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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.

  2. 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.

  3. 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.

  4. 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.

  5. Compilation of Information on the Transonic Attachment of Flows at the Leading Edges of Airfoils

    NASA Technical Reports Server (NTRS)

    Lindsey, Walter F; Landrum, Emma Jean

    1958-01-01

    Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.

  6. 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

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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

  12. The effects of leading-edge serrations on reducing flow unsteadiness about airfoils.

    NASA Technical Reports Server (NTRS)

    Schwind, R. G.; Allen, H. J.

    1973-01-01

    High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 x 1,000,000 to 6.2 x 1,000,000 on a rectangular wing of NACA 63-009 airfoil section. A wide selection of leading-edge serrations were also added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large peak in rms pressure, which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related mathematically to the airfoil trailing-edge and boundary-layer thicknesses.

  13. 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.

  14. Effects of leading and trailing edge flaps on the aerodynamics of airfoil/vortex interactions

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed A.; Sankar, L. N.; Tadghighi, H.

    1994-01-01

    A numerical procedure has been developed for predicting the two-dimensional parallel interaction between a free convecting vortex and a NACA 0012 airfoil having leading and trailing edge integral-type flaps. Special emphasis is placed on the unsteady flap motion effects which result in alleviating the interaction at subcritical and supercritical onset flows. The numerical procedure described here is based on the implicit finite-difference solutions to the unsteady two-dimensional full potential equation. Vortex-induced effects are computed using the Biot-Savart Law with allowance for a finite core radius. The vortex-induced velocities at the surface of the airfoil are incorporated into the potential flow model via the use of the velocity transpiration approach. Flap motion effects are also modeled using the transpiration approach. For subcritical interactions, our results indicate that trailing edge flaps can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, our results demonstrate the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time-dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented

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

    NASA Technical Reports Server (NTRS)

    Potapczuk, Mark G.

    1993-01-01

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

  16. Turbine Airfoil With CMC Leading-Edge Concept Tested Under Simulated Gas Turbine Conditions

    NASA Technical Reports Server (NTRS)

    Robinson, R. Craig; Hatton, Kenneth S.

    2000-01-01

    Silicon-based ceramics have been proposed as component materials for gas turbine engine hot-sections. When the Navy s Harrier fighter experienced engine (Pegasus F402) failure because of leading-edge durability problems on the second-stage high-pressure turbine vane, the Office of Naval Research came to the NASA Glenn Research Center at Lewis Field for test support in evaluating a concept for eliminating the vane-edge degradation. The High Pressure Burner Rig (HPBR) was selected for testing since it could provide temperature, pressure, velocity, and combustion gas compositions that closely simulate the engine environment. The study focused on equipping the stationary metal airfoil (Pegasus F402) with a ceramic matrix composite (CMC) leading-edge insert and evaluating the feasibility and benefits of such a configuration. The test exposed the component, with and without the CMC insert, to the harsh engine environment in an unloaded condition, with cooling to provide temperature relief to the metal blade underneath. The insert was made using an AlliedSignal Composites, Inc., enhanced HiNicalon (Nippon Carbon Co. LTD., Yokohama, Japan) fiber-reinforced silicon carbide composite (SiC/SiC CMC) material fabricated via chemical vapor infiltration. This insert was 45-mils thick and occupied a recessed area in the leading edge and shroud of the vane. It was designed to be free floating with an end cap design. The HPBR tests provided a comparative evaluation of the temperature response and leading-edge durability and included cycling the airfoils between simulated idle, lift, and cruise flight conditions. In addition, the airfoils were aircooled, uniquely instrumented, and exposed to the exact set of internal and external conditions, which included gas temperatures in excess of 1370 C (2500 F). In addition to documenting the temperature response of the metal vane for comparison with the CMC, a demonstration of improved leading-edge durability was a primary goal. First, the

  17. Direct Numerical Simulation of an Airfoil with Sand Grain Roughness on the Leading Edge

    NASA Technical Reports Server (NTRS)

    Ribeiro, Andre F. P.; Casalino, Damiano; Fares, Ehab; Choudhari, Meelan

    2016-01-01

    As part of a computational study of acoustic radiation due to the passage of turbulent boundary layer eddies over the trailing edge of an airfoil, the Lattice-Boltzmann method is used to perform direct numerical simulations of compressible, low Mach number flow past an NACA 0012 airfoil at zero degrees angle of attack. The chord Reynolds number of approximately 0.657 million models one of the test conditions from a previous experiment by Brooks, Pope, and Marcolini at NASA Langley Research Center. A unique feature of these simulations involves direct modeling of the sand grain roughness on the leading edge, which was used in the abovementioned experiment to trip the boundary layer to fully turbulent flow. This report documents the findings of preliminary, proof-of-concept simulations based on a narrow spanwise domain and a limited time interval. The inclusion of fully-resolved leading edge roughness in this simulation leads to significantly earlier transition than that in the absence of any roughness. The simulation data is used in conjunction with both the Ffowcs Williams-Hawkings acoustic analogy and a semi-analytical model by Roger and Moreau to predict the farfield noise. The encouraging agreement between the computed noise spectrum and that measured in the experiment indicates the potential payoff from a full-fledged numerical investigation based on the current approach. Analysis of the computed data is used to identify the required improvements to the preliminary simulations described herein.

  18. Measurements in a leading-edge separation bubble due to a simulated airfoil ice accretion

    NASA Technical Reports Server (NTRS)

    Bragg, M. B.; Khodadoust, A.; Spring, S. A.

    1992-01-01

    The separation bubble formed on an airfoil at low Reynolds number behind a simulated leading-edge glaze ice accretion is studied experimentally. Surface pressure and split hot-film measurements as well as flow visualization studies of the bubble reattachment point are reported. The simulated ice generates an adverse pressure gradient that causes a laminar separation bubble of the long bubble type to form. The boundary layer separates at a location on the ice accretion that is independent of angle of attack and reattaches at a downstream location 5-40 percent chord behind the leading edge, depending on the angle of attack. Velocity profiles show a large region of reverse flow that extends up from the airfoil surface as much as 2.5 percent chord. After reattachment, a thick distorted turbulent boundary layer exists. The separation bubble growth and reattachment are clearly seen in the plots of boundary-layer momentum thickness vs surface distance. Local minima and maxima in the boundary-layer momentum thickness development compare well with the shear layer transition point as indicated by the surface pressures and the reattachment point as measured from surface oil flow, respectively.

  19. Investigation of airfoil leading edge separation control with nanosecond plasma actuator

    NASA Astrophysics Data System (ADS)

    Zheng, J. G.; Cui, Y. D.; Zhao, Z. J.; Li, J.; Khoo, B. C.

    2016-11-01

    A combined numerical and experimental investigation of airfoil leading edge flow separation control with a nanosecond dielectric barrier discharge (DBD) plasma actuator is presented. Our study concentrates on describing dynamics of detailed flow actuation process and elucidating the nanosecond DBD actuation mechanism. A loose coupling methodology is employed to perform simulation, which consists of a self-similar plasma model for the description of pulsed discharge and two-dimensional Reynolds averaged Navier-Stokes (RANS) equations for the calculation of external airflow. A series of simulations of poststall flows around a NACA0015 airfoil is conducted with a Reynolds number range covering both low and high Re at Re=(0.05 ,0.15 ,1.2 ) ×106 . Meanwhile, wind-tunnel experiment is performed for two low Re flows to measure aerodynamic force on airfoil model and transient flow field with time-resolved particle image velocimetry (PIV). The PIV measurement provides possibly the clearest view of flow reattachment process under the actuation of a nanosecond plasma actuator ever observed in experiments, which is highly comparable to that predicted by simulation. It is found from the detailed simulation that the discharge-induced residual heat rather than shock wave plays a dominant role in flow control. For any leading edge separations, the preliminary flow reattachment is realized by residual heat-induced spanwise vortices. After that, the nanosecond actuator functions by continuing exciting flow instability at poststall attack angles or acting as an active trip near stall angle. As a result, the controlled flow is characterized by a train of repetitive, downstream moving vortices over suction surface or an attached turbulent boundary layer, which depends on both angle of attack and Reynolds number. The advection of residual temperature with external flow offers a nanosecond plasma actuator a lot of flexibility to extend its influence region. Animations are provided for

  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. Explanation of the effects of leading-edge tubercles on the aerodynamics of airfoils and finite wings

    NASA Astrophysics Data System (ADS)

    Saadat, Mehdi; Haj-Hariri, Hossein; Fish, Frank

    2010-11-01

    A computational study was conducted to explain the aerodynamic effect of leading edge tubercles on maximum lift coefficient, stall angle of attack (AoA), drag, and post stall characteristics for airfoils as well as finite wings. Past experiments demonstrated airfoils with leading edge tubercles do not improve Clmax, drag, or stall AoA but smoothen post stall characteristics to a great degree. In contrast to airfoils, finite wings with L.E. tubercles improved all aerodynamic characteristics. We explain the stall mechanism of the tubercled wing by considering each L.E. tubercle as a combination of a swept forward and a swept backward wing.There are 3 mechanisms (streamline curvature, accelerated stall, and upwash) that cause Clmax of airfoils with L.E. tubercles always be lower than that of smooth airfoils. We also identify two additional mechanisms which are responsible for improved post-stall characteristics of airfoils with L.E. tubercles. Finally, we discuss why finite wings with L.E. tubercles have higher Clmax and lower drag than their smooth L.E. counterparts by studying effects of wing tip, sweep, and taper ratio.

  2. 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.

  3. 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

  4. Experimental investigation of the transonic flow around the leading edge of an eroded fan airfoil

    NASA Astrophysics Data System (ADS)

    Klinner, Joachim; Hergt, Alexander; Willert, Christian

    2014-09-01

    The influence of leading edge modification on the time-averaged and instantaneous flow around a fan airfoil is investigated by particle image velocimetry (PIV), schlieren imaging and high-speed shock shadowgraphs in a transonic cascade windtunnel. In addition to a global characterization of the time-averaged flow using PIV, the instantaneous passage shock position was extracted from single-shot PIV measurements by matching the tracer velocity across the normal shock with an exponential fit. The instantaneous shock positions are assigned to a probability density distribution in order to obtain the average position and the range of fluctuations of the eroded and reference leading edge. The profiles are used to estimate the response time of the particles to the normal shock which was found to be in the sub-microsecond range. Averaged PIV measurements and the probability density of shock position from both geometries are obtained at near stall and choked conditions. In order to extract the frequency range of the shock motion, the shadow of the shock wave was tracked using high-speed shadowgraphy. The paper also provides details on the experimental implementation such as a specifically designed light-sheet probe.

  5. The effects of leading-edge serrations on reducing flow unsteadiness about airfoils, an experimental and analytical investigation

    NASA Technical Reports Server (NTRS)

    Schwind, R. G.; Allen, H. J.

    1973-01-01

    High frequency surface pressure measurements were obtained from wind-tunnel tests over the Reynolds number range 1.2 times one million to 6.2 times one million on a rectangular wing of NACA 63-009 airfoil section. Measurements were also obtained with a wide selection of leading-edge serrations added to the basic airfoil. Under a two-dimensional laminar bubble very close to the leading edge of the basic airfoil there is a large apatial peak in rms pressure. Frequency analysis of the pressure signals in this region show a large, high-frequency energy peak which is interpreted as an oscillation in size and position of the bubble. The serrations divide the bubble into segments and reduce the peak rms pressures. A low Reynolds number flow visualization test on a hydrofoil in water was also conducted. A von Karman vortex street was found trailing from the rear of the foil. Its frequency is at a much lower Strouhal number than in the high Reynolds number experiment, and is related to the trailing-edge and boundary-layer thicknesses.

  6. 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.

  7. 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.

  8. ALESEP: A computer program for the analysis of airfoil leading edge separation bubbles

    NASA Technical Reports Server (NTRS)

    Davis, R. L.; Vatsa, V. N.; Carter, J. E.

    1984-01-01

    The ALESEP program for the analysis of the inviscid/viscous interaction which occurs due to the presence of a closed laminar transitional separation bubble on an airflow is presented. The ALESEP code provides a iterative solution of the boundary layer equations expressed in an inverse formulation coupled to a Cauchy integral representation of the inviscid flow. This interaction analysis is treated as a local perturbation to a known solution obtained from a global airfoil analysis. Part of the required input to the ALESEP code are the reference displacement thickness and tangential velocity distributions. Special windward differencing may be used in the reversed flow regions of the separation bubble to accurately account for the flow direction in the discretization of the streamwise convection of momentum. The ALESEP code contains a forced transition model based on a streamwise intermittency function and a natural transition model based on a solution of the integral form of the turbulent kinetic energy equation. Instructions for the input/output, and program usage are presented.

  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. Heat-requirements for Ice Protection of a Cyclically Gas-heated, 36 Degree Swept Airfoil with Partial-span Leading-edge Slat

    NASA Technical Reports Server (NTRS)

    Gray, Vernon H; vonGlahn, Uwe H

    1956-01-01

    Heating requirements for satisfactory cyclic de-icing over a wide range of icing and operating conditions have been determined for a gas-heated, 36deg swept airfoil of 6.9-foot chord with a partial-span leading-edge slat. Comparisons of heating requirements and effectiveness were made between the slatted and unslatted portions of the airfoil. Studies were also made comparing cyclic de-icing with continuous anti-icing, and cycll.cde-icing systems with and without leading-edge ice-free parting strips. De-icing heat requirements were approximately the same with either heated or unheated parting strips because of the aerodynamic effects of the 36deg sweep angle and the spanwise saw-tooth profile of leading-edge glaze-ice deposits. Cyclic de-icing heat-source requirements were found to be one-fourth or less of the heat requirements for complete anti-icing. The primary factors that affected the performance of the cyclic de-icing heating system were ambient air temperature, heat distribution, and thermal lag.

  11. 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.

  12. A flow visualization study of the leading edge separation bubble on a NACA 0012 airfoil with simulated glaze ice. Final Report M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Khodadoust, Abdollah

    1988-01-01

    As a part of the ongoing research in aircraft icing, the leading edge separation bubble on the NACA 0012 model with a 5-min simulated glaze ice was investigated. The flow visualization methods used oil, tuft, splitter plate, smoke, and liquid crystals to get reattachment line data for the leading edge separation bubble on both surfaces of the airfoil. On the upper surface, the bubble was found to grow larger with increasing negative angles of attack and reduce in size with increasing angles of attack. The separated flow fails to reattach beyond 6 deg for the upper surface and -5 deg for the lower surface. The results of this study compared well with those of other experiments and computational results.

  13. 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.

  14. 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.

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

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

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

  16. Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partial-Span Leading-Edge Slat for Various Modes of Thermal Ice Protection

    NASA Technical Reports Server (NTRS)

    VonGlahn, Uwe H.; Gray, Vernon H.

    1954-01-01

    The effects of primary and runback ice formations on the section drag of a 36 deg swept NACA 63A-009 airfoil section with a partial-span leading-edge slat were studied over a range of angles of attack from 2 to 8 deg and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.39 to 1.23 grams per cubic meter and datum air temperatures from 10 to 25 F. The results with slat retracted showed that glaze-ice formations caused large and rapid increases in section drag coefficient and that the rate of change in section drag coefficient for the swept 63A-009 airfoil was about 2-1 times that for an unswept 651-212 airfoil. Removal of the primary ice formations by cyclic de-icing caused the drag to return almost to the bare-airfoil drag value. A comprehensive study of the slat icing and de-icing characteristics was prevented by limitations of the heating system and wake interference caused by the slat tracks and hot-gas supply duct to the slat. In general, the studies showed that icing on a thin swept airfoil will result in more detrimental aerodynamic characteristics than on a thick unswept airfoil.

  17. 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.

  18. Experiments on airfoils with trailing edge cut away

    NASA Technical Reports Server (NTRS)

    Ackeret, J

    1927-01-01

    Airfoils with their trailing edge cut away are often found on aircraft, as the fins on the hulls of flying boats and the central section of the wings for affording better visibility. It was therefore of some interest to discover the effect of such cutaways on the lift and drag and on the position of the center of pressure. For this purpose, systematic experiments were performed on two different airfoils, a symmetrical airfoil and an airfoil of medium thickness, with successive shortenings of their chords.

  19. Navier-Stokes analysis of blunt trailing edge airfoils

    NASA Technical Reports Server (NTRS)

    Stanaway, Sharon; Mccroskey, W. J.; Kroo, Ilan

    1992-01-01

    The flow around blunt trailing edge airfoils was studied by solving the Reynolds-averaged Navier-Stokes equations. The solution procedure combines a grid around the airfoil with a second grid for the wake so that the time advancement over the domain is fully implicit. This is not only very efficient for the algorithm but also allows implicit solutions of a one equation turbulence model appropriate for both boundary layers and wakes. An algebraic and two one-equation turbulence models are tested for a blunt RAE 2822 airfoil section and detailed comparisons with experimental data are presented in the trailing edge region.

  20. 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.

  1. 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.

  2. Effects of Compressibility, Pitch Rate and Reynolds Number on Unsteady Incipient Leading-Edge Boundary Layer Separation Over a Pitching Airfoil

    DTIC Science & Technology

    2007-11-02

    STRUCTURED GRID) The governing equations employed for the numerical simulation of unsteady flow past an airfoil utilizing a structured grid are...numerical simulation of aerodynamic flows . The physical boundaries of the flow are mapped into constant trans- formed coordinate lines, and this...damping term. 3.3 Geometric Conservation Law The numerical simulation of unsteady flow past a moving airfoil involves the move- ment of the computational

  3. A Computational Modeling Mystery Involving Airfoil Trailing Edge Treatments

    NASA Astrophysics Data System (ADS)

    Choo, Yeunun; Epps, Brenden

    2015-11-01

    In a curious result, Fairman (2002) observed that steady RANS calculations predicted larger lift than the experimentally-measured data for six different airfoils with non-traditional trailing edge treatments, whereas the time average of unsteady RANS calculations matched the experiments almost exactly. Are these results reproducible? If so, is the difference between steady and unsteady RANS calculations a numerical artifact, or is there a physical explanation? The goals of this project are to solve this thirteen year old mystery and further to model viscous/load coupling for airfoils with non-traditional trailing edges. These include cupped, beveled, and blunt trailing edges, which are common anti-singing treatments for marine propeller sections. In this talk, we present steady and unsteady RANS calculations (ANSYS Fluent) with careful attention paid to the possible effects of asymmetric unsteady vortex shedding and the modeling of turbulence anisotropy. The effects of non-traditional trailing edge treatments are visualized and explained.

  4. 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.

  5. 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.

  6. 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.

  7. Experimental Heat Transfer Coefficients and Friction Factors in a Rib-Roughened Leading-Edge Cooling Cavity of a Gas Turbine Airfoil

    NASA Astrophysics Data System (ADS)

    Hagan, Peter

    A gas turbine airfoil contains multiple coolant passageways. These passages usually have rib roughened wall surfaces in order to increase the heat transfer from the blade to the cooling air. Auxiliary power and compressed air is very valuable in a gas turbine, therefore low pumping power requirements are crucial. The thermal performance of three different coolant channel geometries with three different rib sizes was investigated. Heat transfer calculations were performed for Reynolds numbers ranging from 6,000 to 40,000. The performance characteristics were calculated through the use of the convective heat transfer coefficient and the friction factor. In this study, the most desirable characteristics are a high heat transfer coefficient and minimal pumping power requirements. The thermal performance of each case was determined by comparing the average Nusselt numbers to the friction factor ratio. The resulting value was then plotted against the Reynolds number for each case. The trending data indicated thermal efficiency decreases with an increasing Reynolds number for all cases. The picture data shows increased thermal efficiency at larger distances from the nose portion of the cavity. In addition, thermal efficiency was higher at the half distance of the rib pitch while areas close to the ribs saw a lower thermal efficiency. The following experimental data will show that Rig 2 and 3 are the most thermally efficient geometries, with Rig 2 requiring lower pumping power and Rig 3 having a higher average Nusselt number.

  8. 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.

  9. The effects of leading edge roughness on dynamic stall

    NASA Astrophysics Data System (ADS)

    Hrynuk, John

    2016-11-01

    Dynamic stall is a fundamental flow phenomenon that is commonly observed for insect flight and rotorcraft. Under certain conditions a leading edge vortex forms generating large but temporary lift forces. Historically, computations studying dynamic stall on airfoil shapes have struggled to predict this vortex formation time and separation point. Reduced order models and CFD have performed well when experiments have been performed to develop separation models, but this has limited the development of robust design tools. The current study looks at the effect of leading edge surface roughness on the formation of the Dynamic Stall Vortex (DSV). Roughness elements were applied to the leading edge of a NACA 0012 airfoil and PIV data of the vortex formation process was recorded. Measurements were taken at a Reynolds number of Re = 12,000 and baseline smooth NACA 0012 data was also recorded for comparison. Surface roughness elements, below the typical scale modeled by CFD, are shown to change DSV formation angle and location.

  10. Wing Leading Edge Debris Analysis

    NASA Technical Reports Server (NTRS)

    Shah, Sandeep; Jerman, Gregory

    2004-01-01

    This is a slide presentation showing the Left Wing Leading Edge (WLE) heat damage observations: Heavy "slag" deposits on select RCC panels. Eroded and knife-edged RCC rib sections. Excessive overheating and slumping of carrier panel tiles. Missing or molten attachment bolts but intact bushing. Deposit mainly on "inside" RCC panel. Deposit on some fractured RCC surface

  11. 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.

  12. 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

  13. Wake curvature and trailing edge interaction effects in viscous flow over airfoils

    NASA Technical Reports Server (NTRS)

    Melnik, R. E.

    1979-01-01

    A theory developed for analyzing viscous flows over airfoils at high Reynolds numbers is described. The theory includes a complete treatment of viscous interaction effects induced by the curved wake behind the airfoil and accounts for normal pressure gradients across the boundary layer in the trailing edge region. A brief description of a computer code that was developed to solve the extended viscous interaction equations is given. Comparisons of the theoretical results with wind tunnel data for two rear loaded airfoils at supercritical conditions are presented.

