operation at velocities up to 7.9 km/s at attitudes near 90 km.

  10. High-order aberration control during exposure for leading-edge lithography projection optics

    NASA Astrophysics Data System (ADS)

    Ohmura, Yasuhiro; Tsuge, Yosuke; Hirayama, Toru; Ikezawa, Hironori; Inoue, Daisuke; Kitamura, Yasuhiro; Koizumi, Yukio; Hasegawa, Keisuke; Ishiyama, Satoshi; Nakashima, Toshiharu; Kikuchi, Takahisa; Onda, Minoru; Takase, Yohei; Nagahiro, Akimasa; Isago, Susumu; Kawahara, Hidetaka

    2016-03-01

    High throughput with high resolution imaging has been key to the development of leading-edge microlithography. However, management of thermal aberrations due to lens heating during exposure has become critical for simultaneous achievement of high throughput and high resolution. Thermal aberrations cause CD drift and overlay error, and these errors lead directly to edge placement errors (EPE). Management and control of high order thermal aberrations is a critical requirement. In this paper, we will show practical performance of the lens heating with dipole and other typical illumination conditions for finer patterning. We confirm that our new control system can reduce the high-order aberrations and enable critical-dimension uniformity CDU during the exposure.

  11. Cooling Strategies for Vane Leading Edges in a Syngas Environment Including Effects of Deposition and Turbulence

    SciTech Connect

    Ames, Forrest; Bons, Jeffrey

    2014-09-30

    The Department of Energy has goals to move land based gas turbine systems to alternate fuels including coal derived synthetic gas and hydrogen. Coal is the most abundant energy resource in the US and in the world and it is economically advantageous to develop power systems which can use coal. Integrated gasification combined cycles are (IGCC) expected to allow the clean use of coal derived fuels while improving the ability to capture and sequester carbon dioxide. These cycles will need to maintain or increase turbine entry temperatures to develop competitive efficiencies. The use of coal derived syngas introduces a range of potential contaminants into the hot section of the gas turbine including sulfur, iron, calcium, and various alkali metals. Depending on the effectiveness of the gas clean up processes, there exists significant likelihood that the remaining materials will become molten in the combustion process and potentially deposit on downstream turbine surfaces. Past evidence suggests that deposition will be a strong function of increasing temperature. Currently, even with the best gas cleanup processes a small level of particulate matter in the syngas is expected. Consequently, particulate deposition is expected to be an important consideration in the design of turbine components. The leading edge region of first stage vanes most often have higher deposition rates than other areas due to strong fluid acceleration and streamline curvature in the vicinity of the surface. This region remains one of the most difficult areas in a turbine nozzle to cool due to high inlet temperatures and only a small pressure ratio for cooling. The leading edge of a vane often has relatively high heat transfer coefficients and is often cooled using showerhead film cooling arrays. The throat of the first stage nozzle is another area where deposition potentially has a strongly adverse effect on turbine performance as this region meters the turbine inlet flow. Based on roughness

  12. Leading edge reflection patterns for cylindrical converging shock waves over convex obstacles

    NASA Astrophysics Data System (ADS)

    Vignati, F.; Guardone, A.

    2016-09-01

    The unsteady reflection of cylindrical converging shock waves over convex obstacles is investigated numerically. At the leading edge, numerical simulations show the occurrence of all types of regular and irregular reflections predicted by the pseudo-steady theory for planar shock-wave reflections over planar surfaces, although for different combinations of wedge angles and incident shock Mach number. The domain of occurrence of each reflection type and its evolution in time due to shock acceleration and to the non-planar geometry is determined and it is compared to the results of the pseudo-steady theory. The dependence of the reflection pattern on the (local) values of the wedge angle is in good agreement with the pseudo-steady theory. Less complex reflection patterns are instead observed at larger values of the leading edge shock Mach number at which the pseudo-steady theory predicts the occurrence of more complex reflection patterns.

  13. Test and Analysis of a Hyper-X Carbon-Carbon Leading Edge Chine

    NASA Technical Reports Server (NTRS)

    Smith, Russell W.; Sikora, Joseph G.; Lindell, Michael C.

    2005-01-01

    During parts production for the X43A Mach 10 hypersonic vehicle nondestructive evaluation (NDE) of a leading edge chine detected on imbedded delamination near the lower surface of the part. An ultimate proof test was conducted to verify the ultimate strength of this leading edge chine part. The ultimate proof test setup used a pressure bladder design to impose a uniform distributed pressure field over the bi-planar surface of the chine test article. A detailed description of the chine test article and experimental test setup is presented. Analysis results from a linear status model of the test article are also presented and discussed. Post-test inspection of the specimen revealed no visible failures or areas of delamination.

  14. A Reduced-Complexity Investigation of Blunt Leading-Edge Separation Motivated by UCAV Aerodynamics

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Boelens, Okko J.

    2015-01-01

    A reduced complexity investigation for blunt-leading-edge vortical separation has been undertaken. The overall approach is to design the fundamental work in such a way so that it relates to the aerodynamics of a more complex Uninhabited Combat Air Vehicle (UCAV) concept known as SACCON. Some of the challenges associated with both the vehicle-class aerodynamics and the fundamental vortical flows are reviewed, and principles from a hierarchical complexity approach are used to relate flow fundamentals to system-level interests. The work is part of roughly 6-year research program on blunt-leading-edge separation pertinent to UCAVs, and was conducted under the NATO Science and Technology Organization, Applied Vehicle Technology panel.

  15. 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.

  16. Fracture Mechanics Analyses of the Slip-Side Joggle Regions of Wing-Leading-Edge Panels

    NASA Technical Reports Server (NTRS)

    Raju, Ivatury S.; Knight, Norman F., Jr.; Song, Kyongchan; Phillips, Dawn R.

    2011-01-01

    The Space Shuttle wing-leading edge consists of panels that are made of reinforced carbon-carbon. Coating spallation was observed near the slip-side region of the panels that experience extreme heating. To understand this phenomenon, a root-cause investigation was conducted. As part of that investigation, fracture mechanics analyses of the slip-side joggle regions of the hot panels were conducted. This paper presents an overview of the fracture mechanics analyses.

  17. Visualization of leading edge vortices on a series of flat plate delta wings

    NASA Technical Reports Server (NTRS)

    Payne, Francis M.; Ng, T. Terry; Nelson, Robert C.

    1991-01-01

    A summary of flow visualization data obtained as part of NASA Grant NAG2-258 is presented. During the course of this study, many still and high speed motion pictures were taken of the leading edge vortices on a series of flat plate delta wings at varying angles of attack. The purpose is to present a systematic collection of photographs showing the state of vortices as a function of the angle of attack for the four models tested.

  18. Increased heat transfer to a cylindrical leading edge due to spanwise variations in the freestream velocity

    NASA Technical Reports Server (NTRS)

    Rigby, D. L.; Vanfossen, G. J.

    1991-01-01

    The present study numerically demonstrates how small spanwise variations in velocity upstream of a body can cause relatively large increases in the spanwise-averaged heat transfer to the leading edge. Vorticity introduced by spanwise variations, first decays as it drifts downstream, then amplifies in the stagnation region as a result of vortex stretching. This amplification can cause a periodic array of 3 D structures, similar to horseshoe vortices, to form. The numerical results indicate that, for the given wavelength, there is an amplitude threshold below which a structure does not form. A one-dimensional analysis, to predict the decay of vorticity in the absence of the body, in conjunction with the full numerical results indicated that the threshold is more accurately stated as minimum level of vorticity required in the leading edge region for a structure to form. It is possible, using the one-dimensional analysis, to compute an optimum wavelength in terms of the maximum vorticity reaching the leading edge region for given amplitude. A discussion is presented which relates experimentally observed trends to the trends of the present phenomena.

  19. A theory for the core flow of leading-edge vortices

    NASA Technical Reports Server (NTRS)

    Luckring, James M.

    1986-01-01

    Separation-induced leading-edge vortices can dominate the flow about slender wings at moderate to high angles of attack, often with favorable aerodynamic effects. However, at the high angles of attack which are desirable for takeoff and landing as well as subsonic-transonic maneuver the vortices can breakdown or burst in the vicinity of the aircraft causing many adverse effects; these include lift loss, pitchup, and buffet. The flow in the core of leading-edge vortices is generally affiliated with the vortex breakdown phenomenon. A theory is presented for the flow in the core of separation-induced, leading-edge vortices at practical Reynolds numbers. The theory is based on matching inner and outer representations of the vortex. The inner representation models continuously distributed vorticity and includes an asymptotic viscous subcore. The outer representation models concentrated spiral sheets of vorticity and is fully three dimensional. A parameter is identified which closely tracks the vortex breakdown stability boundary for delta, arrow, and diamond wings.

  20. Actin filament turnover drives leading edge growth during myelin sheath formation in the central nervous system

    PubMed Central

    Schmitt, Sebastian; Snaidero, Nicolas; Mitkovski, Mišo; Velte, Caroline; Brückner, Bastian R.; Alexopoulos, Ioannis; Czopka, Tim; Jung, Sang Y.; Rhee, Jeong S.; Janshoff, Andreas; Witke, Walter; Schaap, Iwan A.T.; Lyons, David A.; Simons, Mikael

    2016-01-01

    Summary During central nervous system development, oligodendrocytes wrap their plasma membrane around axons to generate multi-lamellar myelin sheaths. To drive growth at the leading edge of myelin at the interface with the axon, mechanical forces are necessary, but the underlying mechanisms are not known. Using an interdisciplinary approach that combines morphological, genetic and biophysical analyses, we identified a key role for actin filament network turnover in myelin growth. At the onset of myelin biogenesis, F-actin is redistributed to the leading edge, where its polymerization-based forces push out non-adhesive and motile protrusions. F-actin disassembly converts protrusions into sheets by reducing surface tension and in turn inducing membrane spreading and adhesion. We identified the actin depolymerizing factor ADF/Cofilin1, which mediates high F-actin turnover rates, as essential factor in this process. We propose that F-actin turnover is the driving force in myelin wrapping by regulating repetitive cycles of leading edge protrusion and spreading. PMID:26166299

  1. Simulation of Flow Through Breach in Leading Edge at Mach 24

    NASA Technical Reports Server (NTRS)

    Gnoffo, Peter A.; Alter, Stephen J.

    2004-01-01

    A baseline solution for CFD Point 1 (Mach 24) in the STS-107 accident investigation was modified to include effects of holes through the leading edge into a vented cavity. The simulations were generated relatively quickly and early in the investigation by making simplifications to the leading edge cavity geometry. These simplifications in the breach simulations enabled: 1) A very quick grid generation procedure; 2) High fidelity corroboration of jet physics with internal surface impingements ensuing from a breach through the leading edge, fully coupled to the external shock layer flow at flight conditions. These simulations provided early evidence that the flow through a 2 inch diameter (or larger) breach enters the cavity with significant retention of external flow directionality. A normal jet directed into the cavity was not an appropriate model for these conditions at CFD Point 1 (Mach 24). The breach diameters were of the same order or larger than the local, external boundary-layer thickness. High impingement heating and pressures on the downstream lip of the breach were computed. It is likely that hole shape would evolve as a slot cut in the direction of the external streamlines. In the case of the 6 inch diameter breach the boundary layer is fully ingested.

  2. LFC leading edge glove flight: Aircraft modification design, test article development and systems integration

    NASA Technical Reports Server (NTRS)

    Etchberger, F. R.

    1983-01-01

    Reduction of skin friction drag by suction of boundary layer air to maintain laminar flow has been known since Prandtl's published work in 1904. The dramatic increases in fuel costs and the potential for periods of limited fuel availability provided the impetus to explore technologies to reduce transport aircraft fuel consumption. NASA sponsored the Aircraft Energy Efficiency (ACEE) program in 1976 to develop technologies to improve fuel efficiency. This report documents the Lockheed-Georgia Company accomplishments in designing and fabricating a leading-edge flight test article incorporating boundary layer suction slots to be flown by NASA on their modified JetStar aircraft. Lockheed-Georgia Company performed as the integration contractor to design the JetStar aircraft modification to accept both a Lockheed and a McDonnell Douglas flight test article. McDonnell Douglas uses a porous skin concept. The report describes aerodynamic analyses, fabrication techniques, JetStar modifications, instrumentation requirements, and structural analyses and testing for the Lockheed test article. NASA will flight test the two LFC leading-edge test articles in a simulated commercial environment over a 6 to 8 month period in 1984. The objective of the flight test program is to evaluate the effectiveness of LFC leading-edge systems in reducing skin friction drag and consequently improving fuel efficiency.

  3. Spanwise gradients in flow speed and leading edge vortex attachment on low Reynolds number wings

    NASA Astrophysics Data System (ADS)

    Jardin, Thierry; David, Laurent

    2014-11-01

    It is now accepted that the aerodynamic performance of low aspect ratio revolving wings, such as insect wings or maple seed membranes, largely relies on sustained leading edge vortex attachment. However, the mechanisms responsible for this sustained attachment are still poorly understood. Here, we compute the Navier-Stokes solution of the flow around a finite wing (i) subjected to a uniform oncoming flow, (ii) subjected to a spanwise varying oncoming flow and (iii) revolving about its root. Therefore, we are able to isolate the mechanisms associated with the spanwise gradient of the local wing speed from those associated with centrifugal and Coriolis effects. We show that over flapping amplitudes typical of insect flight the spanwise gradient of the local wing speed may suffice in maintaining leading edge vortex attachment. We correlate this result with the development of spanwise flow and we evaluate the sensitivity of such a mechanism to the Reynolds number. It is noted, however, that leading edge vortex attachment through the spanwise gradient of the local wing speed does not promote large lift, which ultimately arises from centrifugal and Coriolis effects.

  4. Fibronectin Rigidity Response through Fyn and p130Cas Recruitment to the Leading Edge

    PubMed Central

    Kostic, Ana

    2006-01-01

    Cell motility on extracellular matrices critically depends on matrix rigidity, which affects cell adhesion and formation of focal contacts. Receptor-like protein tyrosine phosphatase alpha (RPTPα) and the αvβ3 integrin form a rigidity-responsive complex at the leading edge. Here we show that the rigidity response through increased spreading and growth correlates with leading edge recruitment of Fyn, but not endogenous c-Src. Recruitment of Fyn requires the palmitoylation site near the N-terminus and addition of that site to c-Src enables it to support a rigidity response. In all cases, the rigidity response correlates with the recruitment of the Src family kinase to early adhesions. The stretch-activated substrate of Fyn and c-Src, p130Cas, is also required for a rigidity response and it is phosphorylated at the leading edge in a Fyn-dependent process. A possible mechanism for the fibronectin rigidity response involves force-dependent Fyn phosphorylation of p130Cas with rigidity-dependent displacement. With the greater displacement of Fyn from p130Cas on softer surfaces, there will be less phosphorylation. These studies emphasize the importance of force and nanometer-level movements in cell growth and function. PMID:16597701

  5. Thermal-structural analysis of the platelet heat-pipe-cooled leading edge of hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Hongpeng, Liu; Weiqiang, Liu

    2016-10-01

    One of the main challenges for the hypersonic vehicle is its thermal protection, more specifically, the cooling of its leading edge. To investigate the feasibility of a platelet heat-pipe-cooled leading edge structure, thermal/stress distributions for steady-state flight conditions are calculated numerically. Studies are carried on for IN718/Na, C-103/Na and T-111/Li compatible material combinations of heat pipe under nominal operations and a central heat pipe failure cases, and the influence of wall thickness on the design robustness is also investigated. And the heat transfer limits (the sonic limit, the capillary limit and the boiling limit) are also computed to check the operation of platelet heat pipes. The results indicate that, with a 15 mm leading edge radius and a wall thickness of 0.5 mm, C-103/Na and T-111/Li combinations of heat pipe is capable of withstanding both nominal and failure conditions for Mach 8 and Mach 10 flight respectively.

  6. Controlled vortical flow on delta wings through unsteady leading edge blowing

    NASA Technical Reports Server (NTRS)

    Lee, K. T.; Roberts, Leonard

    1990-01-01

    The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions.

  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. 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.

  9. 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.

  10. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2012-09-01

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

  12. A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3

    NASA Technical Reports Server (NTRS)

    Williams, M. Susan; Owens, Lewis R., Jr.; Chu, Julio

    1999-01-01

    A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).