  14. 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.

  15. Some observations of surface pressures and the near wake of a blunt trailing edge airfoil

    NASA Technical Reports Server (NTRS)

    Digumarthi, R. V.; Koutsoyannis, S. P.; Karamcheti, K.

    1981-01-01

    Experiments with a truncated and untruncated airfoils of profiles NACA 640A10, were carried out in subsonic wind tunnels in a velocity range of 19m/s to 54m/s corresponding to Reynolds numbers of 200,000 to 468,000 based on the chord. Airfoil spanned the test section to achieve two dimensionality of the model. Velocity measurements, pressure measurements, and vortex shedding in the wake were measured using a hotwire and pressure transducers. The measured chordwise static pressure distribution on the smooth trailing edge airfoil along the midspan plane, agreed with the theoretical results calculated on the basis of the potential flow for that airfoil. Boundary layer profiles measured in the midspan plane, behind the maximum thickness of the airfoil show no separation of the flow. Spanwise distribution of the measured static pressure on the upper surface of the airfoil shows uniformity for both configurations with and without the boundary layer trip. This uniformity of pressure distribution and separation indicates that the flow on the airfoil was uniform and two dimensional in character.

  16. 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

  17. 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.

  18. Vertical axis wind turbine airfoil

    DOEpatents

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

    2012-12-18

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

  19. 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.

  20. Lift enhancing tabs for airfoils

    NASA Technical Reports Server (NTRS)

    Ross, James C. (Inventor)

    1994-01-01

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

  1. 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.

  2. Airfoil

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

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

  3. 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.

  4. 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.

  5. 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)

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

    NASA Technical Reports Server (NTRS)

    Manley, M. B.

    1980-01-01

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

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Preliminary Design and Evaluation of an Airfoil with Continuous Trailing-Edge Flap

    NASA Technical Reports Server (NTRS)

    Shen, Jinwei; Thornburgh, Robert P.; Kreshock, Andrew R.; Wilbur, Matthew L.; Liu, Yi

    2012-01-01

    This paper presents the preliminary design and evaluation of an airfoil with active continuous trailing-edge flap (CTEF) as a potential rotorcraft active control device. The development of structural cross-section models of a continuous trailing-edge flap airfoil is described. The CTEF deformations with MFC actuation are predicted by NASTRAN and UM/VABS analyses. Good agreement is shown between the predictions from the two analyses. Approximately two degrees of CTEF deflection, defined as the rotation angle of the trailing edge, is achieved with the baseline MFC-PZT bender. The 2D aerodynamic characteristics of the continuous trailing-edge flap are evaluated using a CFD analysis. The aerodynamic efficiency of a continuous trailing-edge flap is compared to that of a conventional discrete trailing-edge flap (DTEF). It is found that the aerodynamic characteristics of a CTEF are equivalent to those of a conventional DTEF with the same deflection angle but with a smaller flap chord. A fluid structure interaction procedure is implemented to predict the deflection of the continuous trailingedge flap under aerodynamic pressure. The reductions in CTEF deflection are overall small when aerodynamic pressure is applied: 2.7% reduction is shown with a CTEF deflection angle of two degrees and at angle of attack of six degrees. In addition, newly developed MFC-PMN actuator is found to be a good supplement to MFC-PZT when applied as the bender outside layers. A mixed MFC-PZT and MFC-PMN bender generates 3% more CTEF deformation than an MFC-PZT only bender and 5% more than an MFC-PMN only bender under aerodynamic loads.

  12. Applications of Hydrofoils with Leading Edge Protuberances

    DTIC Science & Technology

    2012-03-30

    APPLICATIONS OF HYDROFOILS WITH LEADING EDGE PROTUBERANCES Final Technical Report for Office of Naval Research contract...To) 03/30/2012 Final Technical Report 01-08-2008 to 31-12-2011 4. TITLE AND SUBTITLE Sa. CONTRACT NUMBER Applications of Hydrofoils with Leading...AVAILABILITY STATEMENT Approved for public release; distribution is unlimited. 13. SUPPLEMENTARY NOTES 14. ABSTRACT The leading edge modified hydrofoils

  13. 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.

  14. Navier-Stokes calculations and turbulence modeling in the trailing edge region of a circulation control airfoil

    NASA Technical Reports Server (NTRS)

    Viegas, John R.; Rubesin, Morris W.; Maccormack, Robert W.

    1987-01-01

    The accurate prediction of turbulent flows over curved surfaces in general and over the trailing edge region of circulation control airfoils in particular requires the coupled efforts of turbulence modelers, numerical analysts and experimentalists. The purpose of the research program in this area is described. Then, the influence on turbulence modeling of the flow characteristics over a typical circulation control wing is discussed. Next, the scope of this effort to study turbulence in the trailing edge region of a circulation control airfoil is presented. This is followed by a brief overview of the computation scheme, including the grid, governing equations, numerical method, boundary conditions and turbulence models applied to date. Then, examples of applications of two algebraic eddy viscosity models to the trailing edge region of a circulation control airfoil is presented. The results from the calculations is summarized, and conclusions drawn based on examples. Finally, the future directions of the program is outlined.

  15. Vortical Flow Structures in the Near-Wake of a Heaving Airfoil with Passively Actuated Leading and Trailing Flaps.

    NASA Astrophysics Data System (ADS)

    Siala, Firas; Totpal, Alexander; Liburdy, James

    2015-11-01

    The flow physics of flying animals has recently received significant attention, mostly in the context of developing bio-inspired micro air vehicles and oscillating flow energy harvesters. Of particular interest is the understanding of the impact of airfoil flexibility on the flow physics. Research efforts showed that some degree of surface flexibility enhanced the strength and size of the leading edge vortex. In this study, the influence of flexibility on the near-wake dynamics and flow structures is investigated using 2D PIV measurements. The experiments are conducted in a wind tunnel at a Reynolds number of 30,000 and a range of reduced frequencies from 0.09 to 0.2. The flexibility is attained using a torsion rod forming a hinge between the flap and the main wing. Vortex flow structures are visualized using large eddy scale decomposition technique and quantified using swirling strength analysis. It is found that trailing edge flexibility increases the vortex swirling strength compared to a rigid airfoil, whereas leading edge flexibility decreases the swirling strength. Furthermore, the integral length scale determined from the autocorrelation of the velocity fluctuations is found to be approximately equal to the actual vortex size. The vortex convective velocity is shown to be independent of flexibility and oscillation frequency, and it is represented by a trimodal distribution, with peak values at 0.8, 0.95 and 1 times the free stream velocity. Oregon State University.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. Leading-edge effects in bypass transition

    NASA Astrophysics Data System (ADS)

    Nagarajan, S.; Lele, S. K.; Ferziger, J. H.

    The effect of a blunt leading edge on bypass transition is studied by numerical simulation. A mixed direct and large-eddy simulation of a flat plate with a super-ellipse leading edge is carried out at various conditions. Onset and completion of transition is seen to move upstream with increasing bluntness. For sharper leading edges, at lower levels of turbulence, transition usually occurs through instabilities on low-speed streaks as observed by Jacobs & Durbin (2001) and Brandt et al. (2004) whereas increasing either the turbulence intensity or the leading-edge bluntness brings into play another mechanism. Free-stream vortices are amplified at the leading edge because of stretching. In the case of particularly strong vortices, this interaction induces a localized streamwise vortical disturbance in the boundary layer which then grows as it convects downstream and eventually breaks down to form a turbulent spot. These disturbances, which are localized and hence wavepacket-like, move at speeds in the range 0.55 U_{infty} 0.65 U_{infty} and occur in the lower portion of the boundary layer. Simulations conducted with isolated vortices confirm such a response of the boundary layer.

  1. What is the critical height of leading edge roughness for aerodynamics?

    NASA Astrophysics Data System (ADS)

    Bak, Christian; Gaunaa, Mac; Olsen, Anders S.; Kruse, Emil K.

    2016-09-01

    In this paper the critical leading edge roughness height is analyzed in two cases: 1) leading edge roughness influencing the lift-drag ratio and 2) leading edge roughness influencing the maximum lift. The analysis was based on wind tunnel measurements on the airfoils NACA0015, Risoe-B1-18 and Risoe-C2-18 and at three different Reynolds numbers with two different leading edge roughness tape heights. Firstly, an analysis of the momentum thickness as function of Reynolds number was carried out based on the boundary layer theory by Thwaites. Secondly, the wind tunnel measurements combined with panel code predictions of the boundary layer momentum thickness created the basis for determining the impact of roughness on the aerodynamic performance. The critical heights were related to the Reynolds numbers and thereby the size of the wind turbines.

  2. Pneumatic Spoiler Controls Airfoil Lift

    NASA Technical Reports Server (NTRS)

    Hunter, D.; Krauss, T.

    1991-01-01

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

  3. 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.

  4. A Theoretical Study of Leading Edge Noise

    DTIC Science & Technology

    2008-05-01

    as described by Lockard and Morris [13], have given the solution for the unsteady loading on airfoils of arbitrary shape. In this paper we will show...Proceedings of the 7th AIAA/CEAS Aeroacoustics Conference, Maastricht, may 2001, AIAA paper no 2001-2209 [13] Lockard ,D.P. and Morris, P.J...airfoils in turbulent flows, J. Acoustical Soc. Am. vol 116(3), ppl416-1426, 2004. [8] Lockard D.P. and Morris P.J., Radiated Noise from airfoils in

  5. 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.

  6. Summary of Airfoil Data

    NASA Technical Reports Server (NTRS)

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

    1945-01-01

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

  7. 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.

  8. Manipulation of Leading-Edge Vortex Evolution by Applied Suction

    NASA Astrophysics Data System (ADS)

    Buchholz, James; Akkala, James

    2016-11-01

    The generation and shedding of vortices from unsteady maneuvering bodies can be characterized within a framework of vorticity transport, accounting for the effects of multiple sources and sinks of vorticity on the overall circulation of the vortex system. On a maneuvering wing, the diffusive flux of secondary vorticity from the surface is a critical contributor to the strength and dynamics of the leading-edge vortex, suggesting that flow control strategies targeting the manipulation of the secondary vorticity flux and the secondary vortex may provide an effective means of manipulating vortex development. Suction has been applied in the vicinity of the secondary vortex during the downstroke of a periodically-plunging flat-plate airfoil, and the flow evolution and aerodynamic loads are compared to the baseline case in which suction is not applied. Observation of the resulting surface pressure distribution and flow evolution suggest that the secondary flux of vorticity and the evolution of the flow field can be altered subject to appropriate position of the suction ports relative to the developing vortex structures, and at a specific temporal window in the development of the vortex. This work was supported by the Air Force Office of Scientific Research, Grant Number FA9550-16-1-0107 and NSF EPSCoR Grant Number EPS1101284.

  9. 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.

  10. 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.

  11. Leading-edge vortices in insect flight

    NASA Astrophysics Data System (ADS)

    Ellington, Charles P.; van den Berg, Coen; Willmott, Alexander P.; Thomas, Adrian L. R.

    1996-12-01

    INSECTS cannot fly, according to the conventional laws of aerodynamics: during flapping flight, their wings produce more lift than during steady motion at the same velocities and angles of attack1-5. Measured instantaneous lift forces also show qualitative and quantitative disagreement with the forces predicted by conventional aerodynamic theories6-9. The importance of high-life aerodynamic mechanisms is now widely recognized but, except for the specialized fling mechanism used by some insect species1,10-13, the source of extra lift remains unknown. We have now visualized the airflow around the wings of the hawkmoth Manduca sexta and a 'hovering' large mechanical model-the flapper. An intense leading-edge vortex was found on the down-stroke, of sufficient strength to explain the high-lift forces. The vortex is created by dynamic stall, and not by the rotational lift mechanisms that have been postulated for insect flight14-16. The vortex spirals out towards the wingtip with a spanwise velocity comparable to the flapping velocity. The three-dimensional flow is similar to the conical leading-edge vortex found on delta wings, with the spanwise flow stabilizing the vortex.

  12. Airfoils for wind turbine

    SciTech Connect

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

    2000-05-30

    Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

  13. Airfoils for wind turbine

    DOEpatents

    Tangler, James L.; Somers, Dan M.

    2000-01-01

    Airfoils for the tip and mid-span regions of a wind turbine blade have upper surface and lower surface shapes and contours between a leading edge and a trailing edge that minimize roughness effects of the airfoil and provide maximum lift coefficients that are largely insensitive to roughness effects. The airfoil in one embodiment is shaped and contoured to have a thickness in a range of about fourteen to seventeen percent, a Reynolds number in a range of about 1,500,000 to 2,000,000, and a maximum lift coefficient in a range of about 1.4 to 1.5. In another embodiment, the airfoil is shaped and contoured to have a thickness in a range of about fourteen percent to sixteen percent, a Reynolds number in a range of about 1,500,000 to 3,000,000, and a maximum lift coefficient in a range of about 0.7 to 1.5. Another embodiment of the airfoil is shaped and contoured to have a Reynolds in a range of about 1,500,000 to 4,000,000, and a maximum lift coefficient in a range of about 1.0 to 1.5.

  14. 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.

  15. 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.

  16. A fast leading-edge pulse generator

    NASA Astrophysics Data System (ADS)

    Wang, R.

    1986-01-01

    The pulse generator consists of ECL semiconductor integrated circuits, high speed transistors and step restorer diodes, among others; its circuitry is simple. The leading edge of the output pulse is less than 100 ps, and the output impedance is 50 ohms. An ECL four-wire receiver connected as a closed loop circut is used in the oscillator section of the set. The pulse frequency varies as low as 10 Hz and as high as 100 MHz. The control of pulse with is based on the subtraction of two pulse widths. The output pulse width may be less than 10 ns and the maximum width may be as wide as an oscillator half cycle. The pulse amplitude is continuously adjustable from + or - 35 mV to + or - 5 V. The operating principle of the oscillator stage, a simplified logic diagram, waveforms at various points, a rectifier circuit in the first stage, positive pulse channel circuit, and an adjustable power source are shown.

  17. Observations on Leading-Edge Vortex Development

    NASA Astrophysics Data System (ADS)

    Glenn, Michael; Lang, Amy; Wahidi, Redha; Wilroy, Jacob

    2016-11-01

    Most of an insect's lift comes from the leading edge vortex (LEV) that they produce when flapping their wings. There are many variables that make a LEV either stronger or weaker such as: roughness from the scales on their wings, angle of attack (AoA) of wing, size of the wing, and speed of the wing during flapping motion. Experiments were conducted to study LEV development to gain a better understanding of butterfly flight and the importance of LEV formation. The variables emphasized in this particular experiment were the chord length Reynolds numbers. Two smooth plates of 4 inches and 7 inches were compared in this experiment with Re of 1500 and 3000. Matlab was used to track the LEV location and calculate the vorticity and circulation magnitudes. Differences in LEV vortex strength as a function of chord length will be presented. Funding was provided by NSF REU site Grant EEC 1358991 and CBET Grant 1628600.

  18. 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.

  19. 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.

  20. 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.

  1. Experimental investigation of leading-edge thrust at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Wood, R. M.; Miller, D. S.

    1983-01-01

    Wings, designed for leading edge thrust at supersonic speeds, were investigated in the Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, 2.16, and 2.36. Experimental data were obtained on a uncambered wing which had three interchangeable leading edges that varied from sharp to blunt. The leading edge thrust concept was evaluated. Results from the investigation showed that leading edge flow separation characteristics of all wings tested agree well with theoretical predictions. The experimental data showed that significant changes in wing leading edge bluntness did not affect the zero lift drag of the uncambered wings.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Effect of trailing edge shape on the separated flow characteristics around an airfoil at low Reynolds number: A numerical study

    NASA Astrophysics Data System (ADS)

    Thomareis, Nikitas; Papadakis, George

    2017-01-01

    Direct numerical simulations of the flow field around a NACA 0012 airfoil at Reynolds number 50 000 and angle of attack 5° with 3 different trailing edge shapes (straight, blunt, and serrated) have been performed. Both time-averaged flow characteristics and the most dominant flow structures and their frequencies are investigated using the dynamic mode decomposition method. It is shown that for the straight trailing edge airfoil, this method can capture the fundamental as well as the subharmonic of the Kelvin-Helmholtz instability that develops naturally in the separating shear layer. The fundamental frequency matches well with relevant data in the literature. The blunt trailing edge results in periodic vortex shedding, with frequency close to the subharmonic of the natural shear layer frequency. The shedding, resulting from a global instability, has an upstream effect and forces the separating shear layer. Due to forcing, the shear layer frequency locks onto the shedding frequency while the natural frequency (and its subharmonic) is suppressed. The presence of serrations in the trailing edge creates a spanwise pressure gradient, which is responsible for the development of a secondary flow pattern in the spanwise direction. This pattern affects the mean flow in the near wake. It can explain an unexpected observation, namely, that the velocity deficit downstream of a trough is smaller than the deficit after a protrusion. Furthermore, the insertion of serrations attenuates the energy of vortex shedding by de-correlating the spanwise coherence of the vortices. This results in weaker forcing of the separating shear layer, and both the subharmonics of the natural frequency and the shedding frequency appear in the spectra.

  9. Transition and Turbulence Structure in the Boundary Layers of an Oscillating Airfoil

    DTIC Science & Technology

    1989-12-31

    measurements near the trailing edge of an oscillating NACA 64A010 airfoil operating at various fiequencies and in unstalled condition. It was...to identify: (i) the conditions under which a sinusoidally oscillating NACA 0012 airfoil operates with a leading edge separation bubble; (ii) the...vortex formation. These experiments were conducted using an NACA 0012 airfoil with a tripping wire to promote transition of the boundary layer. The

  10. 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.

  11. 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.

  12. Flow-field measurements on an airfoil with an oscillating trailing-edge using holographic interferometry

    NASA Technical Reports Server (NTRS)

    Bachalo, W. D.

    1984-01-01

    Holographic interferometry data were acquired on an NACA 64A010 airfoil with an oscillating flap. The airfoil was installed in the Ames 11-Foot Transonic Wind Tunnel between splitter plates. Recordings were made at discrete phase angles of the oscillation. The interferometry results provided detailed flow visualization of the shock boundary-layer interaction and the separated flow. Quantitative results were extracted from the interferograms to produce pressure data. These results were compared to the surface pressures obtained with the surface pressure taps. Excellent agreement was found for low angles of incidence. At larger angles of incidence, the flow had greater three-dimensionality, and the results were not in good agreement in some regions of the flow field. Mach contours were traced for representative flow conditions. Wake profiles were also obtained using the assumption of constant pressure across the wake and the Crocco relationship.

  13. 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.

  14. 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.

  15. Wind Tunnel Aeroacoustic Tests of Six Airfoils for Use on Small Wind Turbines: Preprint

    SciTech Connect

    Migliore, P.; Oerlemans, S.

    2003-12-01

    Aeroacoustic tests of seven airfoils were performed in an open jet anechoic wind tunnel. Six of the airfoils are candidates for use on small wind turbines operating at low Reynolds number. One airfoil was tested for comparison to benchmark data. Tests were conducted with and without boundary layer tripping. In some cases a turbulence grid was placed upstream in the test section to investigate inflow turbulence noise. An array of 48 microphones was used to locate noise sources and separate airfoil noise from extraneous tunnel noise. Trailing edge noise was dominant for all airfoils in clean tunnel flow. With the boundary layer untripped, several airfoils exhibited pure tones that disappeared after proper tripping was applied. In the presence of inflow turbulence, leading edge noise was dominant for all airfoils.

  16. 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.

  17. 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.

  18. 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.

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

    NASA Technical Reports Server (NTRS)

    Stivers, Louis S., Jr.

    1947-01-01

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

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. 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.

  5. Leading-Edge Learning: Two Views.

    ERIC Educational Resources Information Center

    Abernathy, Donna J.

    1999-01-01

    Peter Senge and Jack Welch share thoughts about what it means to learn and lead into the next century. Senge urges leaders to be aware of the economic and the natural environment. Welch asserts that an organization's ability to learn and translate learning into action is the ultimate competitive advantage. (JOW)

  6. 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.

  7. 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.

  8. Leading-edge receptivity for blunt-nose bodies

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

    Boundary-layer receptivity in the leading edge region for bodies with blunt leading edges is investigated in this research program. Receptivity theory provides the link between the unsteady disturbance environment in the freestream and the initial amplitudes of 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.

  9. 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.

  10. The Effect of Aerodynamic Evaluators on the Multi-Objective Optimization of Flatback Airfoils

    NASA Astrophysics Data System (ADS)

    Miller, M.; Slew, K. Lee; Matida, E.

    2016-09-01

    With the long lengths of today's wind turbine rotor blades, there is a need to reduce the mass, thereby requiring stiffer airfoils, while maintaining the aerodynamic efficiency of the airfoils, particularly in the inboard region of the blade where structural demands are highest. Using a genetic algorithm, the multi-objective aero-structural optimization of 30% thick flatback airfoils was systematically performed for a variety of aerodynamic evaluators such as lift-to-drag ratio (Cl/Cd), torque (Ct), and torque-to-thrust ratio (Ct/Cn) to determine their influence on airfoil shape and performance. The airfoil optimized for Ct possessed a 4.8% thick trailing-edge, and a rather blunt leading-edge region which creates high levels of lift and correspondingly, drag. It's ability to maintain similar levels of lift and drag under forced transition conditions proved it's insensitivity to roughness. The airfoil optimized for Cl/Cd displayed relatively poor insensitivity to roughness due to the rather aft-located free transition points. The Ct/Cn optimized airfoil was found to have a very similar shape to that of the Cl/Cd airfoil, with a slightly more blunt leading-edge which aided in providing higher levels of lift and moderate insensitivity to roughness. The influence of the chosen aerodynamic evaluator under the specified conditions and constraints in the optimization of wind turbine airfoils is shown to have a direct impact on the airfoil shape and performance.

  11. 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.

  12. 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.

  13. Direct numerical simulation of broadband trailing edge noise from a NACA 0012 airfoil

    NASA Astrophysics Data System (ADS)

    Mehrabadi, Mohammad; Bodony, Daniel

    2016-11-01

    Commercial jet-powered aircraft produce unwanted noise at takeoff and landing when they are close to near-airport communities. Modern high-bypass-ratio turbofan engines have reduced jet exhaust noise sufficiently such that noise from the main fan is now significant. In preparation for a large-eddy simulation of the NASA/GE Source Diagnostic Test Fan, we study the broadband noise due to the turbulent flow on a NACA 0012 airfoil at zero degree angle-of-attack, a chord-based Reynolds number of 408,000 and a Mach number of 0.115 using direct numerical simulation (DNS) and wall-modeled large-eddy simulation (WMLES). The flow conditions correspond to existing experimental data. We investigate the roughness-induced transition-to-turbulence and sound generation from a DNS perspective as well as examine how these two features are captured by a wall model. Comparisons between the DNS- and WMLES-predicted noise are made and provide guidance on the use of WMLES for broadband fan noise prediction. AeroAcoustics Research Consortium.

  14. Streamwise Oscillation of Airfoils into Reverse Flow

    NASA Astrophysics Data System (ADS)

    Granlund, Kenneth; Jones, Anya; Ol, Michael

    2015-11-01

    An airfoil in freestream is oscillated in streamwise direction to cyclically enter reverse flow. Measured lift is compared to analytical blade element theories. Advance ratio, reduced frequency and angle of attack is varied within those typical for helicopters. Experimental results reveal that lift does not become negative in the flow reversal part, contradicting one theory and supported by another. Flow visualization reveal the leading edge vortex advecting against the freestream to a point in front of the leading edge.

  15. 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.

  16. Turbine airfoil to shroud attachment method

    SciTech Connect

    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.

  17. Incidence angle effects on convected gust airfoil noise

    NASA Technical Reports Server (NTRS)

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

    1983-01-01

    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.

  18. Design and Experimental Results for the S411 Airfoil

    DTIC Science & Technology

    2010-08-01

    unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 8...produces a suction peak at higher lift coefficients, which ensures that tran- sition on the upper surface will occur very near the leading edge. Thus...like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar

  19. Design and Experimental Results for the S406 Airfoil

    DTIC Science & Technology

    2010-08-01

    point B is not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly...in a leading edge that produces a suction peak at higher lift coefficients, which ensures that transition on the upper surface will occur very near...3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar flow

  20. 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.

  1. 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.

  2. 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.

  3. 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.

  4. Low speed airfoil study

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.

    1977-01-01

    Airfoil geometries were developed for low speed high lift applications, such as general aviation aircraft, propellers and helicopter rotors. The primary effort was to determine the extent to which the application of turbulent boundary layer separation criteria, plus manipulation of other input parameters, specifically trailing edging velocity ratio, could be utilized to achieve high C sub Lmax airfoils with relatively low drag at C sub Lmax. Both single-element and double-element airfoils were considered. Wind tunnel testing of some airfoils was included.

  5. Reversible airfoils for stopped rotors in high speed flight

    NASA Astrophysics Data System (ADS)

    Niemiec, Robert; Jacobellis, George; Gandhi, Farhan

    2014-10-01

    This study starts with the design of a reversible airfoil rib for stopped-rotor applications, where the sharp trailing-edge morphs into the rounded leading-edge, and vice-versa. A NACA0012 airfoil is approximated in a piecewise linear manner and straight, rigid outer profile links used to define the airfoil contour. The end points of the profile links connect to control links, each set on a central actuation rod via an offset. Chordwise motion of the actuation rod moves the control and the profile links and reverses the airfoil. The paper describes the design methodology and evolution of the final design, based on which two reversible airfoil ribs were fabricated and used to assemble a finite span reversible rotor/wing demonstrator. The profile links were connected by Aluminum strips running in the spanwise direction which provided stiffness as well as support for a pre-tensioned elastomeric skin. An inter-rib connector with a curved-front nose piece supports the leading-edge. The model functioned well and was able to reverse smoothly back-and-forth, on application and reversal of a voltage to the motor. Navier-Stokes CFD simulations (using the TURNS code) show that the drag coefficient of the reversible airfoil (which had a 13% maximum thickness due to the thickness of the profile links) was comparable to that of the NACA0013 airfoil. The drag of a 16% thick elliptical airfoil was, on average, about twice as large, while that of a NACA0012 in reverse flow was 4-5 times as large, even prior to stall. The maximum lift coefficient of the reversible airfoil was lower than the elliptical airfoil, but higher than the NACA0012 in reverse flow operation.

  6. Mechanisms of sound amplification and sound reduction in the flapping flight of side-by-side airfoils

    NASA Astrophysics Data System (ADS)

    Manela, A.; Halachmi, M.

    2015-06-01

    The acoustic signature of side-by-side airfoils, subject to small-amplitude harmonic pitching and incoming flow unsteadiness, is investigated. The two-dimensional near-field problem is formulated using thin-airfoil theory, where flow unsteadiness is modeled as a passing line vortex, and wake evolution is calculated via the Brown and Michael formula. Assuming that the setup is acoustically compact, acoustic radiation is obtained by means of the Powell-Howe acoustic analogy. The associated compact Green's function is calculated numerically using potential-flow analysis of the fluid-structure flow domain. Results, comparing the acoustic radiation of the double-airfoil system to a reference case of a single airfoil, point to several mechanisms of sound attenuation and sound amplification, caused by airfoil-airfoil and airfoils-wake interactions. It is found that counter-phase pitching of the airfoils results in effective cloaking of the system, which otherwise becomes significantly noisy (as a 5/2-power of the pitching frequency) at large frequencies. In addition, depending on the distance between airfoils, in-phase pitching may result in an acoustic signature equivalent to a single airfoil (when the airfoils are adjacent) or to two separate airfoils (when the airfoils are far apart). In general, flow unsteadiness produces more sound when interacting with a double (compared with a single) airfoil setup. However, airfoils' nonlinear wake-wake interactions give rise to a sound reduction mechanism, which becomes most efficient at times when incoming vorticity passes above airfoils' leading and trailing edges. The present scheme can be readily extended to consider the acoustic properties of various double-airfoil configurations, as well as multiple (> 2) airfoil setups.

  7. Pulsed Film Cooling on a Turbine Blade Leading Edge

    DTIC Science & Technology

    2009-09-01

    F = 0.590, DC = 50% ........................................................... 142 Fig. 6.4 Leading edge model and camera positions...steady M = 0.25. ( Camera Angle 1...147 Fig. 7.2 Near parallel to surface view of coolant, steady M = 0.25. ( Camera Angle 2

  8. 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.

  9. 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.

  10. Near-wall serpentine cooled turbine airfoil

    SciTech Connect

    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.

  11. The S415 and S418 Airfoils

    DTIC Science & Technology

    2010-08-01

    airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 4.) This characteristic is related to the...edge with increasing (decreasing) lift coefficient. This feature results in a leading-edge shape that produces a suction peak at higher lift...should look like sketch 3. Sketch 3 1Director, Institute for Aerodynamics and Gas Dynamics, University of Stuttgart, Germany, 1974–1985.5 No suction

  12. Spline-Based Smoothing of Airfoil Curvatures

    NASA Technical Reports Server (NTRS)

    Li, W.; Krist, S.

    2008-01-01

    Constrained fitting for airfoil curvature smoothing (CFACS) is a splinebased method of interpolating airfoil surface coordinates (and, concomitantly, airfoil thicknesses) between specified discrete design points so as to obtain smoothing of surface-curvature profiles in addition to basic smoothing of surfaces. CFACS was developed in recognition of the fact that the performance of a transonic airfoil is directly related to both the curvature profile and the smoothness of the airfoil surface. Older methods of interpolation of airfoil surfaces involve various compromises between smoothing of surfaces and exact fitting of surfaces to specified discrete design points. While some of the older methods take curvature profiles into account, they nevertheless sometimes yield unfavorable results, including curvature oscillations near end points and substantial deviations from desired leading-edge shapes. In CFACS as in most of the older methods, one seeks a compromise between smoothing and exact fitting. Unlike in the older methods, the airfoil surface is modified as little as possible from its original specified form and, instead, is smoothed in such a way that the curvature profile becomes a smooth fit of the curvature profile of the original airfoil specification. CFACS involves a combination of rigorous mathematical modeling and knowledge-based heuristics. Rigorous mathematical formulation provides assurance of removal of undesirable curvature oscillations with minimum modification of the airfoil geometry. Knowledge-based heuristics bridge the gap between theory and designers best practices. In CFACS, one of the measures of the deviation of an airfoil surface from smoothness is the sum of squares of the jumps in the third derivatives of a cubicspline interpolation of the airfoil data. This measure is incorporated into a formulation for minimizing an overall deviation- from-smoothness measure of the airfoil data within a specified fitting error tolerance. CFACS has been

  13. 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

  14. 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.

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

    PubMed

    Borazjani, Iman; Daghooghi, Mohsen

    2013-04-07

    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.

  16. 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.

  17. 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.

  18. 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.

  19. 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

  20. Design and Experimental Results for a Natural-Laminar-Flow Airfoil for General Aviation Applications

    NASA Technical Reports Server (NTRS)

    Somers, D. M.

    1981-01-01

    A natural-laminar-flow airfoil for general aviation applications, the NLF(1)-0416, was designed and analyzed theoretically and verified experimentally in the Langley Low-Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low-speed airfoils with the low cruise drag of the NACA 6-series airfoils was achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge was also met. Comparisons of the theoretical and experimental results show excellent agreement. Comparisons with other airfoils, both laminar flow and turbulent flow, confirm the achievement of the basic objective.

  1. 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.

  2. Effect of pivot location and passive heave on propulsion from a pitching airfoil

    NASA Astrophysics Data System (ADS)

    Mackowski, A. W.; Williamson, C. H. K.

    2017-01-01

    We experimentally investigate the propulsive characteristics of a pitching NACA 0012 airfoil section, with emphasis on thrust and propulsive efficiency, at a Reynolds number of 1.7 ×104 . For the sake of mechanical simplicity, we consider an airfoil restricted to a single actuator in the pitching direction. We examine the effect of changing the airfoil's axis of rotation, finding that contrary to Garrick's linear theory, there exists a pitching axis near the airfoil that maximizes propulsive efficiency. Next, we examine the effect of placing passive springs on the airfoil in the heave (transverse) direction using our Cyber-Physical Fluid Dynamics technique. This elastic heaving motion allows the airfoil to combine pitching and heaving modes while being actuated only in the pitching direction. Two sets of dynamics are considered: one case where the airfoil is weighted unevenly and pitched about its center of mass (so that the resulting heaving motion is independent of inertial forces), and another case where the airfoil's center of mass is fixed at its centroid. For pitching at an amplitude of 8∘ and a reduced frequency k of two, we find that elastic heave produces a maximum propulsive efficiency of 35%, compared to 25% without any heave motion. Further, while operating at the same efficiency as the static-pivot case, we find that passive heaving greatly increases the magnitude of the airfoil's thrust. The airfoil configurations with highest propulsive efficiency generally involve pitching near or ahead of the airfoil's leading edge.

  3. Two-dimensional wind-tunnel tests of a NASA supercritical airfoil with various high-lift systems. Volume 1: Data analysis

    NASA Technical Reports Server (NTRS)

    Omar, E.; Zierten, T.; Mahal, A.

    1977-01-01

    High-lift systems for a NASA, 9.3%, method for calculating the viscous flow about two-dimensional multicomponent airfoils was evaluated by comparing its predictions with test data. High-lift systems derived from supercritical airfoils were compared in terms of performance to high-lift systems derived from conventional airfoils. The high-lift systems for the supercritical airfoil were designed to achieve maximum lift and consisted of: a single-slotted flap; a double-slotted flap and a leading-edge slat; and a triple-slotted flap and a leading-edge slat. Agreement between theoretical predictions and experimental results are also discussed.

  4. 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.

  5. Applications of Euler equations to sharp edge delta wings with leading edge vortices

    NASA Technical Reports Server (NTRS)

    Murman, Earll M.; Rizzi, Arthur

    1986-01-01

    Studies on the solution of discrete Euler equations past swept delta wing configurations with sharp leding edges are presented. Freestream Mach numbers range from zero to supersonic, although the Mach number normal to the leading edge is subsonic for all cases discussed. A few examples are given to show the application of the numerical methods to representative problems. The major dicussion is directed at the application of Computational Fluid Dynamics to the understanding of the fundamental fluid mechanic mechanisms of this class of flows.

  6. 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.

  7. 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.

  8. 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.

  9. Leading Edge Heat Shield for Wings of Spacecraft

    NASA Technical Reports Server (NTRS)

    Stewart, David A. (Inventor)

    1998-01-01

    A heat shield for thermally insulating the leading edge of a wing of a spacecraft during ascent and reentry includes a plurality of rigid tiles. Each tile is formed with a pie-shaped element which interlocks with the complementarily-formed element of another tile. The combination of structure afforded by the pie-shaped elements substantially impedes hypersonic flow of any gases that might enter the gaps between tiles.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. On the leading edge vortex of thin wings

    NASA Astrophysics Data System (ADS)

    Arredondo, Abel; Viola, Ignazio Maria

    2016-11-01

    On thin wings, the sharp leading edge triggers laminar separation followed by reattachment, forming a Leading Edge Vortex (LEV). This flow feature is of paramount importance because, if periodically shed, it leads to large amplitude load fluctuations, while if stably attached to the wing, it can provide lift augmentation. We found that on asymmetric-spinnaker-type yacht sails, the LEV can be stable despite the relatively low sweep (30°). This finding, which was recently predicted numerically by Viola et al., has been confirmed through current flume tests on a 1:115th model scale sail. Forces were measured and Particle Image Velocimetry was performed on four horizontal sail sections at a Reynolds number of 1.7x104. Vortex detection revealed that the LEV becomes progressively larger and more stable towards the highest sections, where its axis has a smaller angle with respect to the freestream velocity. Mapping the sail section on a rotating cylinder through a Joukowski transformation, we quantified the lift augmentation provided by the LEV on each sail section. These results open up new sail design strategies based on the manipulation of the LEV and can be applicable to the wings of unmanned aerial vehicles and underwater vehicles. Project funded by Conacyt.

  15. Wind-tunnel test results of airfoil modifications for the EA-6B

    NASA Technical Reports Server (NTRS)

    Sewall, W. G.; Mcghee, R. J.; Ferris, J. C.

    1987-01-01

    Wind-tunnel tests have been conducted (to determine the effects on airfoil performance for several airfoil modifications) for the EA-6B Wing Improvement Program. The modifications consist of contour changes to the leading-edge slat and trailing-edge flap to provide a higher low-speed maximum lift with no high-speed cruise-drag penalty. Airfoil sections from the 28- and 76-percent span stations were selected as baseline shapes with the major testing devoted to the inboard airfoil section (28-percent span station). The airfoil modifications increased the low-speed maximum lift coefficient between 20 and 35 percent over test conditions of 3 to 14 million chord Reynolds number and 0.14 to 0.34 Mach number. At the high-speed test conditions of 0.4 to 0.80 Mach number and 10 million chord Reynolds number, the modified airfoils had either matched or had lower drag coefficients for all normal-force coefficients above 0.2 as compared to the baseline airfoil. At normal-force coefficients less than 0.2, the baseline (original) airfoil had lower drag coefficients than any of the modified airfoils.

  16. 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.

  17. 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.

  18. 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.

  19. Design and Experimental Results for the S415 Airfoil

    DTIC Science & Technology

    2010-08-01

    polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 8.) This... suction peak at higher lift coefficients, which ensures that transition on the upper surface will occur very near the leading edge. Thus, the...pressure distribution should look like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a moderately adverse pressure

  20. The S411, S412, and S413 Airfoils

    DTIC Science & Technology

    2010-08-01

    not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant...in a leading edge that produces a suction peak at higher lift coefficients, which ensures that tran- sition on the upper surface will occur very near...This concept allows a wide low-drag range to be achieved and increases the loading in the leading-edge region. The forward loading serves to balance

  1. 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.

  2. 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

  3. Deformation Zones along Leading Edges of Thrust Faults

    NASA Astrophysics Data System (ADS)

    Johnsion, A. M.; Huang, W. O.

    2006-12-01

    Deformation zones and concomitant damage along earthquake ruptures were recognized long ago in studies of the 1906 San Francisco earthquake. Most of the previous investigations of deformation zones have been of features along strike slip earthquake ruptures. This research, in contrast, describes and analyzes deformation zones observed along leading edges of two thrusts—the 1999 Chi Chi rupture in Taiwan and the Sylmar segment of the 1971 San Fernando Valley rupture in California. Deformation zones along the leading edges of the Chi Chi and Sylmar thrusts have several features and conditions in common: Both formed over reverse faults that dip 30° to 45° at shallow depths. Both accommodated different amounts of strike slip as well as reverse, dip slip along their traces. Both had associated ground deformation zones containing various kinds of smaller structures, including low amplitude folds, small fractures such as strike slip and thrust faults and tension cracks. Both had broken and tilted dwellings and other man made structures within them. Also, both deformation zones were highly asymmetric: the deformation zone in the hanging wall was much wider than that in the footwall. We have combined a proper yielding criterion for permanent (plastic) deformation at the ground surface produced by slip on a buried dislocation that is propagating upward to the surface. The result is an approximate simulation of the growth of ground deformation zones analogous to those we see in the field. The specific phenomena we investigate with the method include: 1). Compressional deformation zones straddling earthquake thrust ruptures. 2). Asymmetric deformation zones. Compressional deformation zones are much wider in the hanging wall than the footwall of thrusts. 3). A thrust deformation zone also includes an extensional zone in the hanging wall. 4). Where there is also left lateral, strike shift across the deformation zone, a zone of left lateral distortion is sandwiched by zones

  4. 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.

  5. Roughness Effects on the Formation of a Leading Edge Vortex

    NASA Astrophysics Data System (ADS)

    Elliott, Cassidy; Lang, Amy; Wahidi, Redha; Wilroy, Jacob

    2016-11-01

    Microscopic scales cover the wings of Monarch butterflies, creating a patterned surface. This patterning is an important natural flow control mechanism that is thought to delay the growth of the leading edge vortex (LEV) produced by the flapping motion of a wing. The increased skin friction caused by the scales leads to a weaker LEV being shed into the butterfly's wake, lessening drag and increasing flight efficiency. To test this theory, a plate of random roughness was designed in SolidWorks and printed on the Objet 30 Pro 3D printer. A 2x3x5 cubic foot tow tank was used to test the rough plate at Reynold's numbers of 1500, 3000, and 6000 (velocities of 8, 16, and 32 mm/s) at an angle of attack of 45 degrees. Particle Image Velocimetry (PIV) captured images of the LEV generated by the plate when towed upwards through the particle-seeded flow. Codes written in MatLab were used to automatically track and determine the strength of the LEV. Circulation values for the randomly-rough plate were then compared to the same values generated in a previous experiment that used a smooth plate and a grooved plate to determine the effect of the patterning on vortex development. Funding provided by NSF REU site Grant EEC 1358991 and CBET 1628600.

  6. Low-speed aerodynamic characteristics of a highly swept arrow wing configuration with several deflected leading edge concepts

    NASA Technical Reports Server (NTRS)

    Gentry, G. L., Jr.; Coe, P. L., Jr.

    1980-01-01

    The effectiveness of leading edge concepts for minimizing or controlling leading edge flow separation was studied. Emphasis was placed on low speed performance, stability, and control characteristics of configurations with highly swept wings. Simple deflection of the leading edge, a variable camber leading edge system, and a leading edge vortex flow system were among the concepts studied. The data are presented without analysis.

  7. 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.

  8. 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.

  9. 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

  10. 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.

  11. Airfoil structure

    DOEpatents

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

    1998-01-13

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

  12. Airfoil structure

    DOEpatents

    Frey, Gary A.; Twardochleb, Christopher Z.

    1998-01-01

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

  13. 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

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

    PubMed

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

    2016-04-12

    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.

  15. 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

  16. Absolute Instability in Swept Leading-Edge Boundary Layers

    NASA Astrophysics Data System (ADS)

    Lin, R.-S.; Li, F.; Malik, M. R.

    1997-11-01

    Absolute instabilities in the swept Hiemenz flow and flows over Poll's swept cylinder are studied. It is assumed that the span is infinite and the laminar flow field is subjected to a line impulsive excitation so that the spanwise wavenumber (β) is taken to be real, which is akin to the rotating disk study made by Lingwood.footnote Lingwood, R. J., J. Fluid Mech., 299, 17, 1995. We found that these flows can be absolutely unstable in the chordwise (x) direction. The pinch-point singularities formed by the coalescence of two distinct spatial branches can lie either below or above the real α-axis. The pinch points with a positive αi imply the existence of an unstable disturbance propagating against the mainstream, which has never been observed before. It is found that singularities of pinch type occur in a region very close to the leading edge, therefore the attachment-line Reynolds number is used to correlate the onset of absolute instability. The critical Reynolds number for absolute instability is found to be about R=540 compared to 583 for the attachment-line instability. Provided the non-linear behavior of this absolute instability is sufficient to trigger the laminar to turbulent transition, then it would cause a complete loss of laminar flow on a swept wing as does the attachment-line instability.

  17. Input description for Jameson's three-dimensional transonic airfoil analysis program

    NASA Technical Reports Server (NTRS)

    Newman, P. A.; Davis, R. M.

    1974-01-01

    The input parameters are presented for a computer program which performs calculations for inviscid isentropic transonic flow over three dimensional airfoils with straight leading edges. The free stream Mach number is restricted only by the isentropic assumption. Weak shock waves are automatically located where they occur in the flow. The finite difference form of the full equation for the velocity potential is solved by the method of relaxation, after the flow exterior to the airfoil is mapped to the upper half plane.