  13. Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 3. University Park Campus School

    ERIC Educational Resources Information Center

    Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

    2008-01-01

    This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high schools across the…

  14. Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 1. Academy of the Pacific Rim

    ERIC Educational Resources Information Center

    Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

    2008-01-01

    This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high schools across the…

  15. Case Studies of Leading Edge Small Urban High Schools. Core Academic Strategic Designs: 2. Noble Street Charter High School

    ERIC Educational Resources Information Center

    Shields, Regis Anne; Ireland, Nicole; City, Elizabeth; Derderian, Julie; Miles, Karen Hawley

    2008-01-01

    This report is one of nine detailed case studies of small urban high schools that served as the foundation for the Education Resource Strategies (ERS) report "Strategic Designs: Lessons from Leading Edge Small Urban High Schools." These nine schools were dubbed "Leading Edge Schools" because they stand apart from other high schools across the…

  16. Translocating myonuclei have distinct leading and lagging edges that require kinesin and dynein.

    PubMed

    Folker, Eric S; Schulman, Victoria K; Baylies, Mary K

    2014-01-01

    Nuclei are precisely positioned within all cells, and mispositioned nuclei are a hallmark of many muscle diseases. Myonuclear positioning is dependent on Kinesin and Dynein, but interactions between these motor proteins and their mechanisms of action are unclear. We find that in developing Drosophila muscles, Dynein and Kinesin work together to move nuclei in a single direction by two separate mechanisms that are spatially segregated. First, the two motors work together in a sequential pathway that acts from the cell cortex at the muscle poles. This mechanism requires Kinesin-dependent localization of Dynein to cell cortex near the muscle pole. From this location Dynein can pull microtubule minus-ends and the attached myonuclei toward the muscle pole. Second, the motors exert forces directly on individual nuclei independently of the cortical pathway. However, the activities of the two motors on the nucleus are polarized relative to the direction of myonuclear translocation: Kinesin acts at the leading edge of the nucleus, whereas Dynein acts at the lagging edge of the nucleus. Consistent with the activities of Kinesin and Dynein being polarized on the nucleus, nuclei rarely change direction, and those that do, reorient to maintain the same leading edge. Conversely, nuclei in both Kinesin and Dynein mutant embryos change direction more often and do not maintain the same leading edge when changing directions. These data implicate Kinesin and Dynein in two distinct and independently regulated mechanisms of moving myonuclei, which together maximize the ability of myonuclei to achieve their proper localizations within the constraints imposed by embryonic development.

  17. 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.

  18. Precocious reproduction increases at the leading edge of a mangrove range expansion.

    PubMed

    Dangremond, Emily M; Feller, Ilka C

    2016-07-01

    Climate change-driven shifts in species ranges are ongoing and expected to increase. However, life-history traits may interact with climate to influence species ranges, potentially accelerating or slowing range shifts in response to climate change. Tropical mangroves have expanded their ranges poleward in the last three decades. Here, we report on a shift at the range edge in life-history traits related to reproduction and dispersal. With a common garden experiment and field observations, we show that Rhizophora mangle individuals from northern populations reproduce at a younger age than those from southern populations. In a common garden at the northern range limit, 38% of individuals from the northernmost population were reproductive by age 2, but less than 10% of individuals from the southernmost population were reproductive by the same age, with intermediate amounts of reproduction from intermediate latitudes. Field observations show a similar pattern of younger reproductive individuals toward the northern range limit. We also demonstrate a shift toward larger propagule size in populations at the leading range edge, which may aid seedling growth. The substantial increase in precocious reproduction at the leading edge of the R. mangle range could accelerate population growth and hasten the expansion of mangroves into salt marshes. PMID:27547335

  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. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

  2. 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.

  3. 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.

  4. 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.

  5. Turbine nozzle leading edge film cooling study in a high speed wind tunnel

    SciTech Connect

    Zhang, L.; Pudupatty, R.

    1999-07-01

    To achieve high engine efficiencies, modern industrial gas turbine components operate at temperatures that are higher than the maximum allowable airfoil alloy temperatures. Film cooling is an effective method of reducing the heat load to a turbine airfoil and, when combined with internal cooling, the airfoil cooling requirements can be satisfied. Here, film cooling effectiveness was measured from a four-row shower head on a turbine vane surface using the pressure sensitive paint (PSP) technique. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Three blowing ratios were studied for each of the five freestream conditions: a reference condition, a reduced and an increased Reynolds number condition, and a reduced and an increased Mach number condition. The freestream turbulence intensity was kept at 12.0% for all the tests. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. Film effectiveness was measured on both the pressure and suction surfaces. The film effectiveness increased with blowing ratio on the suction surface and on the pressure surface for blowing ratios from 1.5 to 2.0, but decreased on the pressure surface for blowing ratios from 2.0 to 2.5. The effects of freestream Mach number and Reynolds number on shower head film cooling are also discussed.

  6. How differential deflection of the inboard and outboard leading-edge flaps affected the handling qua

    NASA Technical Reports Server (NTRS)

    2002-01-01

    How differential deflection of the inboard and outboard leading-edge flaps affected the handling qualities of this modified F/A-18A was evaluated during the first check flight in the Active Aeroelastic Wing program at NASA's Dryden Flight Research Center. The Active Aeroelastic Wing program at NASA's Dryden Flight Research Center seeks to determine the advantages of twisting flexible wings for primary maneuvering roll control at transonic and supersonic speeds, with traditional control surfaces such as ailerons and leading-edge flaps used to aerodynamically induce the twist. From flight test and simulation data, the program intends to develop structural modeling techniques and tools to help design lighter, more flexible high aspect-ratio wings for future high-performance aircraft, which could translate to more economical operation or greater payload capability. AAW flight tests began in November, 2002 with checkout and parameter-identification flights. Based on data obtained during the first flight series, new flight control software will be developed and a second series of research flights will then evaluate the AAW concept in a real-world environment. The program uses wings that were modified to the flexibility of the original pre-production F-18 wing. Other modifications include a new actuator to operate the outboard leading edge flap over a greater range and rate, and a research flight control system to host the aeroelastic wing control laws. The Active Aeroelastic Wing Program is jointly funded and managed by the Air Force Research Laboratory and NASA Dryden Flight Research Center, with Boeing's Phantom Works as prime contractor for wing modifications and flight control software development. The F/A-18A aircraft was provided by the Naval Aviation Systems Test Team and modified for its research role by NASA Dryden technicians.

  7. Flexible Metallic Overwrap Concept Developed for On-Orbit Repair of Space Shuttle Orbiter Leading Edges

    NASA Technical Reports Server (NTRS)

    Ritzert, Frank J.; Nesbitt, James A.

    2005-01-01

    The Columbia accident has focused attention on the critical need for on-orbit repair concepts for leading edges in the event that damage is incurred during space shuttle orbiter flight. Damage that is considered as potentially catastrophic for orbiter leading edges ranges from simple cracks to holes as large as 16 in. in diameter. NASA is particularly interested in examining potential solutions for areas of larger damage since such a problem was identified as the cause for the Columbia disaster. One possible idea for the on-orbit repair of the reinforced carbon/carbon (RCC) leading edges is an overwrap concept that would use a metallic sheet flexible enough to conform to the contours of the orbiter and robust enough to protect any problem area from catastrophic failure during reentry. The simplified view of the application of a refractory metal sheet over a mockup of shuttle orbiter panel 9, which experiences the highest temperatures on the shuttle during reentry is shown. The metallic overwrap concept is attractive because of its versatility as well as the ease with which it can be included in an onboard repair kit. Reentry of the orbiter into Earth's atmosphere imposes extreme requirements on repair materials. Temperatures can exceed 1650 C for up to 15 min in the presence of an extremely oxidizing plasma environment. Several other factors are critical, including catalysity, emissivity, and vibrational and aerodynamic loads. Materials chosen for this application will need to be evaluated with respect to high-temperature capability, resistance to oxidation, strength, coefficient of thermal expansion, and thermal conductivity. The temperature profile across panel 9 during reentry as well as a schematic of the overwrap concept itself is shown.

  8. Flow visualization of leading-edge vortex enhancement by spanwise blowing. [swept wings - wind tunnel stability tests

    NASA Technical Reports Server (NTRS)

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

    1975-01-01

    Flow visualization studies were conducted in a small pilot wind tunnel to determine qualitative effects of blowing a discrete jet essentially parallel to the leading edge of a 45 deg-swept trapezoidal wing featuring leading- and trailing-edge flaps. Test parameters included wing angle-of-attack, jet momentum coefficient, leading- and trailing-edge flap deflections, and nozzle chordwise displacement. Results of this study indicate that blowing from a reflection plane over the wing enhances the leading-edge vortex and delays vortex bursting to higher angles-of-attack and greater span distances. Increased blowing rates decrease vortex size, growth rate, and vertical displacement above the wing surface at a given span station and also extend the spanwise effectiveness of lateral blowing. Deflection of a leading-edge flap delays the beneficial effects of spanwise blowing to higher angles-of-attack. Nozzle chordwise locations investigated for the wing with and without leading-edge flap deflection appear equally effective in enhancing the separated leading-edge flow.

  9. Hypersonic aerospace vehicle leading-edge cooling using heat-pipe, transpiration and film-cooling techniques

    SciTech Connect

    Modlin, J.M.

    1991-01-01

    The feasibility of cooling hypersonic-vehicle leading-edge structures exposed to severe aerodynamic surface heat fluxes was studied, using a combination of liquid-metal heat pipes and surface-mass-transfer cooling techniques. A generalized, transient, finite-difference-based hypersonic leading-edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading-edge section. The hypersonic leading-edge cooling model was developed using an existing, experimentally verified heat-pipe model. Then the existing heat-pipe model was modified by adding both transpiration and film-cooling options as new surface boundary conditions. The models used to predict the leading-edge surface heat-transfer reduction effects of the transpiration and film cooling were modifications of more-generalized, empirically based models obtained from the literature. It is concluded that cooling leading-edge structures exposed to severe hypersonic-flight environments using a combination of liquid-metal heat pipe, surface transpiration, and film cooling methods appears feasible.

  10. 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.

  11. 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.

  12. The Leading Edge 250: Oblique wing aircraft configuration project, volume 4

    NASA Technical Reports Server (NTRS)

    Schmidt, Andre; Moore, Peri; Nguyen, Dan; Oganesyan, Petros; Palmer, Charles

    1988-01-01

    The design of a high speed transport aircraft using the oblique wing concept as a part of the High Speed Civil Transport (HSCT) aircraft study is the Leading Edge 250 capable of travelling at Mach 4 with 250 passengers and has a 6,500 nautical mile range. Its innovation lies within its use of the unconventional oblique wing to provide efficient flight at any Mach number. Wave drag is kept to a minimum at high speed, while high lift is attained during critical takeoff and landing maneuvers by varying the sweep of the wing.

  13. Leading-edge vortex research: Some nonplanar concepts and current challenges

    NASA Technical Reports Server (NTRS)

    Campbell, J. F.; Osborn, R. F.

    1986-01-01

    Some background information is provided for the Vortex Flow Aerodynamics Conference and that current slender wing airplanes do not use variable leading edge geometry to improve transonic drag polar is shown. Highlights of some of the initial studies combining wing camber, or flaps, with vortex flow are presented. Current vortex flap studies were reviewed to show that there is a large subsonic data base and that transonic and supersonic generic studies have begun. There is a need for validated flow field solvers to calculate vortex/shock interactions at transonic and supersonic speeds. Many important research opportunities exist for fundamental vortex flow investigations and for designing advanced fighter concepts.

  14. Mechanical Properties and Microstructural Characterization of Particulate Reinforced Diboride Composites for High Temperature Leading Edge Applications

    NASA Technical Reports Server (NTRS)

    Ellerby, Donald T.; Johnson, Sylvia M.; Bull, Jeff; Laub, Bernie; Reuther, James; Kinney, David; Kontinos, Dean; Beckman, Sarah; Stuffle, Kevin; Cull, A. D.; Arnold, Jim (Technical Monitor)

    2001-01-01

    Previous work on refractory diboride composites has shown that these systems have the potential for use in high temperature leading edge applications for reusable reentry vehicles. Experiments in reentry environments have shown that these materials have multiple use temperatures greater than 1900 C. The work to be discussed focuses on three compositions: HfB2/SiC, ZrB2/SiC, and ZrB2/C/SiC. These composites have been hot pressed and their mechanical properties measured at room and elevated temperatures. Extensive microstructural characterization has been conducted on polished cross sections and the fracture surfaces have been examined to determine their failure origins.

  15. 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.

  16. Estimating Blade Section Airloads from Blade Leading-Edge Pressure Measurements

    NASA Technical Reports Server (NTRS)

    vanAken, Johannes M.

    2003-01-01

    The Tilt-Rotor Aeroacoustic Model (TRAM) test in the Duitse-Nederlandse Wind (DNW) Tunnel acquired blade pressure data for forward flight test conditions of a tiltrotor in helicopter mode. Chordwise pressure data at seven radial locations were integrated to obtain the blade section normal force. The present investigation evaluates the use of linear regression analysis and of neural networks in estimating the blade section normal force coefficient from a limited number of blade leading-edge pressure measurements and representative operating conditions. These network models are subsequently used to estimate the airloads at intermediate radial locations where only blade pressure measurements at the 3.5% chordwise stations are available.

  17. Carbon dioxide gas purification and analytical measurement for leading edge 193nm lithography

    NASA Astrophysics Data System (ADS)

    Riddle Vogt, Sarah; Landoni, Cristian; Applegarth, Chuck; Browning, Matt; Succi, Marco; Pirola, Simona; Macchi, Giorgio

    2015-03-01

    The use of purified carbon dioxide (CO2) has become a reality for leading edge 193 nm immersion lithography scanners. Traditionally, both dry and immersion 193 nm lithographic processes have constantly purged the optics stack with ultrahigh purity compressed dry air (UHPCDA). CO2 has been utilized for a similar purpose as UHPCDA. Airborne molecular contamniation (AMC) purification technologies and analytical measurement methods have been extensively developed to support the Lithography Tool Manufacturers purity requirements. This paper covers the analytical tests and characterizations carried out to assess impurity removal from 3.0 N CO2 (beverage grade) for its final utilization in 193 nm and EUV scanners.

  18. Nondestructive Evaluation Tests Performed on Space Shuttle Leading- Edge Materials Subjected to Impact

    NASA Technical Reports Server (NTRS)

    Roth, Don J.; Martin, Richard E.; Bodis, James R.

    2005-01-01

    In support of the space shuttle Return To Flight efforts at the NASA Glenn Research Center, a series of nondestructive evaluation (NDE) tests were performed on reinforced carbon/carbon (RCC) composite panels subjected to ballistic foam impact. The impact tests were conducted to refine and verify analytical models of an external tank foam strike on the space shuttle leading edge. The NDE tests were conducted to quantify the size and location of the resulting damage zone as well as to identify hidden damage.

  19. Initial development of an ablative leading edge for the space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Daforno, G.; Rose, L.; Graham, J.; Roy, P.

    1974-01-01

    A state-of-the-art preliminary design for typical wing areas is developed. Seven medium-density ablators (with/without honeycomb, flown on Apollo, Prime, X15A2) are evaluated. The screening tests include: (1) leading-edge models sequentially subjected to ascent heating, cold soak, entry heating, post-entry pressure fluctuations, and touchdown shock, and (2) virgin/charred models subjected to bondline strains. Two honeycomb reinforced 30 pcf elastomeric ablators were selected. Roughness/recession degradation of low speed aerodynamics appears acceptable. The design, including attachments, substructure and joints, is presented.

  20. Increased heat transfer to elliptical leading edges due to spanwise variations in the freestream momentum: Numerical and experimental results

    NASA Technical Reports Server (NTRS)

    Rigby, D. L.; Vanfossen, G. J.

    1992-01-01

    A study of the effect of spanwise variation in momentum on leading edge heat transfer is discussed. Numerical and experimental results are presented for both a circular leading edge and a 3:1 elliptical leading edge. Reynolds numbers in the range of 10,000 to 240,000 based on leading edge diameter are investigated. The surface of the body is held at a constant uniform temperature. Numerical and experimental results with and without spanwise variations are presented. Direct comparison of the two-dimensional results, that is, with no spanwise variations, to the analytical results of Frossling is very good. The numerical calculation, which uses the PARC3D code, solves the three-dimensional Navier-Stokes equations, assuming steady laminar flow on the leading edge region. Experimentally, increases in the spanwise-averaged heat transfer coefficient as high as 50 percent above the two-dimensional value were observed. Numerically, the heat transfer coefficient was seen to increase by as much as 25 percent. In general, under the same flow conditions, the circular leading edge produced a higher heat transfer rate than the elliptical leading edge. As a percentage of the respective two-dimensional values, the circular and elliptical leading edges showed similar sensitivity to span wise variations in momentum. By equating the root mean square of the amplitude of the spanwise variation in momentum to the turbulence intensity, a qualitative comparison between the present work and turbulent results was possible. It is shown that increases in leading edge heat transfer due to spanwise variations in freestream momentum are comparable to those due to freestream turbulence.