  18. 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.

  19. The S407, S409, and S410 Airfoils

    DTIC Science & Technology

    2010-08-01

    coefficient at point B is not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is...feature results in a leading-edge shape that produces a suction peak at higher lift coefficients, which ensures that transition on the upper surface...like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar

  20. 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.

  1. Compressibility effects on dynamic stall of airfoils undergoing rapid transient pitching motion

    NASA Technical Reports Server (NTRS)

    Chandrasekhara, M. S.; Platzer, M. F.

    1992-01-01

    The research was carried out in the Compressible Dynamic Stall Facility, CDSF, at the Fluid Mechanics Laboratory (FML) of NASA Ames Research Center. The facility can produce realistic nondimensional pitch rates experienced by fighter aircraft, which on model scale could be as high as 3600/sec. Nonintrusive optical techniques were used for the measurements. The highlight of the effort was the development of a new real time interferometry method known as Point Diffraction Interferometry - PDI, for use in unsteady separated flows. This can yield instantaneous flow density information (and hence pressure distributions in isentropic flows) over the airfoil. A key finding is that the dynamic stall vortex forms just as the airfoil leading edge separation bubble opens-up. A major result is the observation and quantification of multiple shocks over the airfoil near the leading edge. A quantitative analysis of the PDI images shows that pitching airfoils produce larger suction peaks than steady airfoils at the same Mach number prior to stall. The peak suction level reached just before stall develops is the same at all unsteady rates and decreases with increase in Mach number. The suction is lost once the dynamic stall vortex or vortical structure begins to convect. Based on the knowledge gained from this preliminary analysis of the data, efforts to control dynamic stall were initiated. The focus of this work was to arrive at a dynamically changing leading edge shape that produces only 'acceptable' airfoil pressure distributions over a large angle of attack range.

  2. Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications

    NASA Technical Reports Server (NTRS)

    Somers, D. M.

    1981-01-01

    A flapped natural laminar flow airfoil for general aviation applications, the NLF(1)-0215F, has been designed and analyzed theoretically and verified experimentally in the Langley Low Turbulence Pressure Tunnel. The basic objective of combining the high maximum lift of the NASA low speed airfoils with the low cruise drag of the NACA 6 series airfoils has been achieved. The safety requirement that the maximum lift coefficient not be significantly affected with transition fixed near the leading edge has also been met. Comparisons of the theoretical and experimental results show generally good agreement.

  3. 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

  4. 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

  5. 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

  6. 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.

  7. Techniques for modifying airfoils and fairings on aircraft using foam and fiberglass

    NASA Technical Reports Server (NTRS)

    Meyer, M. B.; Jiran, F.

    1981-01-01

    The concept of using foam and fiberglass reinforced plastic to modify airfoils and fairings was applied successfully to high-speed aircraft at NASA Dryden Flight Research Center. An on-aircraft installation method was used to modify an F-15 wing glove and wing leading edge and an F-104 flap trailing edge in support of the Shuttle tile airload tests. A combination of methods, both an on-aircraft installation and an off-aircraft fabrication for installation on the aircraft, was used to modify a section of an F-111 supercritical wing with a natural laminar flow airfoil. Techniques, methods, problem areas, and recommendations are presented which indicate that using foam and fiberglass to modify airfoils and fairings on high-speed aircraft is a viable means of quickly developing airfoils and fairings with desired aerodynamic characteristics with little risk to the parent or carrier aircraft.

  8. 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.

  9. Interaction between leading and trailing edge vortex shedding: effects of bluff body geometry

    NASA Astrophysics Data System (ADS)

    Taylor, Zachary; Kopp, Gregory; Gurka, Roi

    2011-11-01

    Elongated bluff bodies are distinguished from shorter bluff bodies (e.g., circular cylinders) by the fact that they have separating-reattaching flow at the leading edge as well as having vortex shedding at the trailing edge. Engineering examples of these bodies include heat exchanger fins and long-span suspension bridges. We have performed experiments on elongated bluff bodies of varying geometry. These experiments have been performed at Reynolds numbers O(104) based on the thickness of the model. Both surface pressure measurements (using 512 simultaneously sampled pressure taps) and PIV are used to quantify the flow fields of these bodies. The leading edge separation angle is controlled by changing the leading edge geometry. It is observed that the size of the leading edge separation bubble increases with increasing leading edge separation angle. As the size of the leading edge separation bubble increases, it is shown to continually decrease the shedding frequency for a given elongation ratio. It is suggested that the shedding frequency is diminished because the trailing edge vortex shedding is affected by the structures being shed from the leading edge separation bubble. The implications of this competition between leading and trailing edge flows will be explored.

  10. 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.

  11. 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.

  12. Development of a cyber physical apparatus for investigating fluid structure interaction on leading edge vortex evolution

    NASA Astrophysics Data System (ADS)

    Raghu Gowda, Belagumba Venkatachalaiah

    This dissertation examines how simple structural compliance impacts a specific transient vortex phenomenon that occurs on high angle of attack lifting surfaces termed dynamic stall. In many Fluid structure interaction (FSI) research efforts, a purely physical or purely computational approach is taken. In this work a low cost cyber-physical (CPFD) system is designed and developed for representing the FSI in the leading edge vortex (LEV) development problem. The leading edge compliance appears to be favorable in a specific spring constant range for a given wing. When the leading edge compliance prescribed via CPFD system is too low compared with the moment due to dynamic pressure or fluid unsteady effect, the LEV behavior is similar to that of a rigid wing system. When the leading edge compliance is too high, excessive compliance is introduced into the wing system and the leading edge vortex evolution is affected by the large change in wing angle. At moderate leading edge compliance, a balance appears to be achieved in which the leading edge vorticity shedding rate supports the long term evolution of the leading edge vortex. Further investigation is required to determine specific parameters governing these leading edge compliance ranges.

  13. 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.

  14. 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.

  15. 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.

  16. 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.

  17. Modulation of leading edge vorticity and aerodynamic forces in flexible flapping wings.

    PubMed

    Zhao, Liang; Deng, Xinyan; Sane, Sanjay P

    2011-09-01

    In diverse biological flight systems, the leading edge vortex has been implicated as a flow feature of key importance in the generation of flight forces. Unlike fixed wings, flapping wings can translate at higher angles of attack without stalling because their leading edge vorticity is more stable than the corresponding fixed wing case. Hence, the leading edge vorticity has often been suggested as the primary determinant of the high forces generated by flapping wings. To test this hypothesis, it is necessary to modulate the size and strength of the leading edge vorticity independently of the gross kinematics while simultaneously monitoring the forces generated by the wing. In a recent study, we observed that forces generated by wings with flexible trailing margins showed a direct dependence on the flexural stiffness of the wing. Based on that study, we hypothesized that trailing edge flexion directly influences leading edge vorticity, and thereby the magnitude of aerodynamic forces on the flexible flapping wings. To test this hypothesis, we visualized the flows on wings of varying flexural stiffness using a custom 2D digital particle image velocimetry system, while simultaneously monitoring the magnitude of the aerodynamic forces. Our data show that as flexion decreases, the magnitude of the leading edge vorticity increases and enhances aerodynamic forces, thus confirming that the leading edge vortex is indeed a key feature for aerodynamic force generation in flapping flight. The data shown here thus support the hypothesis that camber influences instantaneous aerodynamic forces through modulation of the leading edge vorticity.

  18. Numerical study on reduction of aerodynamic noise around an airfoil with biomimetic structures

    NASA Astrophysics Data System (ADS)

    Wang, Jing; Zhang, Chengchun; Wu, Zhengyang; Wharton, James; Ren, Luquan

    2017-04-01

    A biomimetic airfoil featuring leading edge waves, trailing edge serrations and surface ridges is proposed in this study, based on flow control with each section meeting the NACA 0012 airfoil profile. Numerical simulations have been conducted to compare aerodynamic and acoustic performances between the NACA 0012 and biomimetic airfoils. These simulations utilize the large eddy simulation (LES) method and aeroacoustic analogy at an angle of attack of 0° and a Reynolds number of 1.0×105, based on using the airfoil chord as the characteristic length. The simulation results reveal the overall sound pressure levels (OASPLs) for all frequencies and at the seven observer points around the biomimetic airfoil, and a decrease of 13.1-13.9 dB is observed, whereas the drag coefficient is almost unchanged. The biomimetic structures can transform the shedding vortices in laminar mode for the NACA 0012 airfoil to regular horseshoe-type vortices in the wake, and reduce the spanwise correlation of the large-scale vortices, thereby restrain the vortex shedding noise around the biomimetic airfoil.

  19. 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

  20. Effects of leading-edge camber on low-speed characteristics of slender delta wings

    NASA Technical Reports Server (NTRS)

    Wentz, W. H., Jr.

    1972-01-01

    Wind-tunnel studies have been conducted to determine the effects of leading-edge camber on the low-speed aerodynamic characteristics of a thin, sharp-edge 74 deg delta wing. The results include force and moment measurements, pressure distributions, and flow visualization patterns determined from oil flow, tuft and water vapor observations. The study indicated that leading-edge camber near the apex is effective in controlling the pitch-up tendency of slender delta wings.

  1. Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 6. Perspectives Charter 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…

  2. 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.

  3. Airfoil for a gas turbine engine

    DOEpatents

    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.

  4. Analysis of airfoil transitional separation bubbles

    NASA Technical Reports Server (NTRS)

    Davis, R. L.; Carter, J. E.

    1984-01-01

    A previously developed local inviscid-viscous interaction technique for the analysis of airfoil transitional separation bubbles, ALESEP (Airfoil Leading Edge Separation) has been modified to utilize a more accurate windward finite difference procedure in the reversed flow region, and a natural transition/turbulence model has been incorporated for the prediction of transition within the separation bubble. Numerous calculations and experimental comparisons are presented to demonstrate the effects of the windward differencing scheme and the natural transition/turbulence model. Grid sensitivity and convergence capabilities of this inviscid-viscous interaction technique are briefly addressed. A major contribution of this report is that with the use of windward differencing, a second, counter-rotating eddy has been found to exist in the wall layer of the primary separation bubble.

  5. Leading Edge. Volume 7, Number 3. Systems Safety Engineering

    DTIC Science & Technology

    2010-01-01

    including both the MOGAS Stowage Room and MOGAS Transfer Room • Remove the external 1,500‑gallon bladder stowage rack and replace with modified low‑ sulfur...the 01 level on the port side deck edge. One rack holds six 55‑gallon drums and the other, a MOGAS stowage locker that is adjacent to the drum...requirements and standards for the safe handling, stowage , and use of all ammunition and explosives afloat. It is applicable to all ships owned or

  6. A method for computing the leading-edge suction in a higher-order panel method

    NASA Technical Reports Server (NTRS)

    Ehlers, F. E.; Manro, M. E.

    1984-01-01

    Experimental data show that the phenomenon of a separation induced leading edge vortex is influenced by the wing thickness and the shape of the leading edge. Both thickness and leading edge shape (rounded rather than point) delay the formation of a vortex. Existing computer programs used to predict the effect of a leading edge vortex do not include a procedure for determining whether or not a vortex actually exists. Studies under NASA Contract NAS1-15678 have shown that the vortex development can be predicted by using the relationship between the leading edge suction coefficient and the parabolic nose drag. The linear theory FLEXSTAB was used to calculate the leading edge suction coefficient. This report describes the development of a method for calculating leading edge suction using the capabilities of the higher order panel methods (exact boundary conditions). For a two dimensional case, numerical methods were developed using the double strength and downwash distribution along the chord. A Gaussian quadrature formula that directly incorporates the logarithmic singularity in the downwash distribution, at all panel edges, was found to be the best method.

  7. 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.

  8. Evaluation of Airfoil Dynamic Stall Characteristics for Maneuverability

    NASA Technical Reports Server (NTRS)

    Bousman, William G.; Aiken, Edwin W. (Technical Monitor)

    2000-01-01

    In severe maneuvers, out of necessity for a military aircraft or inadvertently for a civil aircraft, a helicopter airfoil will stall in a dynamic manner and provide lift beyond what would be calculated based on static airfoil tests. The augmented lift that occurs in dynamic stall is related to a vortex that is shed near the leading edge of the airfoil. However, directly related to the augmented lift that results from the dynamic stall vortex are significant penalties in pitching moment and drag. An understanding of the relationship between the augmented lift in dynamic stall and the associated moment and drag penalties is the purpose of this paper. This relationship is characterized using data obtained in two-dimensional wind tunnel tests and related to the problem of helicopter maneuverability.

  9. Study of supersonic wings employing the attainable leading-edge thrust concept

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.

    1982-01-01

    A theoretical study was made of supersonic wing geometries at Mach 1.8, using the attainable leading-edge thrust concept. The attainable thrust method offers a powerful means to improve overall aerodynamic efficiency by identifying wing leading-edge geometries that promote attached flow and by defining a local angle-of-attack range over which attached flow may be obtained. The concept applies to flat and to cambered wings, which leads to the consideration of drooped-wing leading edges for attached flow at high lift coefficients.

  10. Linearized propulsion theory of flapping airfoils revisited

    NASA Astrophysics Data System (ADS)

    Fernandez-Feria, R.

    2016-12-01

    A vortical impulse theory is used to compute the thrust force of a plunging and pitching airfoil in forward flight at high Reynolds numbers within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick, which considered only two effects, the leading-edge suction and the projection in the flight direction of the pressure force on the airfoil. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains, in addition to the pressure force projection term, a new term that generalizes the leading-edge suction term in Garrick's theory. This term depends on Theodorsen function C (k ) and on a new complex function C1(k ) of the reduced frequency k . The main qualitative difference with Garrick's theory is that the propulsive efficiency, or ratio of the mean thrust power and the mean input power required to drive the airfoil, tends to zero as the reduced frequency increases to infinity (as k-1), in contrast to Garrick's propulsive efficiency that tends to a constant (1 /2 ). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k →∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining nondimensional parameters. The present analytical results are in good agreement, for small amplitude oscillations, with numerical results from unsteady panel methods, and with experimental data and numerical results from the Navier-Stokes equations, except for small reduced frequencies where viscous effects are obviously important.

  11. Subsonic loads on wings having sharp leading edges and tips

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.; Mook, D. T.; Nayfeh, A. H.

    1976-01-01

    A vortex-lattice method for predicting the aerodynamics of wings having separation at the sharp edges in incompressible flows is extended to compressible subsonic flows using a modified Prandtl-Glauert transformation. Numerical results showing the effect of freestream Mach number on the aerodynamic coefficients are compared with available experimental data for several planforms. It is shown that the proposed method is suitable for predicting the aerodynamic loads on low-aspect wings at moderate angles of attack for high subsonic freestream Mach number. The method is limited to angles of attack up to 12 deg for high subsonic freestream Mach number and to angles of attack up to 20 deg for Mach number not exceeding 0.5.

  12. Generation of instability waves at a leading edge

    NASA Technical Reports Server (NTRS)

    Goldstein, M. E.

    1982-01-01

    Two cases are considered. The first is concerned with mean flows of the Blasius type wherein the instabilities are represented by Tollmien-Schlichting waves. It is shown that the latter are generated fairly far downstream of the edge and are the result of a wave length reduction process that tunes the free stream disturbances to the Tollmien-Schlichting wave length. The other case is concerned with inflectional, uni-directional, transversely sheared mean flows. Such idealized flows provide a fairly good local representation to the nearly parallel flows in jets. They can support inviscid instabilities of the Kelvin-Helmholtz type. The various mathematically permissible mechanisms that can couple these instabilities to the upstream disturbances are discussed.

  13. Characteristics of two sharp-nosed airfoils having reduced spinning tendencies

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1932-01-01

    According to Mr. L.D. Bell, of the Consolidated Aircraft Corporation, certain undesirable spinning characteristics of a commercial airplane were eliminated by the addition of a filler to the forward part of the wing to give it a sharp leading edge. To ascertain what aerodynamic effects result from such a change of section, two airfoils having sharp leading edges were tested in the variable-density wind tunnel. Both sections were derived by modifying the Gott. 398. The tests, which were made at a large value of the Reynolds Number, were carried to very large angles of attack to provide data for application to flight at angles of attack well beyond the stall. The characteristics of the sharp-nosed airfoils are compared with those of the normal Gott. 398 airfoil. Both of the sharp-nosed airfoils, which differ in the angle between the upper and lower surfaces at the leading edge, have about the same characteristics. As compared with the normal airfoil, the maximum lift is reduced by approximately 26 per cent, but the objectionable rapidly decreasing lift with angle of attack beyond the stall is eliminated; the profile drag of the section is slightly reduced in the range of the lift coefficient between 0.2 and 0.85, but at higher and lower lift coefficients the drag is increased.

  14. 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 .

  15. 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.

  16. 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.

  17. 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.

  18. 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.

  19. Compressibility and Leading-Edge Bluntness Effects for a 65 Deg Delta Wing

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2004-01-01

    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 a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the compressibility and bluntness effects primarily at a Reynolds number of 6 million from this data set. Emphasis is placed upon on the onset and progression of leading-edge vortex separation, and compressibility is shown to promote this separation. Comparisons with recent publications show that compressibility and Reynolds number have opposite effects on blunt leading edge vortex separation

  20. 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.

  1. Direct simulation Monte Carlo of rarefied hypersonic flow on power law shaped leading edges

    NASA Astrophysics Data System (ADS)

    Santos, Wilson Fernando Nogueira Dos

    A numerical study of several parameters that influence the flowfield structure, aerodynamic surface quantities and shock wave structure at rarefied hypersonic flow conditions is conducted on power law shaped leading edges. The calculations are performed with a detailed computer code that properly accounts for nonequilibrium effects and that has been demonstrated to yield excellent comparisons with flight- and ground-test data. The flowfield structure, aerodynamic surface quantities and shock wave structure of power law shaped leading edges are examined in order to provide information on how well these shapes could stand as possible candidates for blunting geometries of hypersonic leading edges. Newtonian flow analysis has shown that these shapes exhibit both blunt and sharp aerodynamic properties. Moreover, computational investigation of minimum-drag bodies at supersonic and moderate hypersonic speeds has indicated that power law shapes for certain exponents yield the lowest wave drag. These qualities make power law shapes strong candidates for leading edge design. A very detailed description of the impact on the flow properties, such as velocity, density, temperature and pressure, has been presented separately in the vicinity of the nose of the leading edges due to changes in their shapes. Numerical solutions show that the shape of the leading edge disturbed the flowfield far upstream, where the domain of influence decreased as the leading edge became aerodynamically sharp. A detailed procedure is presented to predict the pressure gradient along the body surface in a rarefied environment. Numerical solutions show that the pressure gradient behavior follows that predicted by Newtonian theory. It is found that the pressure gradient along the body surface goes to zero at the nose of the leading edge for power law exponents less than 2/3, a characteristic of a blunt body. It is finite for power law exponent of 2/3 and goes to minus infinite for power law exponents

  2. 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.

  3. 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.

  4. 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.

  5. A compressible solution of the Navier-Stokes equations for turbulent flow about an airfoil

    NASA Technical Reports Server (NTRS)

    Shamroth, S. J.; Gibeling, H. J.

    1979-01-01

    A compressible time dependent solution of the Navier-Stokes equations including a transition turbulence model is obtained for the isolated airfoil flow field problem. The equations are solved by a consistently split linearized block implicit scheme. A nonorthogonal body-fitted coordinate system is used which has maximum resolution near the airfoil surface and in the region of the airfoil leading edge. The transition turbulence model is based upon the turbulence kinetic energy equation and predicts regions of laminar, transitional, and turbulent flow. Mean flow field and turbulence field results are presented for an NACA 0012 airfoil at zero and nonzero incidence angles of Reynolds number up to one million and low subsonic Mach numbers.

  6. 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.

  7. 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.

  8. 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.

  9. Advanced leading edge thermal-structure concept. Direct bond reusable surface insulation to a composite structure

    NASA Technical Reports Server (NTRS)

    Riccitiello, S. R.; Figueroa, H.; Coe, C. F.; Kuo, C. P.

    1984-01-01

    An advanced leading-edge concept was analyzed using the space shuttle leading edge system as a reference model. The comparison indicates that a direct-bond system utilizing a high temperature (2700 F) fibrous refractory composite insulation tile bonded to a high temperature (PI/graphite) composite structure can result in a weight savings of up to 800 lb. The concern that tile damage or loss during ascent would result in adverse entry aerodynamics if a leading edge tile system were used is addressed. It was found from experiment that missing tiles (as many as 22) on the leading edge would not significantly affect the basic force-and-moment aerodynamic coefficients. Additionally, this concept affords a degree of redundancy to a thermal protection system in that the base structure (being a composite material) ablates and neither melts nor burns through when subjected to entry heating in the event tiles are actually lost or damaged during ascent.

  10. 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.

  11. FORTRAN program for calculating leading and trailing-edge geometry of turbomachine blades

    NASA Technical Reports Server (NTRS)

    Schumann, L. F.

    1977-01-01

    A FORTRAN IV program which calculates leading- and trailing-edge circle radii, tangency angles on the leading- and trailing-edge circles, and stagger angles of turbomachinery blade sections using only spline points defining the blade surfaces is described. The program shifts the origin of the blade coordinates to the leading edge of the blade. Required input includes (m, theta) coordinates of a sufficient number of spline points to adequately define the two surfaces of the blade. Other required input are the radii from the axis of rotation of the leading- and trailing-edges. The output from this program is used directly as the geometrical input for a NASA developed program for calculating transonic velocities on a blade-to-blade stream surface of a turbomachine. The program is used for axial, radial, and mixed flow turbomachine blades.

  12. Two leading-edge droop modifications for tailoring stall characteristics of a general aviation trainer configuration

    NASA Technical Reports Server (NTRS)

    Ross, Holly M.; Perkins, John N.

    1992-01-01

    The high-angle-of-attack testing intended to develop leading-edge modifications for tailoring the stall characteristics of model is described. Two different leading-edge modifications are considered: a small profile leading-edge droop on the outboard 24 percent of the wing and a large profile leading-edge droop on the outboard 50 percent of the wing. Results indicate that the longitudinal stability for the unmodified and both modified configurations was good for low angle of attack, but the modified configurations exhibited neutral longitudinal stability just prior to stall. The unmodified and both modified configurations demonstrated good lateral stability characteristics for low angles of attack, but all configurations were directionally unstable for high angles of attack.

  13. 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.

  14. Aerodynamics of a Flapping Airfoil with a Flexible Tail

    NASA Astrophysics Data System (ADS)

    Lai, Alan Kai San

    This dissertation presents computational solutions to an airfoil in a oscillatory heaving motion with a aeroelastically flexible tail attachment. An unsteady potential flow solver is coupled to a structural solver to obtain the aeroelastic flow solution over an inviscid fluid to investigate the propulsive performance of such a configuration. The simulation is then extended to a two-dimensional viscous solver by coupling NASA's CFL3D solver to the structural solver to study how the flow is altered by the presence of viscosity. Finally, additional simulations are done in three dimensions over wings with varying aspect ratio to study the three-dimensional effects on the propulsive performance of an airfoil with an aeroelastic tail. The computation reveals that the addition of the aeroelastic trailing edge improved the thrust generated by a heaving airfoil significantly. As the frequency of the heaving motion increases, the thrust generated by the airfoil with the tail increases exponentially. In an inviscid fluid, the increase in thrust is insufficient to overcome the increase in power required to maintain the motion and as a result the overall propulsive efficiency is reduced. When the airfoil is heaving in a viscous fluid, the presence of a suction boundary layer and the appearance of leading edge vortex increase the thrust generated to such an extent that the propulsive efficiency is increased by about 3% when compared to the same airfoil with a rigid tail. The three-dimensional computations shows that the presence of the tip vorticies suppress some of the increase in thrust observed in the two-dimensional viscous computations for short span wings. For large span wings, the overall thrust enhancing capabilities of the aeroelastic tail is preserved.

  15. Extension of a vortex-lattice method to include the effects of leading-edge separation

    NASA Technical Reports Server (NTRS)

    Mook, D. T.; Maddox, S. A.

    1974-01-01

    Vortex-lattice methods have been used successfully to obtain the aerodynamic coefficients of lifting surfaces without leading-edge separation. It is shown how an existing vortex-lattice method can be modified to include the effects of leading-edge separation. The modified version is then used to calculate the aerodynamic loads on a highly swept delta wing. The results are compared with Peckham's (1958) experimental data.

  16. 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.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. 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.

  2. 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.

  3. Extension of leading-edge-suction analogy to wings with separated flow around the side edges at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Lamar, J. E.

    1974-01-01

    A method for determining the lift, drag, and pitching moment for wings which have separated flow at the leading and side edges with subsequently reattached flow downstream and inboard is presented. Limiting values of the contribution to lift of the side-edge reattached flow are determined for rectangular wings. The general behavior of this contribution is computed for rectangular, cropped-delta, cropped-diamond, and cropped-arrow wings. Comparisons of the results of the method and experiment indicate reasonably good correlation of the lift, drag, and pitching moment for a wide planform range. The agreement of the method with experiment was as good as, or better than, that obtained by other methods. The procedure is computerized and is available from COSMIC as NASA Langley computer program A0313.

  4. Ordered roughness effects on NACA 0026 airfoil

    NASA Astrophysics Data System (ADS)

    Harun, Z.; Abbas, A. A.; Dheyaa, R. Mohammed; Ghazali, M. I.

    2016-10-01

    The effects of highly-ordered rough surface - riblets, applied onto the surface of a NACA 0026 airfoil, are investigated experimentally using wind tunnel. The riblets are arranged in directionally converging - diverging pattern with dimensions of height, h = 1 mm, pitch or spacing, s = 1 mm, yaw angle α = 0o and 10o The airfoil with external geometry of 500 mm span, 600 mm chord and 156 mm thickness has been built using mostly woods and aluminium. Turbulence quantities are collected using hotwire anemometry. Hotwire measurements show that flows past converging and diverging pattern inherit similar patterns in the near-wall region for both mean velocity and turbulence intensities profiles. The mean velocity profiles in logarithmic regions for both flows past converging and diverging riblet pattern are lower than that with yaw angle α = 0o. Converging riblets cause the boundary layer to thicken and the flow with yaw angle α = 0o produces the thinnest boundary layer. Both the converging and diverging riblets cause pronounced outer peaks in the turbulence intensities profiles. Most importantly, flows past converging and diverging pattern experience 30% skin friction reductions. Higher order statistics show that riblet surfaces produce similar effects due to adverse pressure gradient. It is concluded that a small strip of different ordered roughness features applied at a leading edge of an airfoil can change the turbulence characteristics dramatically.

  5. 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.

  6. 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

  7. Investigation of the Effects of Leading-edge Chord-extensions and Fences in Combination with Leading-edge Flaps on the Aerodynamic Characteristics at Mach Numbers from 0.40 to 0.93 of a 45 Degree Sweptback Wing of Aspect Ratio 4

    NASA Technical Reports Server (NTRS)

    Spreeman, Kenneth P; Alford, William J , Jr

    1954-01-01

    This investigation was made to determine the effects of 6 degree full-spoan and 3 degree partial-span leading-edge flaps in combination with chord-extensions or fences on the aerodynamic characteristics of a wing-fuselage configuration with a 45 degree sweptback wing of aspect ratio 4, taper ratio 0.3, and NACA 65A006 airfoil sections. The investigation was made in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.40 to 0.93 and an angle-of-attack range of about -2 degrees to 24 degrees. Lift, drag, and pitching-moment data were obtained for all configurations. From overall considerations of stability and performance it appears that with the model of this investigation the 6 degree full-span leading-edge flaps in combination with the chord-extension over the outboard 35 percent of the span, with or without leading-edge camber, would be the most desirable configuration.

  8. Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes

    NASA Technical Reports Server (NTRS)

    Ivey, Margaret F

    1945-01-01

    Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corresponding cutouts made in wing were tested for various flap deflections, chord-wise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80 percent wing chord behind leading edge. Airflow was smooth and buffeting negligible.

  9. 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.

  10. 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.

  11. 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.

  12. GOLPH3 drives cell migration by promoting Golgi reorientation and directional trafficking to the leading edge

    PubMed Central

    Xing, Mengke; Peterman, Marshall C.; Davis, Robert L.; Oegema, Karen; Shiau, Andrew K.; Field, Seth J.

    2016-01-01

    The mechanism of directional cell migration remains an important problem, with relevance to cancer invasion and metastasis. GOLPH3 is a common oncogenic driver of human cancers, and is the first oncogene that functions at the Golgi in trafficking to the plasma membrane. Overexpression of GOLPH3 is reported to drive enhanced cell migration. Here we show that the phosphatidylinositol-4-phosphate/GOLPH3/myosin 18A/F-actin pathway that is critical for Golgi–to–plasma membrane trafficking is necessary and limiting for directional cell migration. By linking the Golgi to the actin cytoskeleton, GOLPH3 promotes reorientation of the Golgi toward the leading edge. GOLPH3 also promotes reorientation of lysosomes (but not other organelles) toward the leading edge. However, lysosome function is dispensable for migration and the GOLPH3 dependence of lysosome movement is indirect, via GOLPH3’s effect on the Golgi. By driving reorientation of the Golgi to the leading edge and driving forward trafficking, particularly to the leading edge, overexpression of GOLPH3 drives trafficking to the leading edge of the cell, which is functionally important for directional cell migration. Our identification of a novel pathway for Golgi reorientation controlled by GOLPH3 provides new insight into the mechanism of directional cell migration with important implications for understanding GOLPH3’s role in cancer. PMID:27708138

  13. 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.

  14. 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.

  15. 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

  16. Linearized propulsion theory of flapping airfoils revisited

    NASA Astrophysics Data System (ADS)

    Fernandez-Feria, Ramon

    2016-11-01

    A vortical impulse theory is used to compute the thrust of a plunging and pitching airfoil in forward flight within the framework of linear potential flow theory. The result is significantly different from the classical one of Garrick that considered the leading-edge suction and the projection in the flight direction of the pressure force. By taking into account the complete vorticity distribution on the airfoil and the wake the mean thrust coefficient contains a new term that generalizes the leading-edge suction term and depends on Theodorsen function C (k) and on a new complex function C1 (k) of the reduced frequency k. The main qualitative difference with Garrick's theory is that the propulsive efficiency tends to zero as the reduced frequency increases to infinity (as 1 / k), in contrast to Garrick's efficiency that tends to a constant (1 / 2). Consequently, for pure pitching and combined pitching and plunging motions, the maximum of the propulsive efficiency is not reached as k -> ∞ like in Garrick's theory, but at a finite value of the reduced frequency that depends on the remaining non-dimensional parameters. The present analytical results are in good agreement with experimental data and numerical results for small amplitude oscillations. Supported by the Ministerio de Economia y Competitividad of Spain Grant No. DPI2013-40479-P.

  17. Computer Program to Obtain Ordinates for NACA Airfoils

    NASA Technical Reports Server (NTRS)

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

    1996-01-01

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

  18. Characterizing a burst leading-edge vortex on a rotating flat plate wing

    NASA Astrophysics Data System (ADS)

    Jones, Anya R.; Medina, Albert; Spooner, Hannah; Mulleners, Karen

    2016-04-01

    Identifying, characterizing, and tracking incoherent vortices in highly separated flows is of interest for the development of new low-order models for unsteady lift prediction. The current work examines several methods to identify vortex burst and characterize a burst leading-edge vortex. Time-resolved stereoscopic PIV was performed on a rotating flat plate wing at Re = 2500. The burst process was found to occur at mid-span and is characterized by axial flow reversal, the entrainment of opposite-sign vorticity, and a rapid expansion of vortex size. A POD analysis revealed that variations in certain mode coefficients are indicative of the flow state changes characteristics of burst. During burst, the leading-edge vortex evolves to a region of inhomogeneous vorticity distributed over a large area. Several methods of defining the vortex size and circulation are evaluated and a combination of these can be used to characterize the leading-edge vortex both pre- and post-burst.

  19. 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.

  20. 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.

  1. Influence of blade angle distribution along leading edge on cavitation performance of a centrifugal pump

    NASA Astrophysics Data System (ADS)

    Xu, Y.; Tan, L.; Cao, S. L.; Wang, Y. C.; Meng, G.; Qu, W. S.

    2015-01-01

    The influence of blade angle distribution along leading edge on cavitation performance of centrifugal pumps is analysed in the present paper. Three sets of blade angle distribution along leading edge for three blade inlet angles are chosen to design nine centrifugal pump impellers. The RNG k-epsilon turbulence model and the Zwart-Gerber-Belamri cavitation model are employed to simulate the cavitation flows in centrifugal pumps with different impellers and the same volute. The numerical results are compared with the experimental data, and the comparison proves that the numerical simulation can accurately predict the cavitation performance of centrifugal pumps. On the basis of the numerical simulations, the pump head variations with pump inlet pressure, and the flow details in centrifugal pump are revealed to demonstrate the influence of blade angle distribution along leading edge on cavitation performances of centrifugal pumps.

  2. 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.

  3. Application of local indentations for film cooling of gas turbine blade leading edge

    NASA Astrophysics Data System (ADS)

    Petelchyts, V. Yu.; Khalatov, A. A.; Pysmennyi, D. N.; Dashevskyy, Yu. Ya.

    2016-09-01

    The paper presents results of computer simulation of the film cooling on the turbine blade leading edge model where the air coolant is supplied through radial holes and row of cylindrical inclined holes placed inside hemispherical dimples or trench. The blowing factor was varied from 0.5 to 2.0. The model size and key initial parameters for simulation were taken as for a real blade of a high-pressure high-performance gas turbine. Simulation was performed using commercial software code ANSYS CFX. The simulation results were compared with reference variant (no dimples or trench) both for the leading edge area and for the flat plate downstream of the leading edge.

  4. c-Jun is essential for organization of the epidermal leading edge.

    PubMed

    Li, Guochun; Gustafson-Brown, Cindy; Hanks, Steven K; Nason, Katie; Arbeit, Jeffrey M; Pogliano, Kit; Wisdom, Ronald M; Johnson, Randall S

    2003-06-01

    The migration of epithelial layers requires specific and coordinated organization of the cells at the leading edge of the sheet. Mice that are conditionally deleted for the c-jun protooncogene in epidermis are born at expected frequencies, but with open eyes and with defects in epidermal wound healing. Keratinocytes lacking c-Jun are unable to migrate or elongate properly in culture at the border of scratch assays. Histological analyses in vitro and in vivo demonstrate an inability to activate EGF receptor at the leading edge of wounds, and we demonstrate that this can be rescued by supplementation with conditioned medium or the EGF receptor ligand HB-EGF. Lack of c-Jun prevents EGF-induced expression of HB-EGF, indicating that c-jun controls formation of the epidermal leading edge through its control of an EGF receptor autocrine loop.

  5. A feasibility study of heat-pipe-cooled leading edges for hypersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    Silverstein, C. C.

    1971-01-01

    A theoretical study of the use of heat pipe structures for cooling the leading edges of hypersonic cruise aircraft was carried out over a Mach number range of 6 to 12. Preliminary design studies showed that a heat pipe cooling structure with a 33-in. chordwise length could maintain the maximum temperature of a 65 deg sweepback wing with a 0.5-in. leading edge radius below 1600 F during cruise at Mach 8. A few relatively minor changes in the steady-state design of the structure were found necessary to insure satisfactory cooling during the climb to cruise speed and altitude. It was concluded that heat pipe cooling is an attractive, feasible technique for limiting leading edge temperatures of hypersonic cruise aircraft.

  6. 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.

  7. 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.

  8. 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.

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

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    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 a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.

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

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    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 a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

  11. 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.

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

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    2003-01-01

    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 a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.

  13. Thermoviscoplastic analysis of engine cowl leading edge subjected to oscillating shock-shock interaction

    NASA Technical Reports Server (NTRS)

    Pandey, Ajay K.

    1992-01-01

    A finite element thermoviscoplastic analysis method, which employs a unified constitutive model proposed by Bodner and Partom, is used to predict rate-dependent nonlinear structural behavior. The method is evaluated by predicting stress-strain behavior of a uniaxially loaded bar of nickel-based superalloy (B1900 + Hf) material. The method is used to predict the time-dependent thermoviscoplastic response of a B1900 + Hf leading edge subjected to oscillating shock-shock interaction loading. Viscoplastic analysis shows that the leading edge experiences significant plastic straining. The plastic region increases with cyclic loading in the high heat flux area.

  14. 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.

  15. 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.

  16. Development flight tests of JetStar LFC leading-edge flight test experiment

    NASA Technical Reports Server (NTRS)

    Fisher, David F.; Fischer, Michael C.

    1987-01-01

    The overall objective of the flight tests on the JetStar aircraft was to demonstrate the effectiveness and reliability of laminar flow control under representative flight conditions. One specific objective was to obtain laminar flow on the JetStar leading-edge test articles for the design and off-design conditions. Another specific objective was to obtain operational experience on a Laminar Flow Control (LFC) leading-edge system in a simulated airline service. This included operational experience with cleaning requirements, the effect of clogging, possible foreign object damage, erosion, and the effects of ice particle and cloud encounters. Results are summarized.

  17. Determination of forced convective heat transfer coefficients for subsonic flows over heated asymmetric NANA 4412 airfoil

    NASA Astrophysics Data System (ADS)

    Dag, Yusuf

    Forced convection over traditional surfaces such as flat plate, cylinder and sphere have been well researched and documented. Data on forced convection over airfoil surfaces, however, remain very scanty in literature. High altitude vehicles that employ airfoils as lifting surfaces often suffer leading edge ice accretions which have tremendous negative consequences on the lifting capabilities and stability of the vehicle. One of the ways of mitigating the effect of ice accretion involves judicious leading edge convective cooling technique which in turn depends on the accuracy of convective heat transfer coefficient used in the analysis. In this study empirical investigation of convective heat transfer measurements on asymmetric airfoil is presented at different angle of attacks ranging from 0° to 20° under subsonic flow regime. The top and bottom surface temperatures are measured at given points using Senflex hot film sensors (Tao System Inc.) and used to determine heat transfer characteristics of the airfoils. The model surfaces are subjected to constant heat fluxes using KP Kapton flexible heating pads. The monitored temperature data are then utilized to determine the heat convection coefficients modelled empirically as the Nusselt Number on the surface of the airfoil. The experimental work is conducted in an open circuit-Eiffel type wind tunnel, powered by a 37 kW electrical motor that is able to generate subsonic air velocities up to around 41 m/s in the 24 square-inch test section. The heat transfer experiments have been carried out under constant heat flux supply to the asymmetric airfoil. The convective heat transfer coefficients are determined from measured surface temperature and free stream temperature and investigated in the form of Nusselt number. The variation of Nusselt number is shown with Reynolds number at various angles of attacks. It is concluded that Nusselt number increases with increasing Reynolds number and increase in angle of attack from 0

  18. Effect of leading-edge geometry on boundary-layer receptivity to freestream sound

    NASA Technical Reports Server (NTRS)

    Lin, Nay; Reed, Helen L.; Saric, W. S.

    1991-01-01

    The receptivity to freestream sound of the laminar boundary layer over a semi-infinite flat plate with an elliptic leading edge is simulated numerically. The incompressible flow past the flat plate is computed by solving the full Navier-Stokes equations in general curvilinear coordinates. A finite-difference method which is second-order accurate in space and time is used. Spatial and temporal developments of the Tollmien-Schlichting wave in the boundary layer, due to small-amplitude time-harmonic oscillations of the freestream velocity that closely simulate a sound wave travelling parallel to the plate, are observed. The effect of leading-edge curvature is studied by varying the aspect ratio of the ellipse. The boundary layer over the flat plate with a sharper leading edge is found to be less receptive. The relative contribution of the discontinuity in curvature at the ellipse-flat-plate juncture to receptivity is investigated by smoothing the juncture with a polynomial. Continuous curvature leads to less receptivity. A new geometry of the leading edge, a modified super ellipse, which provides continuous curvature at the juncture with the flat plate, is used to study the effect of continuous curvature and inherent pressure gradient on receptivity.

  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. 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.

  2. 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.

  3. Composite airfoil assembly

    DOEpatents

    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.

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

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

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

  5. 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.

    1988-01-01

    Results of preliminary tests to measure ice growth on an airfoil during flight icing conditions are presented. Ultrasonic pulse echo measurements of ice thickness are obtained from an array of eight ultrasonic transducers mounted flush with the leading edge of the airfoil. These thickness measurements are used to document the evolution of the ice shape during the encounter in the form of successive ice profiles. Results from 3 research flights are presented and discussed. The accuracy of the ultrasonic measurements is found to be within 0.5 mm of mechanical and stereo photograph measurements of the ice accretion.

  6. Upper-surface modifications for C sub l max improvement of selected NASA 6-series airfoils

    NASA Technical Reports Server (NTRS)

    Szelazek, C. A.; Hicks, R. M.

    1979-01-01

    The thickness of the upper surface of 64 airfoils was increased from the leading edge to the position of maximum thickness. The modifications were generated using a numerical optimization routine coupled with an aerodynamic analysis code. The type of modification presented can be used for aircraft design or for the retrofit of current aircraft to improve the stall characteristics and climb performance. The coordinates of the modified airfoils are presented with plots of the forward 45% of the profiles and pressure distributions for both the modified and unmodified sections at an angle of attack of 14 degrees.

  7. The Leading Edge of Early Childhood Education: Linking Science to Policy for a New Generation

    ERIC Educational Resources Information Center

    Lesaux, Nonie K., Ed.; Jones, Stephanie M., Ed.

    2016-01-01

    "The Leading Edge of Early Childhood Education" aims to support the effort to simultaneously scale up and improve the quality of early childhood education by bringing together relevant insights from emerging research to provide guidance for this critical, fledgling field. It reflects the growing recognition that early childhood…

  8. 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.

  9. 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…

  10. 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.

  11. Guidance signalling regulates leading edge behaviour during collective cell migration of cardiac cells in Drosophila.

    PubMed

    Raza, Qanber; Jacobs, J Roger

    2016-11-15

    Collective cell migration is the coordinated movement of cells, which organize tissues during morphogenesis, repair and some cancers. The motile cell membrane of the advancing front in collective cell migration is termed the Leading Edge. The embryonic development of the vertebrate and Drosophila hearts are both characterized by the coordinated medial migration of a bilateral cluster of mesodermal cells. In Drosophila, the cardioblasts form cohesive bilateral rows that migrate collectively as a unit towards the dorsal midline to form the dorsal vessel. We have characterized the collective cell migration of cardioblasts as an in vivo quantitative model to study the behaviour of the Leading Edge. We investigated whether guidance signalling through Slit and Netrin pathways plays a role in cell migration during heart development. Through time-lapse imaging and quantitative assessment of migratory behaviour of the cardioblasts in loss-of-function mutants, we demonstrate that both Slit and Netrin mediated signals are autonomously and concomitantly required to maximize migration velocity, filopodial and lamellipodial activities. Additionally, we show that another Slit and Netrin receptor, Dscam1, the role of which during heart development was previously unknown, is required for both normal migration of cardioblasts and luminal expansion. Leading edge behaviour analysis revealed a dosage dependent genetic interaction between Slit and Netrin receptors suggesting that downstream signalling through these receptors converge on a common output that increases leading edge activity of the cardioblasts. Finally, we found that guidance signalling maintains the balance between epithelial and mesenchymal characteristics of the migrating cardioblasts.

  12. Evaluation of a sodium/Hastelloy-X heat pipe for wing leading edge cooling

    SciTech Connect

    Merrigan, M.A.; Sena, J.T.; Glass, D.E.