  1. 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

  2. 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

  3. Subsonic balance and pressure investigation of a 60-deg delta wing with leading-edge devices (data report)

    NASA Technical Reports Server (NTRS)

    Rao, D. M.; Tingas, S. A.

    1981-01-01

    The drag reduction potential of leading edge devices on a 60 degree delta wing at high lift was examined. Geometric variations of fences, chordwise slots, pylon type vortex generators, leading edge vortex flaps, and sharp leading edge extensions were tested individually and in specific combinations to improve high-alpha drag performance with a minimum of low-alpha drag penalty. The force, moment, and surface static pressure data for angles of attack up to 23 degrees, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter are documented.

  4. Flow-induced noise control behind bluff bodies with various leading edges using the surface perturbation technique

    NASA Astrophysics Data System (ADS)

    Lu, Z. B.; Halim, D.; Cheng, L.

    2016-05-01

    The present paper is devoted to an investigation on the flow-induced noise control downstream of bluff bodies with various leading edges using the surface perturbation technique. Four typical leading edges used in various engineering applications were studied in this work: the semi-circular, square, 30° symmetric trapezoid and 30° asymmetric trapezoid leading edges. The surface perturbation was created by piezo-ceramic actuators embedded underneath the surface of a bluff body placed in a cross flow. To suppress the flow-induced noise downstream bluff bodies with those leading edges, the surface perturbation technique was implemented. Based on the experiments, a noise reduction in the duct of more than 14.0 dB has been achieved for all leading-edge cases. These results indicated that the vortex shedding and its flow-induced noise have been successfully suppressed by the proposed control scheme. The flow structure alteration around the bluff bodies and the shear layer shift phenomenon observed on the trailing edges were then investigated for interpreting the control mechanism for this flow-induced noise suppression, which were based on the vortex shedding strength suppression and vortex shedding frequency shift phenomenon. The effective control position for various leading edges was also studied for developing optimal control strategies for practical engineering applications.

  5. Leading-edge receptivity to a vortical freestream disturbance: A numerical analysis

    NASA Technical Reports Server (NTRS)

    Buter, Thomas A.; Reed, Helen L.

    1991-01-01

    The receptivity to freestream vorticity of the boundary layer over a flat plate with an elliptic leading edge is investigated numerically. The flow is simulated by solving the incompressible Navier-Stokes system in general curvilinear coordinates with the vorticity and stream function as dependent variables. A finite-difference scheme which is second-order accurate in both space and time is used. As a first step, the steady basic-state solution is computed. Then a small amplitude vortical disturbance is introduced at the upstream boundary and the governing equations are solved time-accurately to evaluate the spatial and temporal growth of the perturbations leading to instability waves (Tollmien-Schlichting waves) inside the boundary layer. Preliminary results for a symmetric, 2-D disturbance reveal the presence of Tollmien-Schlichting waves aft of the flat-plate/ellipse juncture.

  6. Analog filtering methods improve leading edge timing performance of multiplexed SiPMs

    NASA Astrophysics Data System (ADS)

    Bieniosek, M. F.; Cates, J. W.; Grant, A. M.; Levin, C. S.

    2016-08-01

    Multiplexing many SiPMs to a single readout channel is an attractive option to reduce the readout complexity of high performance time of flight (TOF) PET systems. However, the additional dark counts and shaping from each SiPM cause significant baseline fluctuations in the output waveform, degrading timing measurements using a leading edge threshold. This work proposes the use of a simple analog filtering network to reduce the baseline fluctuations in highly multiplexed SiPM readouts. With 16 SiPMs multiplexed, the FWHM coincident timing resolution for single 3~\\text{mm}× 3~\\text{mm}× 20 mm LYSO crystals was improved from 401  ±  4 ps without filtering to 248  ±  5 ps with filtering. With 4 SiPMs multiplexed, using an array of 3~\\text{mm}× 3~\\text{mm}× 20 mm LFS crystals the mean time resolution was improved from 436  ±  6 ps to 249  ±  2 ps. Position information was acquired with a novel binary positioning network. All experiments were performed at room temperature with no active temperature regulation. These results show a promising technique for the construction of high performance multiplexed TOF PET readout systems using analog leading edge timing pickoff.

  7. Two Functionally Distinct Sources of Actin Monomers Supply the Leading Edge of Lamellipodia

    PubMed Central

    Vitriol, Eric A.; McMillen, Laura M.; Kapustina, Maryna; Gomez, Shawn M.; Vavylonis, Dimitrios; Zheng, James Q.

    2015-01-01

    Summary Lamellipodia, the sheet-like protrusions of motile cells, consist of networks of actin filaments (F-actin) regulated by the ordered assembly from and disassembly into actin monomers (G-actin). Traditionally, G-actin is thought to exist as a homogeneous pool. Here, we show that there are two functionally and molecularly distinct sources of G-actin that supply lamellipodial actin networks. G-actin originating from the cytosolic pool requires the monomer binding protein thymosin β4 (Tβ4) for optimal leading edge localization, is targeted to formins, and is responsible for creating an elevated G/F-actin ratio that promotes membrane protrusion. The second source of G-actin comes from recycled lamellipodia F-actin. Recycling occurs independently of Tβ4 and appears to regulate lamellipodia homeostasis. Tβ4-bound G-actin specifically localizes to the leading edge because it doesn’t interact with Arp2/3-mediated polymerization sites found throughout the lamellipodia. These findings demonstrate that actin networks can be constructed from multiple sources of monomers with discrete spatiotemporal functions. PMID:25865895

  8. Analog filtering methods improve leading edge timing performance of multiplexed SiPMs.

    PubMed

    Bieniosek, M F; Cates, J W; Grant, A M; Levin, C S

    2016-08-21

    Multiplexing many SiPMs to a single readout channel is an attractive option to reduce the readout complexity of high performance time of flight (TOF) PET systems. However, the additional dark counts and shaping from each SiPM cause significant baseline fluctuations in the output waveform, degrading timing measurements using a leading edge threshold. This work proposes the use of a simple analog filtering network to reduce the baseline fluctuations in highly multiplexed SiPM readouts. With 16 SiPMs multiplexed, the FWHM coincident timing resolution for single [Formula: see text] mm LYSO crystals was improved from 401  ±  4 ps without filtering to 248  ±  5 ps with filtering. With 4 SiPMs multiplexed, using an array of [Formula: see text] mm LFS crystals the mean time resolution was improved from 436  ±  6 ps to 249  ±  2 ps. Position information was acquired with a novel binary positioning network. All experiments were performed at room temperature with no active temperature regulation. These results show a promising technique for the construction of high performance multiplexed TOF PET readout systems using analog leading edge timing pickoff.

  9. Analytical impact models and experimental test validation for the Columbia shuttle wing leading edge panels.

    SciTech Connect

    Lu, Wei-Yang; Metzinger, Kurt Evan; Gwinn, Kenneth West; Antoun, Bonnie R.; Korellis, John S.

    2004-10-01

    This paper describes the analyses and the experimental mechanics program to support the National Aeronautics and Space Administration (NASA) investigation of the Shuttle Columbia accident. A synergism of the analysis and experimental effort is required to insure that the final analysis is valid - the experimental program provides both the material behavior and a basis for validation, while the analysis is required to insure the experimental effort provides behavior in the correct loading regime. Preliminary scoping calculations of foam impact onto the Shuttle Columbia's wing leading edge determined if enough energy was available to damage the leading edge panel. These analyses also determined the strain-rate regimes for various materials to provide the material test conditions. Experimental testing of the reinforced carbon-carbon wing panels then proceeded to provide the material behavior in a variety of configurations and strain-rates for flown or conditioned samples of the material. After determination of the important failure mechanisms of the material, validation experiments were designed to provide a basis of comparison for the analytical effort. Using this basis, the final analyses were used for test configuration, instrumentation location, and calibration definition in support of full-scale testing of the panels in June 2003. These tests subsequently confirmed the accident cause.

  10. Dynamics Impact Tolerance of Shuttle RCC Leading Edge Panels Using LS-DYNA

    NASA Technical Reports Server (NTRS)

    Fasanella, Edwin L.; Jackson, Karen E.; Lyle, Karen H.; Jones, Lisa E.; Hardy, Robin C.; Spellman, Regina L.; Carney, Kelly S.; Melis, Matthew E.; Stockwell, Alan E.

    2005-01-01

    This paper describes a research program conducted to enable accurate prediction of the impact tolerance of the shuttle Orbiter leading-edge wing panels using physics-based codes such as LS-DYNA, a nonlinear, explicit transient dynamic finite element code. The shuttle leading-edge panels are constructed of Reinforced-Carbon-Carbon (RCC) composite material, which is used because of its thermal properties to protect the shuttle during reentry into the Earth's atmosphere. Accurate predictions of impact damage from insulating foam and other debris strikes that occur during launch required materials characterization of expected debris, including strain-rate effects. First, analytical models of individual foam and RCC materials were validated. Next, analytical models of foam cylinders impacting 6- in. x 6-in. RCC flat plates were developed and validated. LS-DYNA pre-test models of the RCC flat plate specimens established the impact velocity of the test for three damage levels: no-detectable damage, non-destructive evaluation (NDE) detectable damage, or visible damage such as a through crack or hole. Finally, the threshold of impact damage for RCC on representative Orbiter wing panels was predicted for both a small through crack and for NDE-detectable damage.

  11. Dynamic Impact Tolerance of Shuttle RCC Leading Edge Panels using LS-DYNA

    NASA Technical Reports Server (NTRS)

    Fasanella, Edwin; Jackson, Karen E.; Lyle, Karen H.; Jones, Lisa E.; Hardy, Robin C.; Spellman, Regina L.; Carney, Kelly S.; Melis, Matthew E.; Stockwell, Alan E.

    2008-01-01

    This paper describes a research program conducted to enable accurate prediction of the impact tolerance of the shuttle Orbiter leading-edge wing panels using 'physics-based- codes such as LS-DYNA, a nonlinear, explicit transient dynamic finite element code. The shuttle leading-edge panels are constructed of Reinforced-Carbon-Carbon (RCC) composite material, which issued because of its thermal properties to protect the shuttle during re-entry into the Earth's atmosphere. Accurate predictions of impact damage from insulating foam and other debris strikes that occur during launch required materials characterization of expected debris, including strain-rate effects. First, analytical models of individual foam and RCC materials were validated. Next, analytical models of individual foam cylinders impacting 6-in. x 6-in. RCC flat plates were developed and validated. LS-DYNA pre-test models of the RCC flat plate specimens established the impact velocity of the test for three damage levels: no-detectable damage, non-destructive evaluation (NDE) detectable damage, or visible damage such as a through crack or hole. Finally, the threshold of impact damage for RCC on representative Orbiter wing panels was predicted for both a small through crack and for NDE-detectable damage.

  12. Fracture Mechanics Analyses of the Slip-Side Joggle Regions of Wing-Leading Edge Panels

    NASA Technical Reports Server (NTRS)

    Raju, Ivatury S.; Knight, Norman F., Jr.; Song, Kyongchan; Phillips, Dawn R.

    2010-01-01

    The Space Shuttle Orbiter wing comprises of 22 leading edge panels on each side of the wing. These panels are part of the thermal protection system that protects the Orbiter wings from extreme heating that take place on the reentry in to the earth atmosphere. On some panels that experience extreme heating, liberation of silicon carbon (SiC) coating was observed on the slip side regions of the panels. Global structural and local fracture mechanics analyses were performed on these panels as a part of the root cause investigation of this coating liberation anomaly. The wing-leading-edge reinforced carbon-carbon (RCC) panels, Panel 9, T-seal 10, and Panel 10, are shown in Figure 1 and the progression of the stress analysis models is presented in Figure 2. The global structural analyses showed minimal interaction between adjacent panels and the T-seal that bridges the gap between the panels. A bounding uniform temperature is applied to a representative panel and the resulting stress distribution is examined. For this loading condition, the interlaminar normal stresses showed negligible variation in the chord direction and increased values in the vicinity of the slip-side joggle shoulder. As such, a representative span wise slice on the panel can be taken and the cross section can be analyzed using plane strain analysis.

  13. Some Effects of Leading-Edge Sweep on Boundary-Layer Transition at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Chapman, Gray T.

    1961-01-01

    The effects of crossflow and shock strength on transition of the laminar boundary layer behind a swept leading edge have been investigated analytically and with the aid of available experimental data. An approximate method of determining the crossflow Reynolds number on a leading edge of circular cross section at supersonic speeds is presented. The applicability of the critical crossflow criterion described by Owen and Randall for transition on swept wings in subsonic flow was examined for the case of supersonic flow over swept circular cylinders. A wide range of applicability of the subsonic critical values is indicated. The corresponding magnitude of crossflow velocity necessary to cause instability on the surface of a swept wing at supersonic speeds was also calculated and found to be small. The effects of shock strength on transition caused by Tollmien-Schlichting type of instability are discussed briefly. Changes in local Reynolds number, due to shock strength, were found analytically to have considerably more effect on transition caused by Tollmien-Schlichting instability than on transition caused by crossflow instability. Changes in the mechanism controlling transition from Tollmien-Schlichting instability to crossflow instability were found to be possible as a wing is swept back and to result in large reductions in the length of laminar flow.

  14. Reynolds number effects on leading edge vortex development on a waving wing

    NASA Astrophysics Data System (ADS)

    Jones, A. R.; Babinsky, H.

    2011-07-01

    The waving wing experiment is a fully three-dimensional simplification of the flapping wing motion observed in nature. The spanwise velocity gradient and wing starting and stopping acceleration that exist on an insect-like flapping wing are generated by rotational motion of a finite span wing. The flow development around a waving wing at Reynolds number between 10,000 and 60,000 has been studied using flow visualization and high-speed PIV to capture the unsteady velocity field. Lift and drag forces have been measured over a range of angles of attack, and the lift curve shape was similar in all cases. A transient high-lift peak approximately 1.5 times the quasi-steady value occurred in the first chord length of travel, caused by the formation of a strong attached leading edge vortex. This vortex appears to develop and shed more quickly at lower Reynolds numbers. The circulation of the leading edge vortex has been measured and agrees well with force data.

  15. Segregation of leading-edge and uropod components into specific lipid rafts during T cell polarization

    PubMed Central

    Gómez-Moutón, Concepción; Abad, Jose Luis; Mira, Emilia; Lacalle, Rosa Ana; Gallardo, Eduard; Jiménez-Baranda, Sonia; Illa, Isabel; Bernad, Antonio; Mañes, Santos; Martínez-A., Carlos

    2001-01-01

    Redistribution of specialized molecules in migrating cells develops asymmetry between two opposite cell poles, the leading edge and the uropod. We show that acquisition of a motile phenotype in T lymphocytes results in the asymmetric redistribution of ganglioside GM3- and GM1-enriched raft domains to the leading edge and to the uropod, respectively. This segregation to each cell pole parallels the specific redistribution of membrane proteins associated to each raft subfraction. Our data suggest that raft partitioning is a major determinant for protein redistribution in polarized T cells, as ectopic expression of raft-associated proteins results in their asymmetric redistribution, whereas non-raft-partitioned mutants of these proteins are distributed homogeneously in the polarized cell membrane. Both acquisition of a migratory phenotype and SDF-1α-induced chemotaxis are cholesterol depletion-sensitive. Finally, GM3 and GM1 raft redistribution requires an intact actin cytoskeleton, but is insensitive to microtubule disruption. We propose that membrane protein segregation not only between raft and nonraft domains but also between distinct raft subdomains may be an organizational principle that mediates redistribution of specialized molecules needed for T cell migration. PMID:11493690

  16. Experimental Study of Shock Wave Interference Heating on a Cylindrical Leading Edge. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Wieting, Allan R.