    1996-12-31

    This report covers assembly of a sodium heat pipe, testing to verify performance during start-up and under steady-state conditions with stagnation point heat loads to about 80 W/cm{sup 2}, performance analysis and evaluation. Evaluation of this leading edge cooling concept is offered and recommendations for further research discussed.

  13. 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…

  14. Weak Force Stalls Protrusion at the Leading Edge of the Lamellipodium

    PubMed Central

    Bohnet, Sophie; Ananthakrishnan, Revathi; Mogilner, Alex; Meister, Jean-Jacques; Verkhovsky, Alexander B.

    2006-01-01

    Protrusion, the first step of cell migration, is driven by actin polymerization coupled to adhesion at the cell's leading edge. Polymerization and adhesive forces have been estimated, but the net protrusion force has not been measured accurately. We arrest the leading edge of a moving fish keratocyte with a hydrodynamic load generated by a fluid flow from a micropipette. The flow arrests protrusion locally as the cell approaches the pipette, causing an arc-shaped indentation and upward folding of the leading edge. The effect of the flow is reversible upon pipette removal and dependent on the flow direction, suggesting that it is a direct effect of the external force rather than a regulated cellular response. Modeling of the fluid flow gives a surprisingly low value for the arresting force of just a few piconewtons per micrometer. Enhanced phase contrast, fluorescence, and interference reflection microscopy suggest that the flow does not abolish actin polymerization and does not disrupt the adhesions formed before the arrest but rather interferes with weak nascent adhesions at the very front of the cell. We conclude that a weak external force is sufficient to reorient the growing actin network at the leading edge and to stall the protrusion. PMID:16326894

  15. 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,…

  16. 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.

  17. 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.

  18. Decomposing the aerodynamic forces of low-Reynolds flapping airfoils

    NASA Astrophysics Data System (ADS)

    Moriche, Manuel; Garcia-Villalba, Manuel; Flores, Oscar

    2016-11-01

    We present direct numerical simulations of flow around flapping NACA0012 airfoils at relatively small Reynolds numbers, Re = 1000 . The simulations are carried out with TUCAN, an in-house code that solves the Navier-Stokes equations for an incompressible flow with an immersed boundary method to model the presence of the airfoil. The motion of the airfoil is composed of a vertical translation, heaving, and a rotation about the quarter of the chord, pitching. Both motions are prescribed by sinusoidal laws, with a reduced frequency of k = 1 . 41 , a pitching amplitude of 30deg and a heaving amplitude of one chord. Both, the mean pitch angle and the phase shift between pitching and heaving motions are varied, to build a database with 18 configurations. Four of these cases are analysed in detail using the force decomposition algorithm of Chang (1992) and Martín Alcántara et al. (2015). This method decomposes the total aerodynamic force into added-mass (translation and rotation of the airfoil), a volumetric contribution from the vorticity (circulatory effects) and a surface contribution proportional to viscosity. In particular we will focus on the second, analysing the contribution of the leading and trailing edge vortices that typically appear in these flows. This work has been supported by the Spanish MINECO under Grant TRA2013-41103-P. The authors thankfully acknowledge the computer resources provided by the Red Española de Supercomputacion.

  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. Methodology for heat flux investigation on leading edges using infrared thermography

    NASA Astrophysics Data System (ADS)

    Corre, Y.; Gardarein, J.-L.; Dejarnac, R.; Gaspar, J.; Gunn, J. P.; Aumeunier, M.-H.; Courtois, X.; Missirlian, M.; Rigollet, F.

    2017-01-01

    During steady state plasma operation in fusion devices, leading edges of the actively cooled plasma-facing components can be impacted by plasma flux with nearly normal angle of incidence, causing local overheating. The overheating can be a critical issue in high-power machines, especially in the presence of mechanical misalignments. Due to heat diffusion through the material, the edge power overload also leads to a local increase of temperature on the top part of the tile that can be detected by the infrared imaging system (viewed from the top of the machine). In the Tore Supra tokamak, heat flux impinging on the top and the leading edge of the carbon fibre composite (CFC) flat tiles are characterized with both an infrared (IR) thermographic system and 2D thermal modelling of the tile. A specific sensor correction based on a laboratory blackbody-slit experiment has been developed to simulate the spatial resolution related effects (necessary here since the temperature gradient near the leading edge is smaller than the pixel size of the IR system). The transfer function of the IR system is characterized by a Gaussian distribution function. The standard deviation is found to be σ  =  1.75 mm for a pixel size of 3.1 mm. The heat flux calculation is applied to CFC flat tiles and, after being processed with the transfer function, compared to experimental IR data for two geometrical situations: one with 0.2 mm misalignment between two adjacent tiles and the other without misalignment (well-aligned tiles). The heat flux ratio between the leading edge and top is found to be ~25 in the case of the protruding tile, which is lower than the expected ratio using the guiding-centre ballistic approximation with no cross-field heat flux (57).

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

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  3. Spanwise flow and the attachment of the leading-edge vortex on insect wings.

    PubMed

    Birch, J M; Dickinson, M H

    2001-08-16

    The flow structure that is largely responsible for the good performance of insect wings has recently been identified as a leading-edge vortex. But because such vortices become detached from a wing in two-dimensional flow, an unknown mechanism must keep them attached to (three-dimensional) flapping wings. The current explanation, analogous to a mechanism operating on delta-wing aircraft, is that spanwise flow through a spiral vortex drains energy from the vortex core. We have tested this hypothesis by systematically mapping the flow generated by a dynamically scaled model insect while simultaneously measuring the resulting aerodynamic forces. Here we report that, at the Reynolds numbers matching the flows relevant for most insects, flapping wings do not generate a spiral vortex akin to that produced by delta-wing aircraft. We also find that limiting spanwise flow with fences and edge baffles does not cause detachment of the leading-edge vortex. The data support an alternative hypothesis-that downward flow induced by tip vortices limits the growth of the leading-edge vortex.

  4. 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.

  5. 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

  6. Low speed airfoil design and analysis

    NASA Technical Reports Server (NTRS)

    Eppler, R.; Somers, D. M.

    1979-01-01

    A low speed airfoil design and analysis program was developed which contains several unique features. In the design mode, the velocity distribution is not specified for one but many different angles of attack. Several iteration options are included which allow the trailing edge angle to be specified while other parameters are iterated. For airfoil analysis, a panel method is available which uses third-order panels having parabolic vorticity distributions. The flow condition is satisfied at the end points of the panels. Both sharp and blunt trailing edges can be analyzed. The integral boundary layer method with its laminar separation bubble analog, empirical transition criterion, and precise turbulent boundary layer equations compares very favorably with other methods, both integral and finite difference. Comparisons with experiment for several airfoils over a very wide Reynolds number range are discussed. Applications to high lift airfoil design are also demonstrated.

  7. 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.

  8. 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.

  9. Thermostructural applications of heat pipes for cooling leading edges of high-speed aerospace vehicles

    NASA Technical Reports Server (NTRS)

    Camarda, Charles J.; Glass, David E.

    1992-01-01

    Heat pipes have been considered for use on wing leading edge for over 20 years. Early concepts envisioned metal heat pipes cooling a metallic leading edge. Several superalloy/sodium heat pipes were fabricated and successfully tested for wing leading edge cooling. Results of radiant heat and aerothermal testing indicate the feasibility of using heat pipes to cool the stagnation region of shuttle-type space transportation systems. The test model withstood a total seven radiant heating tests, eight aerothermal tests, and twenty-seven supplemental radiant heating tests. Cold-wall heating rates ranged from 21 to 57 Btu/sq ft-s and maximum operating temperatures ranged from 1090 to 1520 F. Follow-on studies investigated the application of heat pipes to cool the stagnation regions of single-stage-to-orbit and advanced shuttle vehicles. Results of those studies indicate that a 'D-shaped' structural design can reduce the mass of the heat-pipe concept by over 44 percent compared to a circular heat-pipe geometry. Simple analytical models for heat-pipe startup from the frozen state (working fluid initially frozen) were adequate to approximate transient, startup, and steady-state heat-pipe performance. Improvement in analysis methods has resulted in the development of a finite-element analysis technique to predict heat-pipe startup from the frozen state. However, current requirements of light-weight design and reliability suggest that metallic heat pipes embedded in a refractory composite material should be used. This concept is the concept presently being evaluated for NASP. A refractory-composite/heat-pipe-cooled wing leading edge is currently being considered for the National Aero-Space Plane (NASP). This concept uses high-temperature refractory-metal/lithium heat pipes embedded within a refractory-composite structure and is significantly lighter than an actively cooled wing leading edge because it eliminates the need for active cooling during ascent and descent. Since the

  10. Transonic airfoil design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1976-01-01

    A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

  11. 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.

  12. 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.

  13. Navier-Stokes computation of wing leading edge tangential blowing for a tilt rotor in hover

    NASA Technical Reports Server (NTRS)

    Fejtek, Ian; Roberts, Leonard

    1992-01-01

    The effect of a thin tangential jet located at the leading edge of the wing of a tilt rotor configuration in hover is computed using the thin-layer Navier-stokes equations. Computations showed that leading edge tangential blowing is effective in reducing the download caused by the impingement of the rotor download caused by the impingement of the rotor downwash on the wing. Results from the numerical model support previous experimental findings that download reduction is due mainly to a decrease in upper surface pressure and not an increase in pressure on the lower surface. The numerical solution clearly shows that because of the three-dimensionality of the flow field, the download could be reduced further by allowing a spanwise variation in blowing strength.

  14. 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.

  15. 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

  16. 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.

  17. 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.

  18. AMPK activity regulates trafficking of mitochondria to the leading edge during cell migration and matrix invasion

    PubMed Central

    Cunniff, Brian; McKenzie, Andrew J.; Heintz, Nicholas H.; Howe, Alan K.

    2016-01-01

    Cell migration is a complex behavior involving many energy-expensive biochemical events that iteratively alter cell shape and location. Mitochondria, the principal producers of cellular ATP, are dynamic organelles that fuse, divide, and relocate to respond to cellular metabolic demands. Using ovarian cancer cells as a model, we show that mitochondria actively infiltrate leading edge lamellipodia, thereby increasing local mitochondrial mass and relative ATP concentration and supporting a localized reversal of the Warburg shift toward aerobic glycolysis. This correlates with increased pseudopodial activity of the AMP-activated protein kinase (AMPK), a critically important cellular energy sensor and metabolic regulator. Furthermore, localized pharmacological activation of AMPK increases leading edge mitochondrial flux, ATP content, and cytoskeletal dynamics, whereas optogenetic inhibition of AMPK halts mitochondrial trafficking during both migration and the invasion of three-dimensional extracellular matrix. These observations indicate that AMPK couples local energy demands to subcellular targeting of mitochondria during cell migration and invasion. PMID:27385336

  19. 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.

  20. 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.

  1. 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.

  2. Experimental Investigation of a Yawed Airfoil in Reverse Flow Dynamic Stall

    NASA Astrophysics Data System (ADS)

    Smith, Luke; Lind, Andrew, , Dr.; Jones, Anya, , Dr.

    2016-11-01

    When a rotating blade enters high advance ratio flight, a significant portion of the blade is subject to reverse flow, where flow travels from the blade's geometric trailing edge to the geometric leading edge. The purpose of this work is to determine the influence of spanwise flow on a blade undergoing dynamic stall in reverse flow. Without spanwise flow, an oscillating sharp trailing edge airfoil in reverse flow experiences separation about its sharp aerodynamic leading edge, leading to the formation of a dynamic stall vortex at low angles of attack. With spanwise flow, an airfoil experiences a delay in lift stall, possibly due to the convection of a vortex along the freestream. This work characterizes the three-dimensional flow field of an oscillating airfoil at static yaw angles in reverse flow. Time-resolved velocity fields and chordwise pressure distributions are presented for several span locations, reduced frequencies, and Reynolds numbers. The unsteady velocity fields allow for the identification of dynamic stall vortex locations, and the unsteady pressure distributions allow for the analysis of spanwise variation in aerodynamic forces. By comparing the yawed and un-yawed cases, this work illustrates the relative importance of spanwise flow in reverse flow dynamic stall.

  3. A Study of Aerodynamic Control in Stalled Flight Leading-Edge Vortex Formation Analysis.

    DTIC Science & Technology

    1985-02-01

    RD-R153 758 A STUDY OF RERODYNAMIC CONTROL IN STALLED FLIGHT 1/1 LEADING-EDGE VORTEX FORMRTION RNALYSIS(U) ANRLYTICRL METHODS INC REDMOND NR J K...FORMATION ANALYSIS James K. Nathman ANALYTICAL METHODS , INC. 2047 - 152nd Avenue N.E. Redmond, Washington 98052 I February 1985 = Final Report for Period...ORGANIZATION Analytical Methods , Inc. (Ifapplicable) AF Flight Dynamics LaboratoryAFWAL (AFS0C 6c. ADDRESS (City, State and ZIP Code) 7b. ADDRESS

  4. A theory for the core of a three-dimensional leading-edge vortex

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    1985-01-01

    A theory is presented for the flow in the core of a separation-induced leading-edge vortex. 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.

  5. 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.

  6. Effect of leading-edge vortex flaps on aerodynamic performance of delta wings

    NASA Technical Reports Server (NTRS)

    Reddy, C. S.

    1981-01-01

    The effect of leading-edge vortex flaps on the aerodynamic characteristics of highly swept-back wings is analytically investigated, using the free vortex sheet method. The method, based on a three-dimensional inviscid flow model, is an advanced panel type employing quadratic doublet distributions to represent the wing surface, rolled-up vortex sheet and wake and is capable of computing forces, moments and surface pressures.

  7. 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.

  8. Influence of blade leading edge geometry and upstream blowing on the heat/mass transfer in a turbine cascade

    NASA Astrophysics Data System (ADS)

    Papa, Marco

    The effect of secondary flows on mass transfer from a simulated gas turbine blade and hubwall is investigated. Measurements performed using naphthalene sublimation provide non-dimensional mass transfer coefficients, in the form of Sherwood numbers, that can be converted to heat transfer coefficients through the use of an analogy. Tests are conducted in a linear cascade composed of five blades having the profile of a first stage rotor blade of a high-pressure turbine aircraft engine. Detailed mass transfer maps on the airfoil and endwall surfaces allow the identification of significant flow features that are in good agreement with existing secondary flow models. These results are well-suited for validation of numerical codes, as they are obtained with an accurate technique that does not suffer from conduction or radiation errors and allows the imposition of precise boundary conditions. The performance of a RANS (Reynolds Averaged Navier-Stokes) numerical code that simulates the flow and heat/mass transfer in the cascade using the SST (Shear Stress Transport) k-o model is evaluated through a comparison with the experimental results. Tests performed with a modified blade leading edge show that the introduction of a fillet at the junction with the endwall reduces the effects of the horseshoe vortex in the first part of the passage, while no measurable changes in mass transfer are observed further downstream. Air injected through a slot located upstream of the cascade simulates the engine wheelspace coolant injection between the stator and the rotor. Local mass transfer data obtained injecting naphthalene-free and naphthalene-saturated air are reduced to derive maps of cooling effectiveness on the blade and endwall. Oil dot tests show the surface flow on the endwall. The surface downstream of the gap is coplanar to the upstream surface in the baseline configuration and is shifted to form a forward and backward facing step to investigate the effects of component

  9. 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.

  10. 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.

  11. 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.

  12. Lesson from Tungsten Leading Edge Heat Load Analysis in KSTAR Divertor

    NASA Astrophysics Data System (ADS)

    Hong, Suk-Ho; Pitts, Richard Anthony; Lee, Hyeong-Ho; Bang, Eunnam; Kang, Chan-Soo; Kim, Kyung-Min; Kim, Hong-Tack; ITER Organization Collaboration; Kstar Team Team

    2016-10-01

    An important design issue for the ITER tungsten (W) divertor and in fact for all such components using metallic plasma-facing elements and which are exposed to high parallel power fluxes, is the question of surface shaping to avoid melting of leading edges. We have fabricated a series of tungsten blocks with a variety of leading edge heights (0.3, 0.6, 1.0, and 2.0 mm), from the ITER worst case to heights even beyond the extreme value tested on JET. They are mounted into adjacent, inertially cooled graphite tile installed in the central divertor region of KSTAR, within the field of view of an infra-red (IR) thermography system with a spatial resolution to 0.4 mm/pixel. Adjustment of the outer divertor strike point position is used to deposit power on the different blocks in different discharges. The measured power flux density on flat regions of the surrounding graphite tiles is used to obtain the parallel power flux, q|| impinging on the various W blocks. Experiments have been performed in Type I ELMing H-mode with Ip = 600 kA, BT = 2 T, PNBI = 3.5 MW, leading to a hot attached divertor with typical pulse lengths of 10 s. Three dimensional ANSYS simulations using q|| and assuming geometric projection of the heat flux are found to be consistent with the observed edge loading. This research was partially supported by Ministry of Science, ICT, and Future Planning under KSTAR project.

  13. 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.

  14. 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.

  15. Case Studies of Leading Edge Small Urban High Schools. Relevance Strategic Designs: 5. Life Academy of Health and Bioscience

    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…

  16. 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).

  17. Characterization of Unsteady Flow Structures Near Leading-Edge Slat. Part 1; PIV Measurements

    NASA Technical Reports Server (NTRS)

    Jenkins, Luther N.; Khorrami, Mehdi R.; Choudhari, Meelan

    2004-01-01

    A comprehensive computational and experimental study has been performed at the NASA Langley Research Center as part of the Quiet Aircraft Technology (QAT) Program to investigate the unsteady flow near a leading-edge slat of a two-dimensional, high-lift system. This paper focuses on the experimental effort conducted in the NASA Langley Basic Aerodynamics Research Tunnel (BART) where Particle Image Velocimetry (PIV) data was acquired in the slat cove and at the slat trailing edge of a three-element, high-lift model at 4, 6, and 8 degrees angle of attack and a freestream Mach Number of 0.17. Instantaneous velocities obtained from PIV images are used to obtain mean and fluctuating components of velocity and vorticity. The data show the recirculation in the cove, reattachment of the shear layer on the slat lower surface, and discrete vortical structures within the shear layer emanating from the slat cusp and slat trailing edge. Detailed measurements are used to examine the shear layer formation at the slat cusp, vortex shedding at the slat trailing edge, and convection of vortical structures through the slat gap. Selected results are discussed and compared with unsteady, Reynolds-Averaged Navier-Stokes (URANS) computations for the same configuration in a companion paper by Khorrami, Choudhari, and Jenkins (2004). The experimental dataset provides essential flow-field information for the validation of near-field inputs to noise prediction tools.

  18. Analysis of the Flow About Delta Wings with Leading Edge Separation at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Nenni, J. P.; Tung, C.

    1973-01-01

    A research program was conducted to develop an improved theoretical flow model for the flow about sharp edge delta wings with leading-edge separation at supersonic speeds. The flow model incorporates a representation of the secondary separation region which occurs just inboard of the leading edge on such wings and is based on a slender-wing theory whereby the full three-dimensional problem is reduced to a quasi two-dimensional problem in the cross-flow plane. The secondary separation region was modeled by a surface distribution of singularities or a linearized type of cavity representation. The primary vortex and separation were modeled by a concentrated vortex and cut in the cross-flow potential which represents its feeding sheet. The cross-flow solutions for the cavity model were obtained, but these solutions have physical significance only in a very restricted range of angle of attack. The reasons for the failure of the flow model are discussed. The analysis is presented so that other interested researchers may critically review the work.

  19. 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.

  20. 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.

  1. Prediction of unsteady aerodynamic loadings caused by leading edge and trailing edge control surface motions in subsonic compressible flow: Computer program description

    NASA Technical Reports Server (NTRS)

    Redman, M. C.; Rowe, W. S.

    1975-01-01

    A digital computer program has been developed to calculate unsteady loadings caused by motions of lifting surfaces with leading edge or trailing edge controls based on the subsonic kernel function approach. The pressure singularities at hinge line and side edges have been extracted analytically as a preliminary step to solving the integral equation by collocation. The program calculates generalized aerodynamic forces for user supplied deflection modes. Optional intermediate output includes pressure at an array of points, and sectional generalized forces. From one to six controls on the half span can be accommodated.

  2. On the unsteady motion and stability of a heaving airfoil in ground effect

    NASA Astrophysics Data System (ADS)

    Molina, Juan; Zhang, Xin; Angland, David

    2011-04-01

    This study explores the fluid mechanics and force generation capabilities of an inverted heaving airfoil placed close to a moving ground using a URANS solver with the Spalart-Allmaras turbulence model. By varying the mean ground clearance and motion frequency of the airfoil, it was possible to construct a frequency-height diagram of the various forces acting on the airfoil. The ground was found to enhance the downforce and reduce the drag with respect to freestream. The unsteady motion induces hysteresis in the forces' behaviour. At moderate ground clearance, the hysteresis increases with frequency and the airfoil loses energy to the flow, resulting in a stabilizingmotion. By analogy with a pitching motion, the airfoil stalls in close proximity to the ground. At low frequencies, the motion is unstable and could lead to stall flutter. A stall flutter analysis was undertaken. At higher frequencies, inviscid effects overcome the large separation and the motion becomes stable. Forced trailing edge vortex shedding appears at high frequencies. The shedding mechanism seems to be independent of ground proximity. However, the wake is altered at low heights as a result of an interaction between the vortices and the ground.

  3. 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.

  4. 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

  5. Effects of Alternate Leading Edge Cutback on the Space Shuttle Main Engine Low Pressure Fuel Pump

    NASA Technical Reports Server (NTRS)

    Mulder, Andrew; Skelley, Stephen

    2016-01-01

    A higher order cavitation oscillation observed in the SSME low pressure fuel pump has been eliminated in water flow testing of a modified subscale replica of the inducer. The low pressure pump was modified by removing the outboard sections of two opposing blades of the four-bladed inducer, blending the "cutback" regions into the blades at the leading edge and tip, and removing material on the suction sides to decrease the exposed leading edge thickness. The leading edge tips of the cutback blades were moved approximately 25 degrees from their previous locations, thereby increasing one blade to blade spacing, decreasing the second, while simultaneously moving the cutback tips downstream. The test was conducted in MSFC's inducer test loop at scaled operating conditions in degassed and filtered water. In addition to eliminating HOC across the entire scaled operating regime, rotating cavitation was suppressed while the range of both alternate blade and asymmetric cavitation were increased. These latter phenomena, and more significantly, the shifts between these cavitation modes also resulted in significant changes to the head coefficient at low cavitation numbers. Reverse flow was detected at a slightly larger flow coefficient with the cutback inducer and suction capability was reduced by approximately 1 velocity head at and above approximately 90% of the reference flow coefficient. These performance changes along with more intense reverse flow are consistent with poor flow area management and increased incidence in the cutback region. Although the test demonstrated that the inducer modification was successful at eliminating the higher order cavitation across the entire scaled operating regime, different, previously unobserved, cavitation oscillations were introduced and significant performance penalties were imposed.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. Measurement of Leading Edge Vortices from a Delta Wing Using a Three Component Laser Velocimeter

    NASA Technical Reports Server (NTRS)

    Meyers, James F.; Hepner, Timothy E.

    1988-01-01

    A demonstration of the capabilities of a three component laser velocimeter to provide a detailed experimental database of a complex flow field i s presented. The orthogonal three component laser velocimeter was used to measure the leading edge vortex flow field above a 75 degrees delta wing at angles-of-attack of 20.5 degrees and 40.0 degrees. The resulting mean velocity and turbulence intensity measurements are presented. The laser velocimeter is described in detail including a description of the data processing algorithm. A full error analysis was conducted and the results presented.

  12. Case study: on the leading edge of new curricula concepts: systems and safety in nursing education.

    PubMed

    Dick, Diana Davidson; Weisbrod, Lorna; Gregory, David; Dyck, Netha; Neudorf, Kim

    2006-09-01

    The Nursing Division of the Saskatchewan Institute of Applied Science and Technology (SIAST) first included systems and patient safety as a priority in its institutional business and strategic plan in 2003. Three interrelated leading-edge, two-year projects (2004-2006) were launched: Best Practice, Mentorship and Patient Safety, with the intent that each project would enhance the others. This case study focuses on the work of the Patient Safety Project Team. The team developed a project framework and strategic plan, conducted a literature review and identified key concepts related to systems and patient safety. Strategies to integrate these concepts into the school's 15 nursing education programs are being implemented.

  13. 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.

  14. Compressible Navier-Stokes equations: A study of leading edge effects

    NASA Technical Reports Server (NTRS)

    Hariharan, S. I.; Karbhari, P. R.

    1987-01-01

    A computational method is developed that allows numerical calculations of the time dependent compressible Navier-Stokes equations.The current results concern a study of flow past a semi-infinite flat plate.Flow develops from given inflow conditions upstream and passes over the flat plate to leave the computational domain without reflecting at the downstream boundary. Leading edge effects are included in this paper. In addition, specification of a heated region which gets convected with the flow is considered. The time history of this convection is obtained, and it exhibits a wave phenomena.

  15. Electrogasdynamic excitation of controlling disturbances near a swept wing leading edge

    NASA Astrophysics Data System (ADS)

    Chernyshev, Sergey; Kiselev, Andrey; Kuryachii, Aleksandr

    2016-10-01

    New design of multiple plasma actuator intended for the excitation of disturbances in boundary layer near a leading edge of a swept wing is proposed. The excited disturbances have to suppress the cross-flow-type instability modes provoking laminar-to-turbulent transition in usual conditions. Numerical modeling of the excitation of controlling disturbances by plasma actuator has been executed in stationary approximation for the case of infinite span swept wing at subsonic cruise flight conditions. Localized volumetric force and heat impact of actuator periodic along a wing span has been considered. Calculations have been executed for physical parameters of impact typical for surface dielectric barrier discharge.

  16. An attached flow design of a noninterfering leading edge extension to a thick delta wing

    NASA Technical Reports Server (NTRS)

    Ghaffari, F.; Lamar, J. E.

    1985-01-01

    The analytical procedure presented for leading edge extension (LEE) determination, in keeping with such design criteria as noninterference at the wing design point, is applied to thick delta wings. The LEE device thus defined is to be mounted on a wing along a dividing stream surface that is associated with an attached flow design lift coefficient greater than zero. The delta wing in question is of twisted and cambered type. It is demonstrated that span reductions for the candidate LEEs has the most detrimental effect on overall aerodynamic efficiency, irrespective of area or shape.

  17. Analysis of In-Flight Structural Failures of P-3C Wing Leading Edge Segments

    DTIC Science & Technology

    1992-06-01

    the remaining distance from the outboard engines to the wing tips. The length (fore and aft) of these leading-edge sections is 15% of the chord (total... Chord , Root (ft) 18.9 Tip 7.6 Aileron Area, S. (ft 2 ) 45.5 Hinge Line (cw) 0.725 Deflection Limit, Up (degrees) -23.3 Down +16.2 Horizontal Tail Area...playing a large role in the problem because of the location of the wing’s elastic axis at a constant 40 percent of chord , according to available

  18. Spanwise pressure distribution on delta wing with leading-edge vortex flap

    NASA Technical Reports Server (NTRS)

    Reddy, C. S.

    1987-01-01

    The aerodynamic characteristics of a highly swept planar delta wing employing conical leading edge flaps are numerically investigated, using a free vortex sheet method that is based on an advanced, three-dimensional inviscid flow panel method employing quadratic doublet distributions to represent the wing surface and the rolled-up vortex sheet and wake. Upward flap deflection shifts the negative pressure peak inboard of the basic wing and develops a significant suction pressure on the flap that then produces thrust component in the direction of flight; overall drag is thereby reduced.

  19. 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.

  20. Transonic airfoil analysis and design using Cartesian coordinates

    NASA Technical Reports Server (NTRS)

    Carlson, L. A.

    1975-01-01

    An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.

  1. 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.

  2. Localized deformation zones in the offshore leading edge of the Yakutat microplate, Gulf of Alaska

    NASA Astrophysics Data System (ADS)

    Lowe, L. A.; Gulick, S. P.; Pavlis, T.; Bruhn, R. L.; Mann, P.

    2006-12-01

    The Gulf of Alaska margin is dominated by the collision and subduction of the Yakutat microplate as it travels northwest with respect to North America at near Pacific Plate velocities (\\~45 mm/yr). The oblique Yakutat block collision with North America is in transition between convergence to the west and translation along the Queen Charlotte-Fairweather-Denali Fault system to the east and north. Industry seismic reflection and high- resolution seismic reflection data collected by the R/V Maurice Ewing (2004) provides insight into how the Yakutat-North America collision is accommodated by active offshore structures near the leading edge of the Yakutat microplate. A \\~200 km wide area bounded by the Ten Fathom Fault, the offshore N. America-Yakutat contact, to the west and the eastern edge of the Pamplona Zone (PZ) to the east has previously been mapped as a continuous deformation zone consisting of NE-SW trending imbricate thrusts and folds. Though this mapping corroborates onshore measurements of active deformation west of the Bering Glacier in the Yakutat block, the relationship between current onshore deformation and the observed offshore structures remains unclear. Our observations indicate that neotectonic deformation is accommodated offshore by highly localized, asynchronous thrusts that, when analyzed in an accretionary context, may be connected by a sub-horizontal decollement. Data from the eastern edge of the PZ, the proposed deformation front, shows surface deformation caused by east-verging thrust faults. Seismic reflection profiles in the western PZ and the Bering Trough show no evidence of active tectonic deformation and up to \\~200 m of undisturbed sediments indicating that faulting in this part of the Yakutat block has been inactive since the Last Glacial Maximum or earlier. Farther west, above the Kayak Island fault zone, directly east of the Ten Fathom Fault, the presence of up to \\~50 m of undeformed sediments suggests a recent (ca. 14 ka

  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. 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.

  5. Low-order phenomenological modeling of leading-edge vortex formation

    NASA Astrophysics Data System (ADS)

    Wang, Chengjie; Eldredge, Jeff D.

    2013-09-01

    A low-order point vortex model for the two-dimensional unsteady aerodynamics of a flat plate wing section is developed. A vortex is released from both the trailing and leading edges of the flat plate, and the strength of each is determined by enforcing the Kutta condition at the edges. The strength of a vortex is frozen when it reaches an extremum, and a new vortex is released from the corresponding edge. The motion of variable-strength vortices is computed in one of two ways. In the first approach, the Brown-Michael equation is used in order to ensure that no spurious force is generated by the branch cut associated with each vortex. In the second approach, we propose a new evolution equation for a vortex by equating the rate of change of its impulse with that of an equivalent surrogate vortex with identical properties but constant strength. This impulse matching approach leads to a model that admits more general criteria for shedding, since the variable-strength vortex can be exchanged for its constant-strength surrogate at any instant. We show that the results of the new model, when applied to a pitching or perching plate, agree better with experiments and high-fidelity simulations than the Brown-Michael model, using fewer than ten degrees of freedom. We also assess the model performance on the impulsive start of a flat plate at various angles of attack. Current limitations of the model and extensions to more general unsteady aerodynamic problems are discussed.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. 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.

  11. 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.

  12. Polarization of chemokine receptors to the leading edge during lymphocyte chemotaxis.

    PubMed

    Nieto, M; Frade, J M; Sancho, D; Mellado, M; Martinez-A, C; Sánchez-Madrid, F

    1997-07-07

    Leukocyte migration in response to cell attractant gradients or chemotaxis is a key phenomenon both in cell movement and in the inflammatory response. Chemokines are quite likely to be the key molecules directing migration of leukocytes that involve cell polarization with generation of specialized cell compartments. The precise mechanism of leukocyte chemoattraction is not known, however. In this study, we demonstrate that the CC chemokine receptors CCR2 and CCR5, but not cytokine receptors such as interleukin (IL)-2Ralpha, IL-2Rbeta, tumor necrosis factor receptor 1, or transforming growth factor betaR, are redistributed to a pole in T cells that are migrating in response to chemokines. Immunofluorescence and confocal microscopy studies show that the chemokine receptors concentrate at the leading edge of the cell on the flattened cell-substratum contact area, induced specifically by the signals that trigger cell polarization. The redistribution of chemokine receptors is blocked by pertussis toxin and is dependent on cell adhesion through integrin receptors, which mediate cell migration. Chemokine receptor expression on the leading edge of migrating polarized lymphocytes appears to act as a sensor mechanism for the directed migration of leukocytes through a chemoattractant gradient.

  13. Aerothermal Performance Constraints for Small Radius Leading Edges Operating at Hypervelocity

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul; Bull, Jeffrey D.; Milos, Frank S.; Squire, Thomas H.

    1997-01-01

    Small radius leading edges and nosetips were used to minimize wave drag in early hypervelocity vehicle concepts until further analysis demonstrated that extreme aerothermodynamic heating blunted the available thermal protection system materials. Recent studies indicate that ultra-high temperature composite (UHTC) materials are shape stable at temperatures approaching 3033 K and will be available for use as sharp leading edge components in the near future. Steady-state aerothermal performance constraints for UHTC components are presented in this paper to identify their non-ablating operational capability at altitudes from sea level to 90 km. An integrated design tool was developed to estimate these constraints. The tool couples aerothermodynamic heating with material response using commercial finite element analysis software and is capable of both steady-state and transient analysis. Performance during entry is analyzed by transient thermal analysis along the trajectory. The thermal load condition from the transient thermal analysis is used to estimate thermal stress. Applying the tool to UHTC materials shows that steady-state, non-ablating operation of a HfB2/SiC(A-7) (A-7) component is possible at velocities approaching Earth's circular orbital velocity of 7.9 km/s at altitudes approaching 70 km.

  14. 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.

  15. 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.

  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.

    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.

  17. 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.

  18. 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.

  19. DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number

    NASA Astrophysics Data System (ADS)

    Pradhan, Sahadev, , Dr.

    2017-01-01

    The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf /√{kBTinf / m }) in the range leading-edge flat plate boundary layer at high Mach number. Here, LT is the characteristic dimension, Uinf and Tinf are the free stream velocity and temperature, rhoinf is the free stream density, m is the molecular mass, muinf is the molecular viscosity based on the free stream temperature Tinf , and kB is the Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.

  20. 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.

  1. Water tunnel results of leading-edge vortex flap tests on a delta wing vehicle

    NASA Technical Reports Server (NTRS)

    Delfrate, J. H.

    1986-01-01

    A water tunnel flow visualization test on leading edge vortex flaps was conducted at the flow visualization facility of the NASA Ames Research Center's Dryden Flight Research Facility. The purpose of the test was to visually examine the vortex structures caused by various leading edge vortex flaps on the delta wing of an F-106 model. The vortex flaps tested were designed analytically and empirically at the NASA Langley Research Center. The three flap designs were designated as full-span gothic flap, full-span untapered flap, and part-span flap. The test was conducted at a Reynolds number of 76,000/m (25,000/ft). This low Reynolds number was used because of the 0.076-m/s (0.25-ft/s) test section flow speed necessary for high quality flow visualization. However, this low Reynolds number may have influenced the results. Of the three vortex flaps tested, the part-span flap produced what appeared to be the strongest vortex structure over the flap area. The full-span gothic flap provided the next best performance.

  2. Plasma Flow Control Optimized Airfoil

    NASA Astrophysics Data System (ADS)

    Voikov, Vladimir; Patel, Mehul

    2005-11-01

    Recent advances in flow control research have demonstrated that plasma actuators can be efficient in different aerodynamic applications, particularly in providing flight control without conventional moving surfaces. The concept involves the use of a laminar airfoil design that employs a separation ramp at the trailing edge that can be manipulated by a plasma actuator to control lift, similar to trailing-edge flaps. The advantages are lower drag by a combination of the laminar flow design, and elimination of parasitic drag associated with wing-flap junctions. This work involves numerical simulations and experiments on a HSNLF(1)-0213 airfoil. The numerical results are obtained using an unsteady, compressible Navier-Stokes simulation that includes a model for the plasma actuators. The experiments are performed on a 2-D airfoil section that is mounted on a lift-drag force balance. The results demonstrate lift enhancement produced by the plasma actuator that is comparable to a plane flap. They also reveal an optimum actuator unsteady frequency that scales with the length of the separated region and local velocity, and is associated with the generation of a train of spanwise vortices. Other scaling including the effect of Reynolds number is presented.

  3. 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

  4. A Wind Tunnel Experiment for Trailing Edge Circulation Control on a 6 Percent 2-D Airfoil up to Transonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Alexander, Michael G.; Anders, Scott G.; Johnson, Stuart K.

    2005-01-01

    A wind tunnel test was conducted on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at mass flow coefficients (C(sub mu)) from 0.0 to 0.12. The test data was acquired in the NASA Langley Transonic Dynamics Tunnel at Mach numbers of 0.8 and 0.3 at Reynolds numbers per foot of 1.05 x 10(exp 6) and 2.43 x 10(exp 5) respectively. For the transonic condition, (Mach = 0.8 at alpha = +3 deg), it was generally found that the smaller slot and larger Coanda surface were more effective overall than other slot/Coanda surface combinations. Generally it was found at Mach = 0.3 at alpha = 6 deg that the smaller slot and smaller Coanda surface were more effective overall than other slot/Coanda surface combinations.

  5. Formation of Leading-Edge Pinholes in the Space Shuttle Wings Investigated

    NASA Technical Reports Server (NTRS)

    Jacobson, Nathan S.

    2000-01-01

    The space shuttle wing leading edge and nose cap are composed of a carbon/carbon composite that is protected by silicon carbide. The coefficient of thermal expansion mismatch leads to cracks in the silicon carbide. The outer coating of the silicon carbide is a sodium-silicate-based glass that becomes fluid at the shuttles high reentry temperatures and fills these cracks. Small pinholes roughly 0.1 mm in diameter have been observed on these materials after 12 or more flights. These pinholes have been investigated by researchers at the NASA Johnson Space Center, Rockwell International, the Boeing Company, Lockheed Martin Corporation, and the NASA Glenn Research Center at Lewis Field to determine the possible sources and the extent of damage. A typical pinhole is illustrated in the photomicrographs. These pinholes are found primarily on the wing leading edges and not on the nose cap, which is covered when the orbiter is on the launch pad. The pinholes are generally associated with a bead of zincrich glass. Examination of the orbiter and launch structure indicates that weathering paint on the launch structure leads to deposits of zinc-containing paint flakes on the wing leading edge. These may become embedded in the crevices of the wing leading edge and form the observed zinc-rich glass. Laboratory experiments indicate that zinc oxide reacts vigorously with the glass coating on the silicon carbide. Thus, it is likely that this is the reaction that leads to pinhole formation (Christensen, S.V.: Reinforced Carbon/Carbon Pin Hole Formation Through Zinc Oxide Attack. Rockwell International Internal Letter, RDW 96 057, May 1996). Cross-sectional examination of pinholes suggests that they are enlarged thermal expansion mismatch cracks. This is illustrated in the photomicrographs. A careful microstructural analysis indicates that the pinhole walls consist of layers of zinc-containing glass. Thus, pinholes are likely formed by zinc oxide particles lodging in crevices and

  6. 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.

  7. Passive Boundary Layer Separation Control on a NACA2415 Airfoil at High Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Parikh, Agastya; Hultmark, Marcus

    2016-11-01

    The design and analysis of a passive flow control system for a NACA2415 airfoil is undertaken. There exists a vast body of knowledge on airfoil boundary layer control with the use of controlled mass flux, but there is little work investigating passive mass flux-based methods. A simple duct system that uses the upper surface pressure gradient to force blowing near the leading edge and suction near the trailing edge is proposed and evaluated. 2D RANS analyses at Rec 1 . 27 ×106 were used to generate potential configurations for experimental tests. Initial computational results suggest drag reductions of approximately 2 - 7 % as well as lift increases of 4 - 5 % at α = 10 .0° and α = 12 .5° . A carbon composite-aluminum structure model that implements the most effective configurations, according to the CFD predictions, has been designed and fabricated. Experiments are being performed to evaluate the CFD results and the feasibility the duct system.

  8. 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.

  9. Summary of Section Data on Trailing-Edge High-Lift Devices

    NASA Technical Reports Server (NTRS)

    1948-01-01

    A summary has been made of available data on the characteristics of airfoil sections with trailing-edge high-lift devices. Data for plain, split, and slotted flaps are collected and analyzed. The effects of each of the variables involved in the design of the various types of flap are examined and, in cases where sufficient data are given, optimum configurations are deduced. Wherever possible, the effects of airfoil section, Reynolds number, and leading-edge roughness are shown. For single and double slotted flaps, where a great mass of unrelated date are available, maximum lift coefficients of a large number of configurations are presented in tables.

  10. Viscous Thin Airfoil Theory

    DTIC Science & Technology

    1980-02-01

    the elliptic cross section is considered to be more representative of the NACA 64A010 airfoil with boundary layer displacement thickness added on than...section and the flat plate airfoil with Kutta condition. The experimental results are for the NACA 64A010 airfoil at M = 0.5 and Reynolds number between...practice for actual airfoils. The experimental data shown in Fig. 3.5 are for the NACA 4 and 5 digit series airfoils (Ref. 17). The lift curve slope is

  11. Experimental Pressure Distributions over Wing Tips at Mach Number 1.9 I : Wing Tip with Subsonic Leading Edge

    NASA Technical Reports Server (NTRS)

    Jagger, James M; Mirels, Harold

    1949-01-01

    An investigation was conducted at a Mach number of 1.91 to determine spanwise pressure distribution over a wing tip in a region influenced by a sharp subsonic leading edge swept back at 70 degrees. Except for pressure distribution on the top surface in the immediate vicinity of the subsonic leading edge, the maximum difference between linearized theory and experimental data was 2 1/2 percent (of free-stream dynamic pressure) for angles of attack up to 4 degrees and 7 percent for angles of attack up to 8 degrees. Pressures on the top surface nearest the subsonic edge indicated local expansions beyond values predicted by linearized theory.

  12. 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

  13. 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.

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

    NASA Technical Reports Server (NTRS)

    Torres, Francisco J.

    1987-01-01

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

  15. Small-Scale Transonic Investigation of the Effects of Partial-Span Leading-Edge Camber on the Aerodynamic Characteristics of a 50 Deg 38' Sweptback Wing of Aspect Ratio 2.98

    NASA Technical Reports Server (NTRS)

    Alford, William J., Jr.; Byrnes, Andrew L., Jr.

    1952-01-01

    A small-scale transonic investigation of two semispan wings of the same plan form was made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range of 0.70 to 1.10 and a mean-test Reynolds number range of 745,000 to 845,000 to determine the effects of partial-span leading-edge camber on the aerodynamic characteristics of a swept-back wing. This paper presents the results of the investigation of wing-alone and wing-fuselage configurations of the two wings; one, was an uncambered wing and the other had the forward 45 percent of the chord cambered over the outboard 55 percent of the span. The semispan wings had 50deg 38ft sweepback of their quarter-chord lines, aspect ratio of 2.98, taper ratio of 0.45, and modified NACA 64A-series airfoil sections tapered in thickness ratio. Lift, drag, pitching moment, and root-bending moment were obtained for these configurations. The results indicated that, for the wing-alone configuration, use of the partial-span leading-edge camber provided an increase in maximum lift-drag ratios up to a Mach number of 0.95, after which no gain was realized. For the wing-fuselage combination, the partial-span leading-edge camber appeared to cause no gain in maximum lift-drag ratio throughout the test range of Mach numbers. The lift-curve slopes of the partial-span leading-edge camber configurations indicated no significant change over the basic configurations in the subsonic range but resulted in slight reductions at the higher Mach numbers. No significantly large changes in pitching-moment-curve slopes or lateral center of additional loading were indicated because of the modification.

  16. Numerical investigation of multi-element airfoils

    NASA Technical Reports Server (NTRS)

    Cummings, Russell M.