    1987-01-01

    An experimental study of shock wave interference heating on a cylindrical leading edge representative of the cowl of a rectangular hypersonic engine inlet at Mach numbers of 6.3, 6.5, and 8.0 is presented. Stream Reynolds numbers ranged from 0.5 x 106 to 4.9 x 106 per ft. and stream total temperature ranged from 2100 to 3400 R. The model consisted of a 3" dia. cylinder and a shock generation wedge articulated to angles of 10, 12.5, and 15 deg. A fundamental understanding was obtained of the fluid mechanics of shock wave interference induced flow impingement on a cylindrical leading edge and the attendant surface pressure and heat flux distributions. The first detailed heat transfer rate and pressure distributions for two dimensional shock wave interference on a cylinder was provided along with insight into the effects of specific heat variation with temperature on the phenomena. Results show that the flow around a body in hypersonic flow is altered significantly by the shock wave interference pattern that is created by an oblique shock wave from an external source intersecting the bow shock wave produced in front of the body.

  17. Space Shuttle Orbiter Wing-Leading-Edge Panel Thermo-Mechanical Analysis for Entry Conditions

    NASA Technical Reports Server (NTRS)

    Knight, Norman F., Jr.; Song, Kyongchan; Raju, Ivatury S.

    2010-01-01

    Linear elastic, thermo-mechanical stress analyses of the Space Shuttle Orbiter wing-leading-edge panels is presented for entry heating conditions. The wing-leading-edge panels are made from reinforced carbon-carbon and serve as a part of the overall thermal protection system. Three-dimensional finite element models are described for three configurations: integrated configuration, an independent single-panel configuration, and a local lower-apex joggle segment. Entry temperature conditions are imposed and the through-the-thickness response is examined. From the integrated model, it was concluded that individual panels can be analyzed independently since minimal interaction between adjacent components occurred. From the independent single-panel model, it was concluded that increased through-the-thickness stress levels developed all along the chord of a panel s slip-side joggle region, and hence isolated local joggle sections will exhibit the same trend. From the local joggle models, it was concluded that two-dimensional plane-strain models can be used to study the influence of subsurface defects along the slip-side joggle region of these panels.

  18. Analytical observations on the aerodynamics of a delta wing with leading edge flaps

    NASA Technical Reports Server (NTRS)

    Oh, S.; Tavella, D.

    1986-01-01

    The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

  19. Leading edge vortex dynamics on a pitching delta wing. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Lemay, Scott P.

    1988-01-01

    The leading edge flow structure was investigated on a 70 deg flat plate delta wing which was pitched about its 1/2 chord position, to increase understanding of the high angle of attack aerodynamics on an unsteady delta wing. The wing was sinusoidally pitched at reduced frequencies ranging from k being identical with 2pi fc/u = 0.05 to 0.30 at chord Reynolds numbers between 90,000 and 350,000, for angle of attack ranges of alpha = 29 to 39 deg and alpha = 0 to 45 deg. The wing was also impulsively pitched at an approximate rate of 0.7 rad/s. During these dynamic motions, visualization of the leading edge vorticies was obtained by entraining titanium tetrachloride into the flow at the model apex. The location of vortex breakdown was recorded using 16mm high speed motion picture photography. When the wing was sinusoidally pitched, a hysteresis was observed in the location of breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing, and the phase lag between model motion and breakdown location were also determined. When the wing was impulsively pitched, several convective times were required for the vortex flow to reach a steady state. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

  20. Metallic Concepts for Repair of Reinforced Carbon-Carbon Space Shuttle Leading Edges

    NASA Technical Reports Server (NTRS)

    Ritzert, Frank; Nesbitt, James

    2007-01-01

    The Columbia accident has focused attention on the critical need for on-orbit repair concepts for wing leading edges in the event that potentially catastrophic damage is incurred during Space Shuttle Orbiter flight. The leading edge of the space shuttle wings consists of a series of eleven panels on each side of the orbiter. These panels are fabricated from reinforced carbon-carbon (RCC) which is a light weight composite with attractive strength at very high temperatures. The damage that was responsible for the loss of the Colombia space shuttle was deemed due to formation of a large hole in one these RCC leading edge panels produced by the impact of a large piece of foam. However, even small cracks in the RCC are considered as potentially catastrophic because of the high temperature re-entry environment. After the Columbia accident, NASA has explored various means to perform on-orbit repairs in the event that damage is sustained in future shuttle flights. Although large areas of damage, such as that which doomed Columbia, are not anticipated to re-occur due to various improvements to the shuttle, especially the foam attachment, NASA has also explored various options for both small and large area repair. This paper reports one large area repair concept referred to as the "metallic over-wrap." Environmental conditions during re-entry of the orbiter impose extreme requirements on the RCC leading edges as well as on any repair concepts. These requirements include temperatures up to 3000 F (1650 C) for up to 15 minutes in the presence of an extremely oxidizing plasma environment. Figure 1 shows the temperature profile across one panel (#9) which is subject to the highest temperatures during re-entry. Although the RCC possesses adequate mechanical strength at these temperatures, it lacks oxidation resistance. Oxidation protection is afforded by converting the outer layers of the RCC to SiC by chemical vapor deposition (CVD). At high temperatures in an oxidizing

  1. 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.

  2. 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.

  3. Effects of increased leading-edge thickness on performance of a transonic rotor blade. [in single stage transonic compressor

    NASA Technical Reports Server (NTRS)

    Reid, L.; Urasek, D. C.

    1972-01-01

    A single-stage transonic compressor was tested with two rotor blade leading-edge configurations to investigate the effect of increased leading-edge thickness on the performance of a transonic blade row. The original rotor blade configuration was modified by cutting back the leading edge sufficiently to double the blade leading-edge thickness and thus the blade gap blockage in the tip region. At design speed this modification resulted in a decrease in rotor overall peak efficiency of four points. The major portion of this decrement in rotor overall peak efficienty was attributed to the flow conditions in the outer 30 percent of the blade span. At 70 and 90 percent of design speed, the modification had very little effect on rotor overall performance.

  4. Effect of canard position and wing leading-edge flap deflection on wing buffet at transonic speeds

    NASA Technical Reports Server (NTRS)

    Gloss, B. B.; Henderson, W. P.; Huffman, J. K.

    1974-01-01

    A generalized wind-tunnel model, with canard and wing planform typical of highly maneuverable aircraft, was tested. The addition of a canard above the wing chord plane, for the configuration with leading-edge flaps undeflected, produced substantially higher total configuration lift coefficients before buffet onset than the configuration with the canard off and leading-edge flaps undeflected. The wing buffet intensity was substantially lower for the canard-wing configuration than the wing-alone configuration. The low-canard configuration generally displayed the poorest buffet characteristics. Deflecting the wing leading-edge flaps substantially improved the wing buffet characteristics for canard-off configurations. The addition of the high canard did not appear to substantially improve the wing buffet characteristics of the wing with leading-edge flaps deflected.

  5. 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.

  6. 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.

  7. Quasi-chemostat behavior in the leading edge of B. subtilis biofilms

    NASA Astrophysics Data System (ADS)

    Srinivasan, Siddarth; Mahadevan, Lakshminarayanan; Rubinstein, Shmuel

    2015-11-01

    Bacillus subtilis is a gram positive bacterium that is a model system commonly used to study biofilm formation. By performing wide-field time-lapse microscopy on a fluorescently labeled B. subtilis strain, we observe a well defined steady boundary layer at the edge of a biofilm growing on an nutrient infused agar gel substrate, within which the outward radial expansion growth predominantly occurs. Using distinct fluorescent protein markers as proxies of gene expression, we quantitatively measure how the width, velocity and ratio of motile cell to matrix cell phenotypes within this boundary layer responds to changes in environmental conditions (such as substrate agar percentage & temperature). We further propose that the steady state at the leading edge can be interpreted as a quasi-chemostat which may enable well controlled response experiments on a colony scale. Finally, we show that for low agar concentration (0.5 wt%), the cells exhibit swarming behavior, whose dynamics and swimming velocities are characterized using differential dynamic microscopy. We show the swarming state is associated with an unstable front which gives rise to fingering and branching growth patterns, illustrating the varied morphological response of the biofilm to environmental conditions

  8. A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

    NASA Technical Reports Server (NTRS)

    Mehrotra, S. C.; Lan, C. E.

    1978-01-01

    The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

  9. Performance Evaluation of Leading Edge Slats on Rigid Wing Sail Catamarans

    NASA Astrophysics Data System (ADS)

    Johnson, Chelsea; O'Neill, Charles

    2015-11-01

    Rigid wing sails have created the fastest catamarans in history, however with the addition of a leading edge slat higher lift and faster speeds may be achieved. Slats are currently used on airplane wings to increase lift, but have not been implemented on a rigid wing sail catamaran. Using 3D modeling and computational fluid dynamics software, this research investigates the effect that slats have on the performance of rigid wing sail catamarans. Aerodynamics and hydrodynamics form the basis of the research. The preliminary results show an increase in the coefficient of lift for sail models with slats over sail models without slats, allowing the catamaran to perform at higher speeds. The ability of the slat to rotate has also been identified as a key factor in increasing the benefit of the slat. This work was supported by NSF site award 1358991.

  10. Experimental investigation on the composite cooling of a semicylinder leading edge

    NASA Astrophysics Data System (ADS)

    Cheng, Ji-Rui; Ji, Hong-Hu

    Composite cooling of the leading edge region of a semicylinder simulating a turbine vane was studied experimentally. The cylinder's inner surface served as an impingement target for studying impingement cooling; the outer surface was used to study the film cooling effectiveness. The equipment for composite cooling was a sucked-wind tunnel with the test section of 140 x 200-sq m in cross section and with transparent windows for flow visualization. Three test models and five different impingement tubes were used. The effects of geometrical parameters of the test models and impingement tubes and of the location and direction of the film holes on both the impingement and the film cooling are described.

  11. Measurement of the temperature field downstream of simulated leading-edge film-cooling holes

    NASA Astrophysics Data System (ADS)

    Mee, D. J.; Ireland, P. T.; Bather, S.

    Experiments to measure the temperature field downstream of simulated leading-edge-region film-cooling holes were performed in an 11 m/s wind tunnel flow. Heated air was passed to a hollow 140 mm diameter cylinder in which three 10.5 mm diameter, spanwise-inclined, film-cooling holes had been machined. A fine nylon mesh, coated with encapsulated thermochromic liquid crystals, was used to measure temperature contours downstream of the holes by moving the mesh relative to the holes and adjusting the power to the air heater. The measurements indicate the extent of the lateral spreading of the coolant gas and show the influence of hole location and coolant mass flow rate on film trajectory and spreading.

  12. Computational study of transition front on a swept wing leading-edge model

    NASA Technical Reports Server (NTRS)

    Iyer, Venkit; Spall, Robert E.; Dagenhart, J. R.

    1992-01-01

    The e(N) method is employed with Navier-Stokes calculations to study 3D supersonic boundary layers at the leading-edge region of swept wings with attention given to the conditions for linear stability. Mean-flow profiles are computed as solutions to the Navier-Stokes equations, and the e(N) method is used to identify the onset of the transition. Specific treatment is given to the identification of parameters which affect the transition such as free-stream Reynolds number, wall cooling, and boundary-layer suction levels. The boundary layer can be stabilized to below N=10 levels in all locations by utilizing distributed suction, and enhanced stabilization can also be achieved by employing wall cooling and reduced free-stream Reynolds numbers. This study presents key calculations for corresponding experimental investigations by shedding light on requirements for test conditions and measurement requirements.

  13. Multiple leading edge vortices of unexpected strength in freely flying hawkmoth

    PubMed Central

    Johansson, L. Christoffer; Engel, Sophia; Kelber, Almut; Heerenbrink, Marco Klein; Hedenström, Anders

    2013-01-01

    The Leading Edge Vortex (LEV) is a universal mechanism enhancing lift in flying organisms. LEVs, generally illustrated as a single vortex attached to the wing throughout the downstroke, have not been studied quantitatively in freely flying insects. Previous findings are either qualitative or from flappers and tethered insects. We measure the flow above the wing of freely flying hawkmoths and find multiple simultaneous LEVs of varying strength and structure along the wingspan. At the inner wing there is a single, attached LEV, while at mid wing there are multiple LEVs, and towards the wingtip flow separates. At mid wing the LEV circulation is ~40% higher than in the wake, implying that the circulation unrelated to the LEV may reduce lift. The strong and complex LEV suggests relatively high flight power in hawmoths. The variable LEV structure may result in variable force production, influencing flight control in the animals. PMID:24253180

  14. Leading edge serrations which reduce the noise of low-speed rotors

    NASA Technical Reports Server (NTRS)

    Soderman, P. T.

    1973-01-01

    Acoustic effects of serrated brass strips attached near the leading edges of two different size rotors were investigated. The two bladed rotors were tested in hover. Rotor rotational speed, blade angle, serration shape, and serration position were varied. The serrations were more effective as noise suppressors at rotor tip speeds less than 135 m/sec (444 ft/sec) than at higher speeds. high frequency noise was reduced but the low frequency rotational noise was little affected. Noise reductions from 4 to 8 db overall sound pressure level and 3 to 17 db in the upper octave bands were achieved on the 1.52 m (5.0 ft) diameter rotor. Noise reductions up to 4 db overall sound pressure level were measured for the 2.59 m (8.5 ft) diameter rotor at some conditions.

  15. 3D characterization of leading-edge vortex formation and growth

    NASA Astrophysics Data System (ADS)

    Onoue, Kyohei; Breuer, Kenneth

    2015-11-01

    We examine the vorticity transport mechanisms responsible for regulating the stability and strength of the leading-edge vortex (LEV) on rapidly pitching plates with different planforms (swept vs. rectangular) in a uniform airflow. All experiments are carried out using a cyber-physical experimental setup (Onoue et al., 2015, JFS vol. 55) and synchronized 3D PIV measurements. In the case of a swept wing, two distinct regions of intense spanwise flow are observed around the LEV centroid--a feature conspicuously absent on a rectangular pitching plate. The interaction between these spanwise flows and the LEV core seems to play a role in prolonging the LEV residence time at the cost of the vortex circulation growth rate and magnitude. Detailed control volume analysis is performed to elucidate the flow physics at work. This research is funded by the Air Force Office of Scientific Research (AFOSR).

  16. An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

    NASA Technical Reports Server (NTRS)

    Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

    1983-01-01

    An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

  17. Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing

    NASA Technical Reports Server (NTRS)

    Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

    1983-01-01

    Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.

  18. Application of superplastically formed and diffusion bonded aluminum to a laminar flow control leading edge

    NASA Technical Reports Server (NTRS)

    Goodyear, M. D.

    1987-01-01

    NASA sponsored the Aircraft Energy Efficiency (ACEE) program in 1976 to develop technologies to improve fuel efficiency. Laminar flow control was one such technology. Two approaches for achieving laminar flow were designed and manufactured under NASA sponsored programs: the perforated skin concept used at McDonnell Douglas and the slotted design used at Lockheed-Georgia. Both achieved laminar flow, with the slotted design to a lesser degree (JetStar flight test program). The latter design had several fabrication problems concerning springback and adhesive flow clogging the air flow passages. The Lockheed-Georgia Company accomplishments is documented in designing and fabricating a small section of a leading edge article addressing a simpler fabrication method to overcome the previous program's manufacturing problems, i.e., design and fabrication using advanced technologies such as diffusion bonding of aluminum, which has not been used on aerospace structures to date, and the superplastic forming of aluminum.

  19. Leading-edge transition and relaminarization phenomena on a subsonic high-lift system

    NASA Technical Reports Server (NTRS)

    Van Dam, C. P.; Vijgen, P. M. H. W.; Yip, L. P.; Potter, R. C.

    1993-01-01

    Boundary-layer transition and relaminarization may have a critical effect on the flow development about multielement high-lift systems of subsonic transport aircraft with swept wings. The purpose of this paper is a study of transition phenomena in the leading-edge region of the various elements of a high-lift system. The flow phenomena studied include transition of the attachment-line flow, relaminarization, and crossflow instability. The calculations are based on pressure distributions measured in flight on the NASA Transport Systems Research Vehicle (Boeing 737-100) at a wing station where the flow approximated infinite swept wing conditions. The results indicate that significant regions of laminar flow can exist on all flap elements in flight. In future flight experiments the extent of these regions will be measured, and the transition mechanisms and the effect of laminar flow on the high-lift characteristics of the multi-element system will be further explored.

  20. Vibration and sound of an elastic wing actuated at its leading edge

    NASA Astrophysics Data System (ADS)

    Manela, A.

    2012-01-01

    The motion and sound of a thin elastic plate, subject to uniform low-Mach flow and actuated at its leading edge, is studied. The linearized response to arbitrary small-amplitude translation and rotation is analyzed using Fourier decomposition of the forcing signal. Both periodic (sinusoidal) and non-periodic ("step-jump") actuations are investigated. When the frequency spectrum of the forcing signal contains an eigenfrequency Ωres of the unforced system, a resonance motion is excited and the plate oscillates at the corresponding eigenmode. The dynamical description is applied to formulate the acoustic problem, where the sources of sound include the plate velocity and fluid vorticity. Acoustic radiation of a dipole type is calculated and discussed in the limit where the plate is acoustically compact. In the case of sinusoidal excitation, plate elasticity has two opposite effects on sound radiation, depending on the forcing frequency: at frequencies close to Ωres, the near-resonance motion results in the generation of high sound levels; however, at frequencies far from Ωres, plate elasticity reduces the amplitude of plate deflection (compared to that of a rigid plate), leading to noise reduction. In the case of non-periodic actuation, the plate-fluid system amplifies those frequencies that are closest to Ωres, which, in turn, dominate the acoustic signature. The results identify the trailing edge noise as the main source of sound, dominating the sound generated by direct plate motion. We suggest the present theory as a preliminary tool for examining the acoustic signature of flapping flight, common in insects and flapping micro-air-vehicles.

  1. Cavitation on a semicircular leading-edge plate and NACA0015 hydrofoil: Visualization and velocity measurement

    NASA Astrophysics Data System (ADS)

    Kravtsova, A. Yu.; Markovich, D. M.; Pervunin, K. S.; Timoshevskii, M. V.; Hanjalić, K.

    2014-12-01

    Using high-speed visualization and particle image velocimetry (PIV), cavitating flows near a plane plate with a rounded leading edge and NACA0015 hydrofoil at angles of attack from 0° to 9° are studied. In the experiments, several known types of cavitation, as well as some differences, were detected with variation of the cavitation number. In particular, at small angles of attack (up to 3°), cavitation on the plate appears in the form of a streak array; on the hydrofoil, it appears in the form of individual bubbles. For the NACA0015 hydrofoil, isolated and intermittent streaks are divided and grow in regimes with developed cavitation; then, however, they merge in bubble clouds and form an extremely regular cellular structure. With an increase in the angle of attack to 9°, the structure of the cavitation cavity on the hydrofoil is changed by the streak structure, like in the case with the plate. In this work, it is shown that PIV permits one to measure the velocity in cavitating flows, in particular, within the gas-vapor phase. It was established from the analysis of distributions of the average flow velocity and moments of velocity fluctuations that the cavitation generation is caused by the development of the carrier fluid flow near the leading edge of the hydrofoil. Down the stream, however, the flow structure strongly depends on the cavitation regime, which is seen from the comparison of the distributions with the case of a single-phase flow. The presented measurements qualitatively verify general trends and show some quantitative distinctions for the two considered flowpast bodies.

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

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

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

  3. 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.

  4. 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.

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

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Pinkerton, Robert M

    1931-01-01

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

  6. Theoretical and Experimental Unsteady Aerodynamics Compared for a Linear Oscillating Cascade With a Supersonic Leading-Edge Locus

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2005-01-01

    Experimental data were obtained to help validate analytical and computational fluid dynamics (CFD) codes used to compute unsteady cascade aerodynamics in a supersonicaxial- flow regime. Results from two analytical codes and one CFD code were compared with experimental data. One analytical code did not account for airfoil thickness or camber; another, using piston theory (piston code), accounted for thickness and camber upstream of the first shockwave/airfoil impingement locations. The Euler CFD code accounted fully for airfoil shape.

  7. Performance of advanced wind turbine airfoils with vortex generators

    SciTech Connect

    Wetzel, K.K.; Farokhi, S.

    1995-12-31

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

  8. 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%.

  9. A comparison of experimental and calculated thin-shell leading-edge buckling due to thermal stresses

    NASA Technical Reports Server (NTRS)

    Jenkins, Jerald M.

    1988-01-01

    High-temperature thin-shell leading-edge buckling test data are analyzed using NASA structural analysis (NASTRAN) as a finite element tool for predicting thermal buckling characteristics. Buckling points are predicted for several combinations of edge boundary conditions. The problem of relating the appropriate plate area to the edge stress distribution and the stress gradient is addressed in terms of analysis assumptions. Local plasticity was found to occur on the specimen analyzed, and this tended to simplify the basic problem since it effectively equalized the stress gradient from loaded edge to loaded edge. The initial loading was found to be difficult to select for the buckling analysis because of the transient nature of thermal stress. Multiple initial model loadings are likely required for complicated thermal stress time histories before a pertinent finite element buckling analysis can be achieved. The basic mode shapes determined from experimentation were correctly identified from computation.

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

    NASA Astrophysics Data System (ADS)

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

    2014-07-01

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

  11. A Mesh Refinement Study on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

    NASA Technical Reports Server (NTRS)

    Fasanella, Edwin L.; Jackson, Karen E.; Lyle, Karen H.; Spellman, Regina L.

    2006-01-01

    A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0-in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at four discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density.

  12. A Three-Dimensional Solution of Flows over Wings with Leading-Edge Vortex Separation. Part 1: Engineering Document

    NASA Technical Reports Server (NTRS)

    Brune, G. W.; Weber, J. A.; Johnson, F. T.; Lu, P.; Rubbert, P. E.

    1975-01-01

    A method of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings with leading-edge vortex separation in incompressible flow is presented. The method employs an inviscid flow model in which the wing and the rolled-up vortex sheets are represented by piecewise, continuous quadratic doublet sheet distributions. The Kutta condition is imposed on all wing edges. Computed results are compared with experimental data and with the predictions of the leading-edge suction analogy for a selected number of wing planforms over a wide range of angle of attack. These comparisons show the method to be very promising, capable of producing not only force predictions, but also accurate predictions of detailed surface pressure distributions, loads, and moments.

  13. Influence of leading edge bluntness on hypersonic flow in a generic internal-compression inlet

    NASA Astrophysics Data System (ADS)

    Borovoy, V.; Egorov, I.; Mosharov, V.; Radchenko, V.; Skuratov, A.; Struminskaya, I.

    2015-06-01

    Flow and heat transfer inside a generic inlet are investigated experimentally. The cross section of the inlet is rectangular. The inlet is installed on a flat plat at a significant distance from the leading edge. The experiments are performed in TsAGI wind tunnel UT-1M working in the Ludwieg tube mode at Mach number M∞ = 5 and Reynolds numbers (based on the plate length L = 320 mm) Re∞L = 23 · 106 and 13 · 106. Steady flow duration is 40 ms. Optical panoramic methods are used for investigation of flow outside and inside the inlet as well. For this purpose, the cowl and one of two compressing wedges are made of a transparent material. Heat flux distribution is measured by thin luminescent Temperature Sensitive Paint (TSP). Surface flow and shear stress visualization is performed by viscous oil containing luminophor particles. The investigation shows that at high contraction ratio of the inlet, an increase of plate or cowl bluntness to some critical value leads to sudden change of the flow structure.

  14. Spin-dependent transport for armchair-edge graphene nanoribbons between ferromagnetic leads.

    PubMed

    Zhou, Benhu; Chen, Xiongwen; Zhou, Benliang; Ding, Kai-He; Zhou, Guanghui

    2011-04-01

    We theoretically investigate the spin-dependent transport for the system of an armchair-edge graphene nanoribbon (AGNR) between two ferromagnetic (FM) leads with arbitrary polarization directions at low temperatures, where a magnetic insulator is deposited on the AGNR to induce an exchange splitting between spin-up and -down carriers. By using the standard nonequilibrium Green's function (NGF) technique, it is demonstrated that the spin-resolved transport property for the system depends sensitively on both the width of AGNR and the polarization strength of FM leads. The tunneling magnetoresistance (TMR) around zero bias voltage possesses a pronounced plateau structure for a system with semiconducting 7-AGNR or metallic 8-AGNR in the absence of exchange splitting, but this plateau structure for the 8-AGNR system is remarkably broader than that for the 7-AGNR one. Interestingly, an increase of the exchange splitting Δ suppresses the amplitude of the structure for the 7-AGNR system. However, the TMR is much enhanced for the 8-AGNR system under a bias amplitude comparable to the splitting strength. Further, the current-induced spin-transfer torque (STT) for the 7-AGNR system is systematically larger than that for the 8-AGNR one. The findings here suggest the design of GNR-based spintronic devices by using a metallic AGNR, but it is more favorable to fabricate a current-controlled magnetic memory element by using a semiconducting AGNR.

  15. 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.

  16. A simplified method for thermal analysis of a cowl leading edge subject to intense local shock-wave-interference heating

    NASA Astrophysics Data System (ADS)

    McGowan, David M.; Camarda, Charles J.; Scotti, Stephen J.

    1992-03-01

    Type IV shock wave interference heating on a blunt body causes extremely intense heating over a very localized region of the body. An analytical solution is presented to a heat transfer problem that approximates the shock wave interference heating of an engine cowl leading edge of the National Aero-Space Plane. The problem uses a simplified geometry to represent the leading edge. An analytical solution is developed that provides a means for approximating maximum temperature differences between the outer and inner surface temperatures of the leading edge. The solution is computationally efficient and, as a result, is well suited for conceptual and preliminary design or trade studies. Transient and steady state analyses are conducted, and results obtained from the analytical solution are compared with results of 2-D thermal finite element analyses over a wide range of design parameters. Isotropic materials as well as laminated composite materials are studied. Results of parametric studies are presented to indicate the effects of the thickness of the cowl leading edge and the width of the region heated by the shock wave interference on the thermal response of the leading edge.

  17. A simplified method for thermal analysis of a cowl leading edge subject to intense local shock-wave-interference heating

    NASA Technical Reports Server (NTRS)

    Mcgowan, David M.; Camarda, Charles J.; Scotti, Stephen J.

    1992-01-01

    Type IV shock wave interference heating on a blunt body causes extremely intense heating over a very localized region of the body. An analytical solution is presented to a heat transfer problem that approximates the shock wave interference heating of an engine cowl leading edge of the National Aero-Space Plane. The problem uses a simplified geometry to represent the leading edge. An analytical solution is developed that provides a means for approximating maximum temperature differences between the outer and inner surface temperatures of the leading edge. The solution is computationally efficient and, as a result, is well suited for conceptual and preliminary design or trade studies. Transient and steady state analyses are conducted, and results obtained from the analytical solution are compared with results of 2-D thermal finite element analyses over a wide range of design parameters. Isotropic materials as well as laminated composite materials are studied. Results of parametric studies are presented to indicate the effects of the thickness of the cowl leading edge and the width of the region heated by the shock wave interference on the thermal response of the leading edge.

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

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H.

    1950-01-01

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

  19. 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

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Preservation of wing leading edge suction at the plane of symmetry as a factor in wing-fuselage design

    NASA Technical Reports Server (NTRS)

    Larrabee, E. E.

    1975-01-01

    Most fuselage geometries cover a portion of the wing leading edge near the plane of symmetry, and it seems reasonable to expect that a large fraction of the leading edge suction which would be developed by the covered wing at high angles of attack is not developed on the fuselage. This is one of the reasons that the Oswald span efficiency factor for the wing body combination fails to approach the value predicted by lifting line theory for the isolated wing. Some traditional and recent literature on wing-body interference is discussed and high Reynolds number data on wing-body-nacelle drag are reviewed. An exposed central leading edge geometry has been developed for a sailplane configuration. Low Reynolds number tests have not validated the design concept.

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

    SciTech Connect

    Somers, D. M.

    2005-01-01

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

  6. Stroke plane angle controls leading edge vortex in a bat-inspired flapper

    NASA Astrophysics Data System (ADS)

    Koekkoek, Gide; Muijres, Florian T.; Johansson, L. Christoffer; Stuiver, Melanie; van Oudheusden, Bas W.; Hedenström, Anders

    2012-01-01

    The present interest in micro air vehicles has given the research on bat flight a new impulse. With the use of high speed cameras and improved PIV techniques, the kinematics and aerodynamics of bats have been studied in great detail. A robotic flapper makes it possible to do measurements by systematically changing only one parameter at a time and investigate the parameter space outside the natural flight envelope of bats without risking animal safety. For this study, a robotic flapper (RoBat), inspired by Leptonycteris yerbabuenae was developed and tested over the speed range 1-7 m/s, with variable maximum angles of attacks ( AoA=55° and 15°, respectively) and constant AoA=55°. These measurements show the presence of a leading edge vortex (LEV) for low speeds and a fully attached flow for high speeds at low AoA, which is in line with natural bat flight. A LEV occurs for AoA=55° throughout the complete flight speed range, and throughout which the LEV circulation coefficient remains rather constant. This implies that bats and micro air vehicles could use LEVs for high load maneuvers also at relatively high flight speeds. However, at high flight speeds the LEV bursts, which causes increased drag, most likely due to a decrease in Strouhal number.

  7. Prediction and Assessment of Reynolds Number Sensitivities Associated with Wing Leading-Edge Radius Variations

    NASA Technical Reports Server (NTRS)

    Wahls, Richard A.; Rivers, Melissa B.; Owens, Lewis R., Jr.

    1999-01-01

    The primary objectives of this study were to expand the data base showing the effects of LE radius distribution and corresponding sensitivity to Rn at subsonic and transonic conditions, and to assess the predictive capability of CFD for these effects. Several key elements led to the initiation of this project: 1) the necessity of meeting multipoint design requirements to enable a viable HSCT, 2) the demonstration that blunt supersonic leading-edges can be associated with performance gain at supersonic speeds , and 3) limited data. A test of a modified Reference H model with the TCA planform and 2 LE radius distributions was performed in the NTF, in addition to Navier-Stokes analysis for an additional 3 LE radius distributions. Results indicate that there is a tremendous potential to improve high-lift performance through the use of a blunt LE across the span given an integrated, fully optimized design, and that low Rn data alone is probably not sufficient to demonstrate the benefit.

  8. Prediction and Assessment of Reynolds Number Sensitivities Associated with Wing Leading-Edge Radius Variations

    NASA Technical Reports Server (NTRS)

    Wahls, Richard A.; Rivers, Melissa B.; Owen, Lewis R., Jr.

    1999-01-01

    The primary objectives of this study were to expand the data base showing the effects of LE radius distribution and corresponding . sensitivity to Rn at subsonic and transonic conditions, and to assess the predictive capability of CFD for these effects. Several key elements led to the initiation of this project: 1) the necessity of meeting multipoint design requirements to enable a viable HSCT, 2) the demonstration that blunt supersonic leading-edges can be associated with performance gain at supersonic speeds , and 3) limited data. A test of a modified Reference H model with the TCA planform and 2 LE radius distributions was performed in the NTF, in addition to Navier-Stokes analysis for an additional 3 LE radius distributions. Results indicate that there is a tremendous potential to improve high-lift performance through the use of a blunt LE across the span given an integrated, fully optimized design, and that low Rn data alone is probably not sufficient to demonstrate the benefit.

  9. Development and Validation of a Novel Bird Strike Resistant Composite Leading Edge Structure

    NASA Astrophysics Data System (ADS)

    Kermanidis, Th.; Labeas, G.; Sunaric, M.; Ubels, L.

    2005-11-01

    A novel design of a fibre-reinforced composite Leading Edge (LE) of a Horizontal Tail Plain (HTP) is proposed. The development and validation approach of the innovative composite LE structure are described. The main design goal is the satisfactory impact resistance of the novel composite LE in the case of bird strike. The design concept is based on the absorption of the major portion of the bird kinetic energy by the composite skins, in order to protect the ribs and the inner LE structure from damaging, thus preserving the tail plane functionality for safe landing. To this purpose, the LE skin is fabricated from specially designed composite panels, so called ‘tensor skin’ panels, comprising folded layers, which unfold under the impact load and increase the energy absorption capability of the LE. A numerical model simulating the bird strike process is developed and bird strike experimental testing is performed, in order to validate the proposed layout and prove the capability of the structure to successfully withstand the impact loading. The numerical modelling issues and the critical parameters of the simulation are discussed. The present work is part of the European Aeronautics Research Project, ‘Crashworthiness of aircraft for high velocity impact CRAHVI’ [1].

  10. Augmentation of fighter-aircraft performance by spanwise blowing over the wing leading edge

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Salomon, M.

    1983-01-01

    Spanwise blowing over the wing and canard of a 1:35 model of a close-coupled-canard fighter-airplane configuration (similar to the Kfir-C2) was investigated experimentally in low-speed flow. Tests were conducted at airspeeds of 30 m/sec (Reynolds number of 1.8 x 10 to the 5th power based on mean aerodynamic chord) with angle-of-attack sweeps from -8 deg to 60 deg, and yaw-angle sweeps from -8 deg to 36 deg at fixed angles of attack 0 deg, 10 deg, 20 deg, 25 deg, 30 deg, and 35 deg. Significant improvement in lift-curve slope, maximum lift, drag polar and lateral/directional stability was found, enlarging the flight envelope beyond its previous low-speed/maximum-lift limit. In spite of the highly swept (60 deg) leading edge, the efficiency of the lift augmentation by blowing was relatively high and was found to increase with increasing blowing momentum on the close-coupled-canard configuration. Interesting possibilities of obtaining much higher efficiencies with swirling jets were indicated.

  11. Thermostructural Evaluation of Joggle Region on the Shuttle Orbiter's Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Walker, Sandra P.; Warren, Jerry E.

    2012-01-01

    An investigation was initiated to determine the cause of coating spallation occurring on the Shuttle Orbiter's wing leading edge panels in the slip-side joggle region. The coating spallation events were observed, post flight, on differing panels on different missions. As part of the investigation, the high re-entry heating occurring on the joggles was considered here as a possible cause. Thus, a thermostructural evaluation was conducted to determine the detailed state-of-stress in the joggle region during re-entry and the feasibility of a laboratory test on a local joggle specimen to replicate this state-of-stress. A detailed three-dimensional finite element model of a panel slip-side joggle region was developed. Parametric and sensitivity studies revealed significant stresses occur in the joggle during peak heating. A critical interlaminar normal stress concentration was predicted in the substrate at the coating interface and was confined to the curved joggle region. Specifically, the high interlaminar normal stress is identified to be the cause for the occurrence of failure in the form of local subsurface material separation occurring in the slip-side joggle. The predicted critical stresses are coincident with material separations that had been observed with microscopy in joggle specimens obtained from flight panels.

  12. Wing Leading Edge RCC Rapid Response Damage Prediction Tool (IMPACT2)

    NASA Technical Reports Server (NTRS)

    Clark, Robert; Cottter, Paul; Michalopoulos, Constantine

    2013-01-01

    This rapid response computer program predicts Orbiter Wing Leading Edge (WLE) damage caused by ice or foam impact during a Space Shuttle launch (Program "IMPACT2"). The program was developed after the Columbia accident in order to assess quickly WLE damage due to ice, foam, or metal impact (if any) during a Shuttle launch. IMPACT2 simulates an impact event in a few minutes for foam impactors, and in seconds for ice and metal impactors. The damage criterion is derived from results obtained from one sophisticated commercial program, which requires hours to carry out simulations of the same impact events. The program was designed to run much faster than the commercial program with prediction of projectile threshold velocities within 10 to 15% of commercial-program values. The mathematical model involves coupling of Orbiter wing normal modes of vibration to nonlinear or linear springmass models. IMPACT2 solves nonlinear or linear impact problems using classical normal modes of vibration of a target, and nonlinear/ linear time-domain equations for the projectile. Impact loads and stresses developed in the target are computed as functions of time. This model is novel because of its speed of execution. A typical model of foam, or other projectile characterized by material nonlinearities, impacting an RCC panel is executed in minutes instead of hours needed by the commercial programs. Target damage due to impact can be assessed quickly, provided that target vibration modes and allowable stress are known.

  13. Fracture Mechanics Analyses of Reinforced Carbon-Carbon Wing-Leading-Edge Panels

    NASA Technical Reports Server (NTRS)

    Raju, Ivatury S.; Phillips, Dawn R.; Knight, Norman F., Jr.; Song, Kyongchan

    2010-01-01

    Fracture mechanics analyses of subsurface defects within the joggle regions of the Space Shuttle wing-leading-edge RCC panels are performed. A 2D plane strain idealized joggle finite element model is developed to study the fracture behavior of the panels for three distinct loading conditions - lift-off and ascent, on-orbit, and entry. For lift-off and ascent, an estimated bounding aerodynamic pressure load is used for the analyses, while for on-orbit and entry, thermo-mechanical analyses are performed using the extreme cold and hot temperatures experienced by the panels. In addition, a best estimate for the material stress-free temperature is used in the thermo-mechanical analyses. In the finite element models, the substrate and coating are modeled separately as two distinct materials. Subsurface defects are introduced at the coating-substrate interface and within the substrate. The objective of the fracture mechanics analyses is to evaluate the defect driving forces, which are characterized by the strain energy release rates, and determine if defects can become unstable for each of the loading conditions.

  14. Effect of Impact Location on the Response of Shuttle Wing Leading Edge Panel 9

    NASA Technical Reports Server (NTRS)

    Lyle, Karen H.; Spellman, Regina L.; Hardy, Robin C.; Fasanella, Edwin L.; Jackson, Karen E.

    2005-01-01

    The objective of this paper is to compare the results of several simulations performed to determine the worst-case location for a foam impact on the Space Shuttle wing leading edge. The simulations were performed using the commercial non-linear transient dynamic finite element code, LS-DYNA. These simulations represent the first in a series of parametric studies performed to support the selection of the worst-case impact scenario. Panel 9 was selected for this study to enable comparisons with previous simulations performed during the Columbia Accident Investigation. The projectile for this study is a 5.5-in cube of typical external tank foam weighing 0.23 lb. Seven locations spanning the panel surface were impacted with the foam cube. For each of these cases, the foam was traveling at 1000 ft/s directly aft, along the orbiter X-axis. Results compared from the parametric studies included strains, contact forces, and material energies for various simulations. The results show that the worst case impact location was on the top surface, near the apex.

  15. Pressure-Velocity Correlations in the Cove of a Leading Edge Slat

    NASA Astrophysics Data System (ADS)

    Wilkins, Stephen; Richard, Patrick; Hall, Joseph

    2015-11-01

    One of the major sources of aircraft airframe noise is related to the deployment of high-lift devices, such as leading-edge slats, particularly when the aircraft is preparing to land. As the engines are throttled back, the noise produced by the airframe itself is of great concern, as the aircraft is low enough for the noise to impact civilian populations. In order to reduce the aeroacoustic noise sources associated with these high lift devices for the next generation of aircraft an experimental investigation of the correlation between multi-point surface-mounted fluctuating pressures measured via flush-mounted microphones and the simultaneously measured two-component velocity field measured via Particle Image Velocimetry (PIV) is studied. The development of the resulting shear-layer within the slat cove is studied for Re =80,000, based on the wing chord. For low Mach number flows in air, the major acoustic source is a dipole acoustic source tied to fluctuating surface pressures on solid boundaries, such as the underside of the slat itself. Regions of high correlations between the pressure and velocity field near the surface will likely indicate a strong acoustic dipole source. In order to study the underlying physical mechanisms and understand their role in the development of aeroacoustic noise, Proper Orthogonal Decomposition (POD) by the method of snapshots is employed on the velocity field. The correlation between low-order reconstructions and the surface-pressure measurements are also studied.

  16. Numerical Predictions of Sonic Boom Signatures for a Straight Line Segmented Leading Edge Model

    NASA Technical Reports Server (NTRS)

    Elmiligui, Alaa A.; Wilcox, Floyd J.; Cliff, Susan; Thomas, Scott

    2012-01-01

    A sonic boom wind tunnel test was conducted on a straight-line segmented leading edge (SLSLE) model in the NASA Langley 4- by 4- Foot Unitary Plan Wind Tunnel (UPWT). The purpose of the test was to determine whether accurate sonic boom measurements could be obtained while continuously moving the SLSLE model past a conical pressure probe. Sonic boom signatures were also obtained using the conventional move-pause data acquisition method for comparison. The continuous data acquisition approach allows for accurate signatures approximately 15 times faster than a move-pause technique. These successful results provide an incentive for future testing with greatly increased efficiency using the continuous model translation technique with the single probe to measure sonic boom signatures. Two widely used NASA codes, USM3D (Navier-Stokes) and CART3D-AERO (Euler, adjoint-based adaptive mesh), were used to compute off-body sonic boom pressure signatures of the SLSLE model at several different altitudes below the model at Mach 2.0. The computed pressure signatures compared well with wind tunnel data. The effect of the different altitude for signature extraction was evaluated by extrapolating the near field signatures to the ground and comparing pressure signatures and sonic boom loudness levels.

  17. Control of Flow Structure on Low Swept Delta Wing with Steady Leading Edge Blowing

    NASA Astrophysics Data System (ADS)

    Ozturk, Ilhan; Zharfa, Mohammadreza; Yavuz, Mehmet Metin

    2014-11-01

    Interest in unmanned combat air vehicles (UCAVs) and micro air vehicles (MAVs) has stimulated investigation of the flow structure, as well as its control, on delta wings having low and moderate values of sweep angle. In the present study, the flow structure is characterized on a delta wing of low sweep 35-degree angle, which is subjected to steady leading edge blowing. The techniques of laser illuminated smoke visualization, laser Doppler anemometry (LDA), and surface pressure measurements are employed to investigate the steady and unsteady nature of the flow structure on delta wing, in relation to the dimensionless magnitude of the blowing coefficient. Using statistics and spectral analysis, unsteadiness of the flow structure is studied in detail. Different injection locations are utilized to apply different blowing patterns in order to identify the most efficient control, which provides the upmost change in the flow structure with the minimum energy input. The study aims to find the optimum flow control strategy to delay or to prevent the stall and possibly to reduce the buffeting on the wing surface. Since the blowing set-up is computer controlled, the unsteady blowing patterns compared to the present steady blowing patterns will be studied next. This project was supported by the Scientific and Technological Research Council of Turkey (Project Number: 3501 111M732).

  18. Augmentation of Fighter-Aircraft Performance by Spanwise Blowing over the Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Seginer, A.; Salomon, M.

    1983-01-01

    Spanwise blowing over the wing and canard of a 1:35 model of a close-coupled-canard fighter airplane configuration (similar to the Kfir-C2) was investigated experimentally in low-speed flow. Tests were conducted at airspeeds of 30 m/sec (Reynolds number of 1.8 x 10 to the 5th power based on mean aerodynamic chord) with angle-of-attack sweeps from -8 to 60 deg, and yaw-angle sweeps from -8 to 36 deg at fixed angles of attack 0, 10, 20, 25, 30, and 35 deg. Significant improvement in lift-curve slope, maximum lift, drag polar and lateral/directional stability was found, enlarging the flight envelope beyond its previous low-speed/maximum-lift limit. In spite of the highly swept (60 deg) leading edge, the efficiency of the lift augmentation by blowing was relatively high and was found to increase with increasing blowing momentum on the close-coupled-canard configuration. Interesting possibilities of obtaining much higher efficiencies with swirling jets were indicated.

  19. Development of Detectability Limits for On-Orbit Inspection of Space Shuttle Wing Leading Edge

    NASA Technical Reports Server (NTRS)

    Stephan, Ryan A.; Johnson, David G.; Mastropietro, A. J.; Ancarrow, Walt C.

    2005-01-01

    At the conclusion of the Columbia Accident Investigation, one of the recommendations of the Columbia Accident Investigation Board (CAIB) was that NASA develop and implement an inspection plan for the Reinforced Carbon-Carbon (RCC) system components of the Space Shuttle. To address these issues, a group of scientists and engineers at NASA Langley Research Center proposed the use of an IR camera to inspect the RCC. Any crack in an RCC panel changes the thermal resistance of the material in the direction perpendicular to the crack. The change in thermal resistance can be made visible by introducing a heat flow across the crack and using an IR camera to image the resulting surface temperature distribution. The temperature difference across the crack depends on the change in the thermal resistance, the length of the crack, the local thermal gradient, and the rate of radiation exchange with the environment. This paper describes how the authors derived the minimum thermal gradient detectability limits for a through crack in an RCC panel. This paper will also show, through the use of a transient, 3-dimensional, finite element model, that these minimum gradients naturally exist on-orbit. The results from the finite element model confirm that there are sufficient thermal gradient to detect a crack on 96% of the RCC leading edge.

  20. Performance of laminar-flow leading-edge test articles in cloud encounters

    NASA Technical Reports Server (NTRS)

    Davis, Richard E.; Maddalon, Dal V.; Wagner, Richard D.

    1987-01-01

    An extensive data bank of concurrent measurements of laminar flow (LF), particle concentration, and aircraft charging state was gathered for the first time. From this data bank, 13 flights in the simulated airline service (SAS) portion were analyzed to date. A total of 6.86 hours of data at one-second resolution were analyzed. An extensive statistical analysis, for both leading-edge test articles, shows that there is a significant effect of cloud and haze particles on the extent of laminar flow obtained. Approximately 93 percent of data points simulating LFC flight were obtained in clear air conditions; approximately 7 percent were obtained in cloud and haze. These percentages are consistent with earlier USAF and NASA estimates and results. The Hall laminar flow loss criteria was verified qualitatively. Larger particles and higher particle concentrations have a more marked effect on LF than do small particles. A particle spectrometer of a charging patch are both acceptable as diagnostic indicators of the presence of particles detrimental to laminar flow.

  1. Turbulent Wing-Leading-Edge Correlation Assessment for the Shuttle Orbiter

    NASA Technical Reports Server (NTRS)

    King, Rudolph A.; Vaughan, Matthew P.

    2009-01-01

    This study was conducted in support of the Orbiter damage assessment activity that takes place for each Shuttle mission since STS-107 (STS - Space Transportation System). As part of the damage assessment activity, the state of boundary layer (laminar or turbulent) during reentry needs to be estimated in order to define the aerothermal environment on the Orbiter. Premature turbulence on the wing leading edge (WLE) is possible if a surface irregularity promotes early transition and the resulting turbulent wedge flow contaminates the WLE flow. The objective of this analysis is to develop a criterion to determine if and when the flow along the WLE experiences turbulent heating given an incoming turbulent boundary layer that contaminates the attachment line. The data to be analyzed were all obtained as part of the MH-13 Space Shuttle Orbiter Aerothermodynamic Test conducted on a 1.8%-scale Orbiter model at Calspan/University of Buffalo Research Center in the Large Energy National Shock Tunnels facility. A rational framework was used to develop a means to assess the state of the WLE flow on the Orbiter during reentry given a contaminated attachment-line flow. Evidence of turbulent flow on the WLE has been recently documented for a few STS missions during the Orbiter s flight history, albeit late in the reentry trajectory. The criterion developed herein will be compared to these flight results.

  2. Tropomyosin Promotes Lamellipodial Persistence by Collaborating with Arp2/3 at the Leading Edge.

    PubMed

    Brayford, Simon; Bryce, Nicole S; Schevzov, Galina; Haynes, Elizabeth M; Bear, James E; Hardeman, Edna C; Gunning, Peter W

    2016-05-23

    At the leading edge of migrating cells, protrusion of the lamellipodium is driven by Arp2/3-mediated polymerization of actin filaments [1]. This dense, branched actin network is promoted and stabilized by cortactin [2, 3]. In order to drive filament turnover, Arp2/3 networks are remodeled by proteins such as GMF, which blocks the actin-Arp2/3 interaction [4, 5], and coronin 1B, which acts by directing SSH1L to the lamellipodium where it activates the actin-severing protein cofilin [6, 7]. It has been shown in vitro that cofilin-mediated severing of Arp2/3 actin networks results in the generation of new pointed ends to which the actin-stabilizing protein tropomyosin (Tpm) can bind [8]. The presence of Tpm in lamellipodia, however, is disputed in the literature [9-19]. Here, we report that the Tpm isoforms 1.8/9 are enriched in the lamellipodium of fibroblasts as detected with a novel isoform-specific monoclonal antibody. RNAi-mediated silencing of Tpm1.8/9 led to an increase of Arp2/3 accumulation at the cell periphery and a decrease in the persistence of lamellipodia and cell motility, a phenotype consistent with cortactin- and coronin 1B-deficient cells [2, 7]. In the absence of coronin 1B or cofilin, Tpm1.8/9 protein levels are reduced while, conversely, inhibition of Arp2/3 with CK666 leads to an increase in Tpm1.8/9 protein. These findings establish a novel regulatory mechanism within the lamellipodium whereby Tpm collaborates with Arp2/3 to promote lamellipodial-based cell migration. PMID:27112294

  3. Method and System for Weakening Shock Wave Strength at Leading Edge Surfaces of Vehicle in Supersonic Atmospheric Flight

    NASA Technical Reports Server (NTRS)

    Daso, Endwell O. (Inventor); Pritchett, Victor E., II (Inventor); Wang, Ten-See (Inventor); Farr, Rebecca Ann (Inventor); Auslender, Aaron Howard (Inventor); Blankson, Isaiah M. (Inventor); Plotkin, Kenneth J. (Inventor)

    2015-01-01

    A method and system are provided to weaken shock wave strength at leading edge surfaces of a vehicle in atmospheric flight. One or more flight-related attribute sensed along a vehicle's outer mold line are used to control the injection of a non-heated, non-plasma-producing gas into a local external flowfield of the vehicle from at least one leading-edge surface location along the vehicle's outer mold line. Pressure and/or mass flow rate of the gas so-injected is adjusted in order to cause a Rankine-Hugoniot Jump Condition along the vehicle's outer mold line to be violated.

  4. An overview of the fundamental aerodynamics branch's research activities in wing leading-edge vortex flows at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Miller, D. S.; Wood, R. M.; Covell, P. F.

    1986-01-01

    For the past 3 years, a research program pertaining to the study of wing leading edge vortices at supersonic speeds has been conducted in the Fundamental Aerodynamics Branch of the High-Speed Aerodynamics Division at the Langley Research Center. The purpose of the research is to provide an understanding of the factors governing the formation and the control of wing leading-edge vortices and to evaluate the use of these vortices for improving supersonic aerodynamic performance. The studies include both experimental and theoretical investigations and focus primarily on planform, thickness and camber effects for delta wings. An overview of this research activity is presented.

  5. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1985-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Chen, Fang-Jenq; Beeler, George B.

    2002-01-01

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

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. Subsonic Investigation of a Leading-Edge Boundary Layer Control Suction System on a High-Speed Civil Transport Configuration

    NASA Technical Reports Server (NTRS)

    Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.; Coe, Paul L., Jr.; Owens, D. Bruce; Gile, Brenda E.; Parikh, Pradip G.; Smith, Don

    1999-01-01

    A wind tunnel investigation of a leading edge boundary layer control system was conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.08 to 0.27, with corresponding chord Reynolds numbers of 1.79 x 10(exp 6) to 5.76 x 10(exp 6). Variations in the amount of suction, as well as the size and location of the suction area, were tested with outboard leading edge flaps deflected 0 and 30 deg and trailing-edge flaps deflected 0 and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein.

  14. 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.

  15. 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.

  16. Data acquisition electronics for gamma ray emission tomography using width-modulated leading-edge discriminators

    NASA Astrophysics Data System (ADS)

    Lage, E.; Tapias, G.; Villena, J.; Desco, M.; Vaquero, J. J.

    2010-08-01

    We present a new high-performance and low-cost approach for implementing radiation detection acquisition systems. The basic elements used are charge-integrating ADCs and a set of components encapsulated in an HDL (hardware definition language) library which makes it possible to implement several acquisition tasks such as time pickoff and coincidence detection using a new and simple trigger technique that we name WMLET (width-modulated leading-edge timing). As proof of concept, a 32-channel hybrid PET/SPECT acquisition system based on these elements was developed and tested. This demonstrator consists of a master module responsible for the generation and distribution of trigger signals, 2 × 16-channel ADC cards (12-bit resolution) for data digitization and a 32-bit digital I/O PCI card for handling data transmission to a personal computer. System characteristics such as linearity, maximum transmission rates or timing resolution in coincidence mode were evaluated with test and real detector signals. Imaging capabilities of the prototype were also evaluated using different detector configurations. The performance tests showed that this implementation is able to handle data rates in excess of 600k events s-1 when acquiring simultaneously 32 channels (96-byte events). ADC channel linearity is >98.5% in energy quantification. Time resolution in PET mode for the tested configurations ranges from 3.64 ns FWHM to 7.88 ns FWHM when signals from LYSO-based detectors are used. The measured energy resolution matched the expected values for the detectors evaluated and single elements of crystal matrices can be neatly separated in the acquired flood histograms.

  17. Data acquisition electronics for gamma ray emission tomography using width-modulated leading-edge discriminators.

    PubMed

    Lage, E; Tapias, G; Villena, J; Desco, M; Vaquero, J J

    2010-08-01

    We present a new high-performance and low-cost approach for implementing radiation detection acquisition systems. The basic elements used are charge-integrating ADCs and a set of components encapsulated in an HDL (hardware definition language) library which makes it possible to implement several acquisition tasks such as time pickoff and coincidence detection using a new and simple trigger technique that we name WMLET (width-modulated leading-edge timing). As proof of concept, a 32-channel hybrid PET/SPECT acquisition system based on these elements was developed and tested. This demonstrator consists of a master module responsible for the generation and distribution of trigger signals, 2 x 16-channel ADC cards (12-bit resolution) for data digitization and a 32-bit digital I/O PCI card for handling data transmission to a personal computer. System characteristics such as linearity, maximum transmission rates or timing resolution in coincidence mode were evaluated with test and real detector signals. Imaging capabilities of the prototype were also evaluated using different detector configurations. The performance tests showed that this implementation is able to handle data rates in excess of 600k events s(-1) when acquiring simultaneously 32 channels (96-byte events). ADC channel linearity is >98.5% in energy quantification. Time resolution in PET mode for the tested configurations ranges from 3.64 ns FWHM to 7.88 ns FWHM when signals from LYSO-based detectors are used. The measured energy resolution matched the expected values for the detectors evaluated and single elements of crystal matrices can be neatly separated in the acquired flood histograms.

  18. Space environmental effects on LDEF composites: A leading edge coated graphite epoxy panel

    NASA Technical Reports Server (NTRS)

    George, Pete E.; Dursch, Harry W.; Hill, Sylvester G.

    1993-01-01

    The electronics module cover for the leading edge (Row D 9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a (O(sub 2), +/- 45, O(sub 2), +/- 45, 90, 0)(sub s) layup. This 11.75 in x 16.75 in panel was covered with thermal control coatings in three of the four quadrants with the fourth quadrant uncoated. The composite panel experienced different thermal cycling extremes in each quadrant due to the different optical properties of the coatings and bare composite. The panel also experienced ultraviolet (UV) and atomic oxygen (AO) attack as well as micrometeoroid and space debris impacts. An AO reactivity of 0.99 x 10(exp -24) cm(sup 3)/atom was calculated for the bare composite based on thickness loss. The white urethane thermal control coatings (A276 and BMS 1060) prevented AO attack of the composite substrate. However, the black urethane thermal control coating (Z306) was severely eroded by AO, allowing some AO attack of the composite substrate. An interesting banding pattern on the AO eroded bare composite surface was investigated and found to match the dimensions of the graphite fiber tow widths as prepregged. Also, erosion depths were greater in the darker bands. Five micrometeoroid/space debris impacts were cross sectioned to investigate possible structural damage as well as impact/AO interactions. Local crushing and delaminations were found to some extent in all of the impacts. No signs of coating undercutting were observed despite the extensive AO erosion patterns seen in the exposed composite material at the impact sites. An extensive microcrack study was performed on the panel along with modeling of the thermal environment to estimate temperature extremes and thermal shock. The white coated composite substrate displayed almost no microcracking while the black coated and bare composite showed extensive microcracking. Significant AO erosion was seen in many of the cracks in the bare composite.

  19. Heat transfer and material temperature conditions in the leading edge area of impingement-cooled turbine vanes

    NASA Astrophysics Data System (ADS)

    Berg, H. P.; Pfaff, K.; Hennecke, D. K.

    The resultant effects on the cooling effectiveness at the leading edge area of an impingement-cooled turbine vane by varying certain geometrical parameters is described with reference to local internal heat transfer coefficients determined from experiment and temperature calculations. The local heat transfer on the cooling-air side is determined experimentally with the aid of the analogy between heat- and mass transfer. The impingement cooling is provided from an inserted sheet-metal containing a single row of holes. The Reynolds Number and several of the cooling geometry parameters were varied. The results demonstrate the high local resolution of the method of measurement, which allows improved analytical treatment of the leading-edge cooling configuration. These experiments also point to the necessity of not always performing model tests under idealized conditions. This becomes very clear in the case of the tests performed on an application-oriented impingement-cooling configuration like that often encountered in engine manufacture. In conclusion, as an example, temperature calculations are employed to demonstrate the effect on the cooling effectiveness of varying the distances between insert and inner surface of the leading edge. It shows how the effectiveness of the leading edge cooling can be increased by simple geometrical measures, which results in a considerable improvement in service life.

  20. Turbulent Vortex-Flow Simulation Over a 65 deg Sharp and Blunt Leading-Edge Delta Wing at Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Ghaffari, Farhad

    2005-01-01

    Turbulent thin-layer, Reynolds-Averaged Navier-Stokes solutions, based on a multi-block structured grid, are presented for a 65 deg delta wing having either a sharp leading edge (SLE) or blunt leading edge (BLE) geometry. The primary objective of the study is to assess the prediction capability of the method for simulating the leading-edge flow separation and the ensuing vortex flow characteristics. Computational results are obtained for two angles of attack of approximately 13 and 20 deg, at free-stream Mach number of 0.40 and Reynolds number of 6 million based on the wing mean aerodynamic chord. The effects of two turbulence models of Baldwin-Lomax with Degani-Schiff (BL/DS) and the Spalart-Allmaras (SA) on the numerical results are also discussed. The computations also explore the effects of two numerical flux-splitting schemes, i.e., flux difference splitting (fds) and flux vector splitting (fvs), on the solution development and convergence characteristics. The resulting trends in solution sensitivity to grid resolution for the selected leading-edge geometries, angles of attack, turbulence models and flux splitting schemes are also presented. The validity of the numerical results is evaluated against a unique set of experimental wind-tunnel data that was obtained in the National Transonic Facility at the NASA Langley Research Center.

  1. Grainyhead-like factor Get1/Grhl3 regulates formation of the epidermal leading edge during eyelid closure.

    PubMed

    Yu, Zhengquan; Bhandari, Ambica; Mannik, Jaana; Pham, Thu; Xu, Xiaoman; Andersen, Bogi

    2008-07-01

    Grainyhead transcription factors play an evolutionarily conserved role in regulating epidermal terminal differentiation. One such factor, the mammalian Grainyhead-like epithelial transactivator (Get1/Grhl3), is important for epidermal barrier formation. In addition to a role in barrier formation, Grainyhead genes play roles in closure of several structures such as the mouse neural tube and Drosophila wounds. Consistent with these observations, we found that Get1 knockout mice have an eye-open at birth phenotype. The failure of eyelid closure appears to be due to critical functions of Get1 in promoting F-actin polymerization, filopodia formation, and the cell shape changes that are required for migration of the keratinocytes at the leading edge during eyelid closure. The expression of TGFalpha, a known regulator of leading edge formation, is decreased in the eyelid tip of Get1(-/-) mice. Levels of phospho-EGFR and phospho-ERK are also decreased at the leading edge tip. Furthermore, in an organ culture model, TGFalpha can increase levels of phospho-EGFR and promote cell shape changes as well as leading edge formation in Get1(-/-) eyelids, indicating that in eyelid closure Get1 acts upstream of TGFalpha in the EGFR/ERK pathway.

  2. "Partners in Science": A Model Cooperative Program Introducing High School Teachers and Students to Leading-Edge Pharmaceutical Science

    ERIC Educational Resources Information Center

    Woska, Joseph R., Jr.; Collins, Danielle M.; Canney, Brian J.; Arcario, Erin L.; Reilly, Patricia L.

    2005-01-01

    "Partners in Science" is a cooperative program between Boehringer Ingelheim Pharmaceuticals, Inc. and area high schools in the community surrounding our Connecticut campus. It is a two-phase program that introduces high school students and teachers to the world of drug discovery and leading-edge pharmaceutical research. Phase 1 involves a series…

  3. 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.

  4. Effects of leading-edge devices on the low-speed aerodynamic characteristics of a highly-swept arrow-wing

    NASA Technical Reports Server (NTRS)

    Scott, S. J.; Nicks, O. W.; Imbrie, P. K.

    1985-01-01

    An investigation was conducted in the Texas A&M University 7 by 10 foot Low Speed Wind Tunnel to provide a direct comparison of the effect of several leading edge devices on the aerodynamic performance of a highly swept wing configuration. Analysis of the data indicates that for the configuration with undeflected leading edges, vortex separation first occurs on the outboard wing panel for angles of attack of approximately 2, and wing apex vorticies become apparent for alpha or = 4 deg. However, the occurrence of the leading edge vortex flow may be postponed with leading edge devices. Of the devices considered, the most promising were a simple leading edge deflection of 30 deg and a leading edge slat system. The trailing edge flap effectiveness was found to be essentially the same for the configuration employing either of these more promising leading edge devices. Analysis of the lateral directional data showed that for all of the concepts considered, deflecting leading edge downward in an attempt to postpone leading edge vortex flows, has the favorable effect of reducing the effective dihedral.

  5. 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.

  6. The Influence of Mesh Density on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

    NASA Technical Reports Server (NTRS)

    Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.; Spellman, Regina L.

    2004-01-01

    A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0- in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA version 960 for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at five discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density. As a final comparison, the model with a fine panel and fine foam mesh was executed with slightly different material properties for the RCC. For this model, the average degraded properties of the RCC were replaced with the maximum degraded properties. Similar comparisons of panel and foam responses were made for the average and maximum degraded models.

  7. The International Polar Year (IPY) Circumpolar Flaw Lead (CFL) system study: a focus on fast ice edge systems

    NASA Astrophysics Data System (ADS)

    Barber, D. G.; Mundy, C. J.; Papakyriakou, T. N.; MacDonald, R. W.; Gratton, Y.; Fortier, L.; Gosselin, M.; Hanesiak, J.; Tremblay, J.; Ferguson, S.; Stern, G.; Meakin, S.; Deming, J. W.; Leitch, D.

    2009-12-01

    The Circumpolar Flaw Lead (CFL) system study supported a large multidisciplinary overwintering in the Banks Island (NT) flaw lead over the period September 2007 to August 2008 as part of the International Polar Year (IPY). The CFL system is formed when the central pack ice (which is mobile) moves away from coastal fast ice, opening a flaw lead. The CFL forms in the fall and continues as thin ice or open water throughout the winter. The flaw lead is circumpolar, with recurrent and interconnected polynyas occurring throughout the Arctic. The overarching objectives of the CFL project were to contrast the physical and biological systems of the flaw lead open water and thin ice to the adjacent landfast ice cover. The Canadian Research Icebreaker (NGCC Amundsen) completed the first-ever overwintering of a research icebreaker in the flaw lead. She supported a total of 11,000 person days distributed across 295 investigators from 28 different countries. The project obtained many first-ever measurements of a complete suite of physical, biogeochemical, contaminant and marine ecosystem variables across the open water - fast ice contrast. In this paper we present an overview of the early results from this study with a particular focus on the role of fast ice edges in modifying the physical, biological and geochemical processes that occur at the fast ice edge interface. Traditional knowledge of these ice edges illustrates a rich physical and ecological environment. Western science results show the temporal nature of fast ice edges and the associated stimulation of biological productivity make these a particularly important part of the flaw lead system.

  8. In-flight measurement of ice growth on an airfoil using an array of ultrasonic transducers

    NASA Technical Reports Server (NTRS)

    Hansman, R. John, Jr.; Kirby, Mark S.; Mcknight, Robert C.; Humes, Robert L.

    1987-01-01

    Results from three research flights to obtain in-flight ultrasonic pulse-echo measurements of airfoil ice thickness as a function of time using an array of eight ultrasonic transducers mounted flush with the leading edge of the airfoil are presented. The accuracy of the thickness measurements is found to be within 0.5 mm of mechanical and stereophotograph measurements of the ice accretion. The ultrasonic measurements demonstrate that the ice growth rate typically varies during the flight, with variations in the ice growth rate for dry ice growth being primarily due to fluctuations in the cloud liquid water content. Discrepancies between experimental results and results predicted by an analytic icing code underline the need for a better understanding of the physics of wet ice growth.

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

    SciTech Connect

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

    1995-12-01

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

  10. Subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration having spanwise leading-edge vortex enhancement

    NASA Technical Reports Server (NTRS)

    Huffman, J. K.; Fox, C. H., Jr.

    1977-01-01

    A configuration which integrates a close coupled canard wing combination, spanwise blowing for enhancement of the wing leading edge vortex, an engine-over-wing concept, and a wing trailing edge coanda-effect flap is studied. The data on the configuration are presented in tabular from without discussion. The investigation was conducted in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.166 through an angle-of-attack range from -2 to 22 deg. Rectangular main engine nozzles of aspect ratio 4, 6, and 8 were tested over a momentum coefficient range from 1.0 to 1.8.

  11. Measurement of Fuel Concentration Profile at Leading Edge of Lifted Flame with Acetone Laser-Induced Fluorescence

    NASA Astrophysics Data System (ADS)

    Hirota, Mitsutomo; Sekine, Kazushi; Hashimoto, Kouta; Saiki, Atsushi; Takahashi, Hidemi; Masuya, Goro

    This is a study of the leading-edge characteristics of a methane-air triple flame. Few experiment results are available for physical examination of such characteristics, so further experimental investigations are strongly needed to understand the stability mechanism in a mixture with a steep concentration gradient. To this end, we measured concentration profiles at the leading edge of a flame using acetone laser-induced fluorescence (acetone LIF). The results demonstrated that the lifted height of the flame changed when acetone was added to the mixture and correlated well with increased C2 radical behind the flame edge. However, the OH radical luminous intensity, measured with a spectroscope, did not change with addition of acetone. Moreover, the burning velocity obtained by the Bunsen-burner method remained constant when acetone was added to the mixture. Therefore, acetone had little influence on burning intensity. Acetone LIF can thus be employed to measure the local concentration gradient at the leading edge of a flame. The acetone LIF signals could be corrected to consider the thermal effect by using silicone oil vanishing-plane data. From the corrected acetone LIF data, the width between the lean and rich flammability limits (flammability limit width) in the flow upstream of the flame with a steep concentration gradient was clearly observed and could be quantitatively compared with the recent numerical results.

  12. 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.

  13. 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.

  14. The aerodynamics of Manduca sexta: digital particle image velocimetry analysis of the leading-edge vortex.

    PubMed

    Bomphrey, Richard J; Lawson, Nicholas J; Harding, Nicholas J; Taylor, Graham K; Thomas, Adrian L R

    2005-03-01

    Here we present the first digital particle image velocimetry (DPIV) analysis of the flow field around the wings of an insect (the tobacco hawkmoth Manduca sexta, tethered to a 6-component force-moment balance in a wind tunnel). A leading-edge vortex (LEV) is present above the wings towards the end of the downstroke, as the net upward force peaks. Our DPIV analyses and smoke visualisations match the results of previous flow visualisation experiments at midwing, and we extend the experiments to provide the first analysis of the flow field above the thorax. Detailed DPIV measurements show that towards the end of the downstroke, the LEV structure is consistent with that recently reported in free-flying butterflies and dragonflies: the LEV is continuous across the thorax and runs along each wing to the wingtip, where it inflects to form the wingtip trailing vortices. The LEV core is 2-3 mm in diameter (approximately 10% of local wing chord) both at the midwing position and over the centreline at 1.2 m s(-1) and at 3.5 m s(-1) flight speeds. At 1.2 m s(-1) the measured LEV circulation is 0.012+/-0.001 m(2) s(-1) (mean +/-S.D.) at the centreline and 0.011+/-0.001 m(2) s(-1) halfway along the wing. At 3.5 m s(-1) LEV circulation is 0.011+/-0.001 m(2) s(-1) at the centreline and 0.020+/-0.004 m(2) s(-1) at midwing. The DPIV measurements suggest that if there is any spanwise flow in the LEV towards the end of the downstroke its velocity is less than 1 m s(-1). Estimates of force production show that the LEV contributes significantly to supporting body weight during bouts of flight at both speeds (more than 10% of body weight at 1.2 m s(-1) and 35-65% of body weight at 3.5 m s(-1)).

  15. Effect of spanwise blowing on leading-edge vortex bursting of a highly swept aspect ratio 1.18 delta wing

    NASA Technical Reports Server (NTRS)

    Scantling, W. L.; Gloss, B. B.

    1974-01-01

    An investigation was conducted in the Langley 1/8-scale V/STOL model tunnel on a semispan delta wing with a leading-edge sweep of 74 deg, to determine the effectiveness of various locations of upper surface and reflection plane blowing on leading-edge vortex bursting. Constant area nozzles were located on the wing upper surface along a ray swept 79 deg, which was beneath the leading-edge vortex core. The bursting and reformation of the leading-edge vortex was viewed by injecting helium into the vortex core, and employing a schlieren system.

  16. Exploratory study of the effects of wing-leading-edge modifications on the stall/spin behavior of a light general aviation airplane

    NASA Technical Reports Server (NTRS)

    1979-01-01

    Configurations with full-span and segmented leading-edge flaps and full-span and segmented leading-edge droop were tested. Studies were conducted with wind-tunnel models, with an outdoor radio-controlled model, and with a full-scale airplane. Results show that wing-leading-edge modifications can produce large effects on stall/spin characteristics, particularly on spin resistance. One outboard wing-leading-edge modification tested significantly improved lateral stability at stall, spin resistance, and developed spin characteristics.

  17. Subscale, hydrogen-burning, airframe-integrated-scramjet: Experimental and theoretical evaluation of a water cooled strut airframe-integrated-scramjet: Experimental leading edge

    NASA Technical Reports Server (NTRS)

    Pinckney, S. Z.; Guy, R. W.; Beach, H. L., Jr.; Rogers, R. C.

    1975-01-01

    A water-cooled leading-edge design for an engine/airframe integrated scramjet model strut leading edge was evaluated. The cooling design employs a copper cooling tube brazed just downstream of the leading edge of a wedge-shaped strut which is constructed of oxygen-free copper. The survival of the strut leading edge during a series of tests at stagnation point heating rates confirms the practicality of the cooling design. A finite difference thermal model of the strut was also proven valid by the reasonable agreement of calculated and measured values of surface temperature and cooling-water heat transfer.

  18. 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.

  19. The Influence of Clocking Angle of the Projectile on the Simulated Impact Response of a Shuttle Leading Edge Wing Panel

    NASA Technical Reports Server (NTRS)

    Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.; Spellman, Regina L.

    2005-01-01

    An analytical study was conducted to determine the influence of clocking angle of a foam projectile impacting a space shuttle leading edge wing panel. Four simulations were performed using LS-DYNA. The leading edge panels are fabricated of multiple layers of reinforced carbon-carbon (RCC) material. The RCC material was represented using Mat 58, which is a material property that can be used for laminated composite fabrics. Simulations were performed of a rectangular-shaped foam block, weighing 0.23-lb., impacting RCC Panel 9 on the top surface. The material properties of the foam were input using Mat 83. The impact velocity was 1,000 ft/s along the Orbiter X-axis. In two models, the foam impacted on a corner, in one model the foam impacted the panel initially on the 2-in.-long edge, and in the last model the foam impacted the panel on the 7-in.- long edge. The simulation results are presented as contour plots of first principal infinitesimal strain and time history plots of contact force and internal and kinetic energy of the foam and RCC panel.

  20. Thelma and Louise Do Religious Education: A Dialogue from the Edge for Leading with Hope

    ERIC Educational Resources Information Center

    Meyers, Patty; Willhauck, Susan

    2003-01-01

    The 1991 movie Thelma and Louise and its protagonists continue to be cultural icons for many women of all ages. With quotations, song lyrics, the metaphor of the edge from the film, and collected wisdom from pedagogy, two religious educators reflect on their vocations and leadership drawing implications for the teaching ministry. The themes…

  1. 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.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  3. Flow visualization of the effect of pitch amplitude changes on the vortical signatures behind a three-dimensional flapping airfoil

    NASA Astrophysics Data System (ADS)

    Parker, Kamalluddien; von Ellenrieder, K. D.; Soria, J.

    2003-04-01

    The structure of the vortical flow behind a symmetrical airfoil of finite aspect ratio undergoing combinations of heave and pitch motions is investigated using qualitative dye flow visualization. The results are contrasted with flow visualizations obtained using electrolytic precipitation. The effect of changing the pitch amplitude is observed from the plan from view and wingtip view of the airfoil. With a Strouhal number of 0.35, Reyholds number based on airfoil chord of 164 and a phase angle of 90o, the maximum pitch amplitude is varied from 0° to 20°. The geometry of the downstream vortical flow is observed to change suggesting that the induced velocity from interacting structures decreases at lower pitch amplitudes. The rate of dynamic stall development may also be affected by variations in pitch amplitude since it appears that the timing of leading edge separation is affected. The flow field of an airfoil flapping periodically about a fixed axis appears to be influenced by the amplitude of pitching oscillations. At the tested Strouhal numbers the vortex formations appear to be primarily dependent on airfoil oscillation rather than heave translation. Furthermore, the results suggest that the wake structures originating from the dynamic stall process are important for the analysis of these complex flows. While the results from the two flow visualization techniques are similar, the dye flow visualization images provide greater qualitative insight. Inherently, precipitative techniques such as the one used here could provide good flow visualizations since the smoke/particles leave the surface of the airfoil, but the setup is found to be very sensitive to potential changes. The ion content in the electrolytic material was also found to play a role. Furthermore, the high ablation rate of the technique presented some practical problems.

  4. A Unit-Problem Investigation of Blunt Leading-Edge Separation Motivated by AVT-161 SACCON Research

    NASA Technical Reports Server (NTRS)

    Luckring, James M.; Boelens, Okko J.

    2011-01-01

    A research effort has been initiated to examine in more detail some of the challenging flow fields discovered from analysis of the SACCON configuration aerodynamics. This particular effort is oriented toward a diamond wing investigation specifically designed to isolate blunt leading-edge separation phenomena relevant to the SACCON investigations of the present workshop. The approach taken to design this new effort is reviewed along with the current status of the program.

  5. Materials, Manufacturing, and Test Development of a Composite Fan Blade Leading Edge Subcomponent for Improved Impact Resistance

    NASA Technical Reports Server (NTRS)

    Miller, Sandi G.; Handschuh, Katherine; Sinnott, Matthew J.; Kohlman, Lee W.; Roberts, Gary D.; Martin, Richard E.; Ruggeri, Charles R.; Pereira, J. Michael

    2015-01-01

    Application of polymer matrix composite materials for jet engine fan blades is becoming attractive as an alternative to metallic blades; particularly for large engines where significant weight savings are recognized on moving to a composite structure. However, the weight benefit of the composite is offset by a reduction of aerodynamic efficiency resulting from a necessary increase in blade thickness; relative to the titanium blades. Blade dimensions are largely driven by resistance to damage on bird strike. Further development of the composite material is necessary to allow composite blade designs to approximate the dimensions of a metallic fan blade. The reduction in thickness over the state of the art composite blades is expected to translate into structural weight reduction, improved aerodynamic efficiency, and therefore reduced fuel consumption. This paper presents test article design, subcomponent blade leading edge fabrication, test method development, and initial results from ballistic impact of a gelatin projectile on the leading edge of composite fan blades. The simplified test article geometry was developed to realistically simulate a blade leading edge while decreasing fabrication complexity. Impact data is presented on baseline composite blades and toughened blades; where a considerable improvement to impact resistance was recorded.

  6. Model for adhesion clutch explains biphasic relationship between actin flow and traction at the cell leading edge

    PubMed Central

    Craig, Erin M.; Stricker, Jonathan; Gardel, Margaret L.; Mogilner, Alex

    2015-01-01

    Cell motility relies on the continuous reorganization of a dynamic actin-myosin-adhesion network at the leading edge of the cell, in order to generate protrusion at the leading edge and traction between the cell and its external environment. We analyze experimentally measured spatial distributions of actin flow, traction force, myosin density, and adhesion density in control and pharmacologically perturbed epithelial cells in order to develop a mechanical model of the actin-adhesion-myosin self-organization at the leading edge. A model in which the F-actin network is treated as a viscous gel, and adhesion clutch engagement is strengthened by myosin but weakened by actin flow, can explain the measured molecular distributions and correctly predict the spatial distributions of the actin flow and traction stress. We test the model by comparing its predictions with measurements of the actin flow and traction stress in cells with fast and slow actin polymerization rates. The model predicts how the location of the lamellipodium-lamellum boundary depends on the actin viscosity and adhesion strength. The model further predicts that the location of the lamellipodium-lamellum boundary is not very sensitive to the level of myosin contraction. PMID:25969948

  7. Materials, Manufacturing and Test Development of a Composite Fan Blade Leading Edge Subcomponent for Improved Impact Resistance

    NASA Technical Reports Server (NTRS)

    Handschuh, Katherine M.; Miller, Sandi G.; Sinnott, Matthew J.; Kohlman, Lee W.; Roberts, Gary D.; Pereira, J. Michael; Ruggeri, Charles R.

    2014-01-01

    Application of polymer matrix composite materials for jet engine fan blades is becoming attractive as an alternative to metallic blades; particularly for large engines where significant weight savings are recognized on moving to a composite structure. However, the weight benefit of the composite of is offset by a reduction of aerodynamic efficiency resulting from a necessary increase in blade thickness; relative to the titanium blades. Blade dimensions are largely driven by resistance to damage on bird strike. Further development of the composite material is necessary to allow composite blade designs to approximate the dimensions of a metallic fan blade. The reduction in thickness over the state of the art composite blades is expected to translate into structural weight reduction, improved aerodynamic efficiency, and therefore reduced fuel consumption. This paper presents test article design, subcomponent blade leading edge fabrication, test method development, and initial results from ballistic impact of a gelatin projectile on the leading edge of composite fan blades. The simplified test article geometry was developed to realistically simulate a blade leading edge while decreasing fabrication complexity. Impact data is presented on baseline composite blades and toughened blades; where a considerable improvement to impact resistance was recorded.

  8. Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

    NASA Technical Reports Server (NTRS)

    Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.

    1992-01-01

    Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

  9. 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.

  10. Flight tests of a supersonic natural laminar flow airfoil

    NASA Astrophysics Data System (ADS)

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

    2015-06-01

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

  11. Flight Tests of a Supersonic Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  12. Flight Tests of a Supersonic Natural Laminar Flow Airfoil

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  13. 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.

  14. An approach to constrained aerodynamic design with application to airfoils

    NASA Technical Reports Server (NTRS)

    Campbell, Richard L.

    1992-01-01

    An approach was developed for incorporating flow and geometric constraints into the Direct Iterative Surface Curvature (DISC) design method. In this approach, an initial target pressure distribution is developed using a set of control points. The chordwise locations and pressure levels of these points are initially estimated either from empirical relationships and observed characteristics of pressure distributions for a given class of airfoils or by fitting the points to an existing pressure distribution. These values are then automatically adjusted during the design process to satisfy the flow and geometric constraints. The flow constraints currently available are lift, wave drag, pitching moment, pressure gradient, and local pressure levels. The geometric constraint options include maximum thickness, local thickness, leading-edge radius, and a 'glove' constraint involving inner and outer bounding surfaces. This design method was also extended to include the successive constraint release (SCR) approach to constrained minimization.

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

    NASA Technical Reports Server (NTRS)

    Ratnayake, Nalin A.

    2009-01-01

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

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

    PubMed

    Johnston, Joe; Gopalarathnam, Ashok

    2012-09-01

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

  17. Low-Speed Aerodynamic Data for an 0.18-Scale Model of an F-16XL with Various Leading-Edge Modifications

    NASA Technical Reports Server (NTRS)

    Hahne, Daniel E.

    1999-01-01

    Using the F-16XL as a test-bed, two strategies for improving the low-speed flying characteristics that had minimal impact on high-speed performance were evaluated. In addition to the basic F-16XL configuration several modifications to the baseline configuration were tested in the Langley 30- X 60-Foot Tunnel: 1) the notched area at the wing leading edge and fuselage juncture was removed resulting in a continuous 70 deg leading-edge sweep on the inboard portion of the wing; 2) an integral attached-flow leading-edge flap concept was added to the continuous leading edge; and 3) a deployable vortex flap concept was added to the continuous leading edge. The purpose of this report is simply to document the test configurations, test conditions, and data obtained in this investigation for future reference and analysis. No analysis is presented herein and the data only appear in tabulated format.

  18. 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.

  19. 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.

  20. 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.