    1993-01-01

    The flow over multi-element airfoils with flat-plate lift-enhancing tabs was numerically investigated. Tabs ranging in height from 0.25 percent to 1.25 percent of the reference airfoil chord were studied near the trailing edge of the main-element. This two-dimensional numerical simulation employed an incompressible Navier-Stokes solver on a structured, embedded grid topology. New grid refinements were used to improve the accuracy of the solution near the overlapping grid boundaries. The effects of various tabs were studied at a constant Reynolds number on a two-element airfoil with a slotted flap. Both computed and measured results indicated that a tab in the main-element cove improved the maximum lift and lift-to-drag ratio relative to the baseline airfoil without a tab. Computed streamlines revealed that the additional turning caused by the tab may reduce the amount of separated flow on the flap. A three-element airfoil was also studied over a range of Reynolds numbers. For the optimized flap rigging, the computed and measured Reynolds number effects were similar. When the flap was moved from the optimum position, numerical results indicated that a tab may help to reoptimize the airfoil to within 1 percent of the optimum flap case.

  17. 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.

  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. New convergence criteria for the vortex-lattice models of the leading-edge separation

    NASA Technical Reports Server (NTRS)

    Kandil, O. A.; Mook, D. T.; Nayfeh, A. H.

    1976-01-01

    The convergence criterion for the vortex-lattice technique which deals with delta wings exhibiting significant leading-edge separation was studied. It was shown that one can predict pressure distributions without irregularities which agree fairly well with experimental data (which show some irregularities of their own) by replacing the system of discrete vortex lines with a single concentrated core. This core has a circulation equal to the algebraic sum of the circulations around the discrete lines and is located at the centroid of these lines. Moreover, there is a requirement that the position and strength of the core must converge as the number of elements increases. Because the calculation of the position and strength of the core is much less involved than the calculation of the loads, this approach has the additional desirable feature of requiring less computational time.

  20. Comparison of two leading uniform theories of edge diffraction with the exact uniform asymptotic solution

    NASA Technical Reports Server (NTRS)

    Boersma, J.; Rahmat-Samii, Y.

    1980-01-01

    The diffraction of an arbitrary cylindrical wave by a half-plane has been treated by Rahmat-Samii and Mittra who used a spectral domain approach. In this paper, their exact solution for the total field is expressed in terms of a new integral representation. For large wave number k, two rigorous procedures are described for the exact uniform asymptotic expansion of the total field solution. The uniform expansions obtained are valid in the entire space, including transition regions around the shadow boundaries. The final results are compared with the formulations of two leading uniform theories of edge diffraction, namely, the uniform asymptotic theory and the uniform theory of diffraction. Some unique observations and conclusions are made in relating the two theories.

  1. Computation of leading edge film cooling from a CONSOLE geometry (CONverging Slot hOLE)

    NASA Astrophysics Data System (ADS)

    Guelailia, A.; Khorsi, A.; Hamidou, M. K.

    2016-01-01

    The aim of this study is to investigate the effect of mass flow rate on film cooling effectiveness and heat transfer over a gas turbine rotor blade with three staggered rows of shower-head holes which are inclined at 30° to the spanwise direction, and are normal to the streamwise direction on the blade. To improve film cooling effectiveness, the standard cylindrical holes, located on the leading edge region, are replaced with the converging slot holes (console). The ANSYS CFX has been used for this computational simulation. The turbulence is approximated by a k-ɛ model. Detailed film effectiveness distributions are presented for different mass flow rate. The numerical results are compared with experimental data.

  2. Application of finite element and remeshing technique to shock interference on a cylindrical leading edge

    NASA Technical Reports Server (NTRS)

    Stewart, James R.; Thareja, Rajiv R.; Wieting, Allan R.; Morgan, Ken

    1988-01-01

    The problem of planar oblique shock impingement on a cylindrical leading edge in hypersonic flow is modeled using a Galerkin-Runge Kutta finite element method. The method utilizes a four stage Runge-Kutta time stepping scheme to solve the compressible Euler equations. Freestream Mach numbers of 6.5, 8.0 and 16.0 are studied. The computed surface pressure distributions consistently agree well with available experimental data. The peak pressure amplification ranges from 5.45 at M = 6.5 to approximately 17.0 at M = 16.0. Stagnation point heat transfer rate amplifications are calculated from the inviscid solution using the method of Fay and Riddell. The value and wall location of the peak pressure and heat transfer rate amplifications are extremely sensitive to the location of the impinging shock/bow shock intersection point.

  3. 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.

  4. Aerothermal Performance Envelopes for Hypersonic Small Radius Unswept Leading Edges and Nosetips

    NASA Technical Reports Server (NTRS)

    Kolodziej, Paul; Rasky, Daniel J. (Technical Monitor)

    1995-01-01

    Small radius leading edges and nosetips were utilized to minimize wave drag in early hypersonic 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 diboride composite materials are shape stable under extreme aerothermodynamic heating at ultra high temperatures. Aerothermal performance envelopes 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 altitudes near 90 km.

  5. 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.

  6. Computational Modeling of a Mechanized Benchtop Apparatus for Leading-Edge Slat Noise Treatment Device Prototypes

    NASA Technical Reports Server (NTRS)

    Turner, Travis L.; Moore, James B.; Long, David L.

    2017-01-01

    Airframe noise is a growing concern in the vicinity of airports because of population growth and gains in engine noise reduction that have rendered the airframe an equal contributor during the approach and landing phases of flight for many transport aircraft. The leading-edge-slat device of a typical high-lift system for transport aircraft is a prominent source of airframe noise. Two technologies have significant potential for slat noise reduction; the slat-cove filler (SCF) and the slat-gap filler (SGF). Previous work was done on a 2D section of a transport-aircraft wing to demonstrate the implementation feasibility of these concepts. Benchtop hardware was developed in that work for qualitative parametric study. The benchtop models were mechanized for quantitative measurements of performance. Computational models of the mechanized benchtop apparatus for the SCF were developed and the performance of the system for five different SCF assemblies is demonstrated.

  7. Leading-edge vortex improves lift in slow-flying bats.

    PubMed

    Muijres, F T; Johansson, L C; Barfield, R; Wolf, M; Spedding, G R; Hedenström, A

    2008-02-29

    Staying aloft when hovering and flying slowly is demanding. According to quasi-steady-state aerodynamic theory, slow-flying vertebrates should not be able to generate enough lift to remain aloft. Therefore, unsteady aerodynamic mechanisms to enhance lift production have been proposed. Using digital particle image velocimetry, we showed that a small nectar-feeding bat is able to increase lift by as much as 40% using attached leading-edge vortices (LEVs) during slow forward flight, resulting in a maximum lift coefficient of 4.8. The airflow passing over the LEV reattaches behind the LEV smoothly to the wing, despite the exceptionally large local angles of attack and wing camber. Our results show that the use of unsteady aerodynamic mechanisms in flapping flight is not limited to insects but is also used by larger and heavier animals.

  8. Development of a Transition Edge Sensor Gamma Ray Microcalorimeter with an Epoxy Coupled Bulk Lead Absorber

    SciTech Connect

    Damayanthi, R. M. T.; Iyomoto, N.; Takahashi, H.; Minamigawa, Y.; Nishimura, K.; Ohno, M.

    2009-12-16

    Transition edge sensor (TES)-based gamma ray detectors have been developed primarily for use up to energies of {approx}100 keV. However, there are many interesting applications at higher energies. We have started to develop a TES gamma-ray detector to apply to Positron Annihilation Spectroscopy analysis at 511 keV. Our detector is composed of a bulk lead absorber, which is coupled to a thin-film TES using a small amount of epoxy. The response of our first detector showed a very long decay tail of {approx}135 ms. To improve the device response time we have designed a new detector in which the response time is improved by a factor of five.

  9. 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.

  10. Calibration of sonic valves for the laminar flow control, leading-edge flight test

    NASA Technical Reports Server (NTRS)

    Petley, D. H.; Alexander, W., Jr.; Wright, A. S., Jr.; Vallas, M.

    1985-01-01

    Sonic needle valves were calibrated to measure and control airflow in the suction system for the leading-edge flight test. The procedure and results for the calibration flow test of 4:41 flight valves are given. Mass-flow rates, which ranged from 0.001 to 0.012 lbm/sec, and maximum back pressure were measured for total temperatures from -30 F to 75 F and total pressures from 120 to 540 psf. Correlating equations are obtained for mass-flow rate as a function of total pressure, total temperature, and valve opening length. The most important aspect of flow measurement and control is found to be the measurement of valve opening length.

  11. 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.

  12. 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.

  13. An Attached Flow Design of a Noninterferring Leading Edge Extension to a Thick Delta Wing

    NASA Technical Reports Server (NTRS)

    Lamar, John E.; Ghaffari, Farhad

    1985-01-01

    An analytical procedure for the determination of the shape of a Leading-Edge Extension (LEE) which satisfies design criteria, including especially noninterference at the wing design point, has been developed for thick delta wings. The LEE device best satisfying all criteria is designed to be mounted on a wing along a dividing stream surface associated with an attached flow design lift coefficient (C(sub L,d)) of greater than zero. This device is intended to improve the aerodynamic performance of transonic aircraft at C(sub L) greater than C(sub L,d) system emanating from the LEE leading edge. In order to quantify this process a twisted and cambered thick delta wing was chosen for the initial application of this design procedure. Appropriate computer codes representing potential and vortex flows were employed to determine the dividing stream surface at C(sub L,d) and an optimized LEE planform shape at C(sub L) greater than C(sub L,d), respectively. To aid in the LEE selection, the aerodynamic effectiveness of 36 planforms was investigated at C(sub L) greater than C(sub L,d). This study showed that reducing the span of the candidate LEEs has the most detrimental effect on overall aerodynamic efficiency, regardless of the shape or area. Furthermore, for a fixed area, constant-chord LEE candidates were relatively more efficient than those with sweep less than the wing. At C(sub L,d), the presence of the LEE planform best satisfying the design criteria was found to have no effect on the wing alone aerodynamic performance.

  14. 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.

  15. Two experimental supercritical laminar-flow-control swept-wing airfoils

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Dagenhart, J. Ray

    1987-01-01

    Two supercritical laminar-flow-control airfoils were designed for a large-chord swept-wing experiment in the Langley 8-Foot Transonic Pressure Tunnel where suction was provided through most of the model surface for boundary-layer control. The first airfoil was derived from an existing full-chord laminar airfoil by extending the trailing edge and making changes in the two lower-surface concave regions. The second airfoil differed from the first one in that it was designed for testing without suction in the forward concave region of the lower surface. Differences between the first airfoil and the one from which it was derived as well as between the first and second airfoils are discussed. Airfoil coordinates and predicted pressure distributions for the design normal Mach number of 0.755 and section lift coefficient of 0.55 are given for the three airfoils.

  16. Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Hernandez, Gloria; Wood, Richard M.; Covell, Peter F.

    1994-01-01

    An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

  17. 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.

  18. 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.

  19. 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.

  20. 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.

  1. The effect of butterfly scales on flight efficiency and leading edge vortex formation

    NASA Astrophysics Data System (ADS)

    Lang, Amy; Wilroy, Jacob; Wahidi, Redha; Slegers, Nathan; Heilman, Micahel; Cranford, Jacob

    2016-11-01

    It is hypothesized that the scales on a butterfly wing lead to increased flight efficiency. Recent testing of live butterflies tracked their motion over 246 flights for 11 different specimens. Results show a 37.8 percent mean decrease in flight efficiency and a flapping amplitude reduction of 6.7 percent once the scales were removed. This change could be largely a result of how the leading edge vortex (LEV) interacts with the wing. To simplify this complex flow problem, an experiment was designed to focus on the alteration of 2-D vortex development with a variation in surface patterning. Specifically, the secondary vorticity generated by the LEV interacting at the patterned surface was studied, as well as the subsequent effect on the LEV's growth rate and peak circulation. For this experiment butterfly inspired grooves were created using additive manufacturing and were attached to a flat plate with a chordwise orientation, thus increasing plate surface area. The vortex generated by the grooved plate was then compared to a smooth case as the plate translated vertically through a tow tank at Re = 1500, 3000, and 6000. Using DPIV, the vortex formation was documented and a maximum vortex formation time of 4.22 was found based on the flat plate travel distance and chord length. Results indicate that the patterned surface slows down the growth of the vortex which corroborates the flight test results. Funding from NSF CBET Fluid Dynamcis is gratefully acknowledged.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. Blowing Circulation Control on a Seaplane Airfoil

    NASA Astrophysics Data System (ADS)

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

    2011-09-01

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

  7. 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

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

    NASA Technical Reports Server (NTRS)

    Amari, S.; Foote, J.; 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.

  9. Experimental Analysis of the Vorticity and Turbulent Flow Dynamics of a Pitching Airfoil at Realistic Flight Conditions

    DTIC Science & Technology

    2007-08-31

    insert: midspan at the leading edge) 99 Figure 4.48 Hydraulic drive system reservoir 100 Figure 4.49 Accumulator with servo valve 100 Figure 4.50...supplied the required amount of hydraulic fluid through the servo valve . During downward movement of the piston, oil returned to the accumulator...through a similar high pressure hose. A Parker Hannifin servo valve controlled the amount of hydraulic fluid required for the pitching of the airfoil

  10. 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.

  11. 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.

  12. Influence of high-intensity turbulence on laminar boundary layer development on a cylindrical leading edge: Enhancement to eddy diffusivity

    NASA Astrophysics Data System (ADS)

    Pearson, Juli K.

    The growing demand for increased efficiency in turbine engine designs has sparked a growing interest for research of air flow around curved surfaces. The turbine's operating conditions result in material property constraints, especially in the first stage turbine vanes and blades. These turbine vane components experience extreme loading conditions of both high temperature and high turbulence intensities exiting the combustor. The surface of the turbine blades has cylindrical leading edges that promote stabilizing flow accelerations. These convex surfaces can cause a reduced eddy diffusivity across the boundary layer. This thesis reviews measurements of velocity and turbulence intensities taken just shy of the thirty degrees offset from the stagnation line of a two-dimensional cylindrical leading edge under a wide range of turbulence and flow conditions flow conditions. Flow conditions and velocity measurements were gathered with respect to the distance to the surface. The length of the measurements extended from the surface to beyond the boundary layer's edge. The instrumentation used to collect data was a single wire driven by a constant temperature anemometer bridge. The hot wire is specially modified to measure data near the cylindrical leading edges curved surface. The traversing system allowed the acquisition of high-resolution boundary layer data. The traversing system was installed internally to the cylindrical leading edge to reduce probe blockage.

  13. Multiple element airfoils optimized for maximum lift coefficient.

    NASA Technical Reports Server (NTRS)

    Ormsbee, A. I.; Chen, A. W.

    1972-01-01

    Optimum airfoils in the sense of maximum lift coefficient are obtained for incompressible fluid flow at large Reynolds number. The maximum lift coefficient is achieved by requiring that the turbulent skin friction be zero in the pressure rise region on the airfoil upper surface. Under this constraint, the pressure distribution is optimized. The optimum pressure distribution is a function of Reynolds number and the trailing edge velocity. Geometries of those airfoils which will generate these optimum pressure distributions are obtained using a direct-iterative method which is developed in this study. This method can be used to design airfoils consisting of any number of elements. Numerical examples of one- and two-element airfoils are given. The maximum lift coefficients obtained range from 2 to 2.5.

  14. Simultaneous monitoring of ice accretion and thermography of an airfoil: an IR imaging methodology

    NASA Astrophysics Data System (ADS)

    Mohseni, M.; Frioult, M.; Amirfazli, A.

    2012-10-01

    A novel image analysis methodology based on infrared (IR) imaging was developed for simultaneous monitoring of ice accretion and thermography of airfoils. In this study, an IR camera was calibrated and used to measure the surface temperature of the energized airfoils, and monitor the ice accretion and growth pattern on the airfoils’ surfaces. The methodology comprises the automatic processing of a series of IR video frames with the purpose of detecting ice pattern evolution during the icing test period. A specially developed MATLAB code was used to detect the iced areas in the IR images, and simultaneously monitor surface temperature evolution of the airfoil during an icing test. Knowing the correlation between the icing pattern and surface temperature changes during an icing test is essential for energy efficient design of thermal icing mitigation systems. Processed IR images were also used to determine the ice accumulation rate on the airfoil's surface in a given icing test. The proposed methodology has been demonstrated to work successfully, since the optical images taken at the end of icing tests from the airfoils’ surfaces compared well with the processed IR images detecting the ice grown outward from the airfoils’ leading edge area.

  15. 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.

  16. 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.

  17. 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.

  18. Flow characteristics of infinite-span wings with wavy leading edges

    NASA Astrophysics Data System (ADS)

    Perez-Torro, Rafael; Kim, Jae-Wook

    2016-11-01

    Implicit LES computations are performed for an infinite-span wing based on the NACA0021 aerofoil section with a sinusoidal wavy leading edge (WLE). At Re∞ = 1 . 2 ×105 and M∞ = 0 . 3 , the computations performed in this study show that three-dimensional laminar separation bubbles (LSBs) form at troughs of the undulated wing. Prior to stall, LSBs can be found in all troughs. However, past the stall angle, LSBs tend to group together in a collocated fashion, leaving regions of complete separation in between groups where a separated shear layer (SSL) is formed. It is found that the size of the LSB group is highly dependent on the number of WLE wavelengths used in the spanwise-periodic domain. The LSB group formation process is investigated by means of simulations where the geometry is slowly pitched from an angle of attack of α =10° to α =20° . The study also includes the analysis of instantaneous flow fields using Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD) techniques. The authors acknowledge the HPC facilities of the UK National Supercomputer Archer via the support of the UK Turbulence Consortium (EP/L000261/1) and the local Iridis4 at the University of Southampton.

  19. The ubiquitin-proteasome system regulates focal adhesions at the leading edge of migrating cells

    PubMed Central

    Teckchandani, Anjali; Cooper, Jonathan A

    2016-01-01

    Cell migration requires the cyclical assembly and disassembly of focal adhesions. Adhesion induces phosphorylation of focal adhesion proteins, including Cas (Crk-associated substrate/p130Cas/BCAR1). However, Cas phosphorylation stimulates adhesion turnover. This raises the question of how adhesion assembly occurs against opposition from phospho-Cas. Here we show that suppressor of cytokine signaling 6 (SOCS6) and Cullin 5, two components of the CRL5SOCS6 ubiquitin ligase, inhibit Cas-dependent focal adhesion turnover at the front but not rear of migrating epithelial cells. The front focal adhesions contain phospho-Cas which recruits SOCS6. If SOCS6 cannot access focal adhesions, or if cullins or the proteasome are inhibited, adhesion disassembly is stimulated. This suggests that the localized targeting of phospho-Cas within adhesions by CRL5SOCS6 and concurrent cullin and proteasome activity provide a negative feedback loop, ensuring that adhesion assembly predominates over disassembly at the leading edge. By this mechanism, ubiquitination provides a new level of spatio-temporal control over cell migration. DOI: http://dx.doi.org/10.7554/eLife.17440.001 PMID:27656905

  20. 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.

  1. 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].

  2. Generation of Fullspan Leading-Edge 3D Ice Shapes for Swept-Wing Aerodynamic Testing

    NASA Technical Reports Server (NTRS)

    Camello, Stephanie C.; Lee, Sam; Lum, Christopher; Bragg, Michael B.

    2016-01-01

    The deleterious effect of ice accretion on aircraft is often assessed through dry-air flight and wind tunnel testing with artificial ice shapes. This paper describes a method to create fullspan swept-wing artificial ice shapes from partial span ice segments acquired in the NASA Glenn Icing Reserch Tunnel for aerodynamic wind-tunnel testing. Full-scale ice accretion segments were laser scanned from the Inboard, Midspan, and Outboard wing station models of the 65% scale Common Research Model (CRM65) aircraft configuration. These were interpolated and extrapolated using a weighted averaging method to generate fullspan ice shapes from the root to the tip of the CRM65 wing. The results showed that this interpolation method was able to preserve many of the highly three dimensional features typically found on swept-wing ice accretions. The interpolated fullspan ice shapes were then scaled to fit the leading edge of a 8.9% scale version of the CRM65 wing for aerodynamic wind-tunnel testing. Reduced fidelity versions of the fullspan ice shapes were also created where most of the local three-dimensional features were removed. The fullspan artificial ice shapes and the reduced fidelity versions were manufactured using stereolithography.

  3. 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.

  4. 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.

  5. Modeling Shock Train Leading Edge Detection in Dual-Mode Scramjets

    NASA Astrophysics Data System (ADS)

    Ladeinde, Foluso; Lou, Zhipeng; Li, Wenhai

    2016-11-01

    The objective of this study is to accurately model the detection of shock train leading edge (STLE) in dual-mode scramjet (DMSJ) engines intended for hypersonic flight in air-breathing propulsion systems. The associated vehicles have applications in military warfare and intelligence, and there is commercial interest as well. Shock trains are of interest because they play a significant role in the inability of a DMSJ engine to develop the required propulsive force. The experimental approach to STLE detection has received some attention; as have numerical calculations. However, virtually all of the numerical work focus on mechanically- (i.e., pressure-) generated shock trains, which are much easier to model relative to the phenomenon in the real system where the shock trains are generated by combustion. A focus on combustion, as in the present studies, enables the investigation of the effects of equivalence ratio, which, together with the Mach number, constitutes an important parameter determining mode transition. The various numerical approaches implemented in our work will be reported, with result comparisons to experimental data. The development of an STLE detection procedure in an a priori manner will also be discussed.

  6. 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.

  7. A practical study of the aerodynamic impact of wind turbine blade leading edge erosion

    NASA Astrophysics Data System (ADS)

    Gaudern, N.

    2014-06-01

    During operation wind turbine blades are exposed to a wide variety of atmospheric and environmental conditions; inspection reports for blades that have been operating for several years show varying degrees of leading edge erosion. It is important to be able to estimate the impact of different stages of erosion on wind turbine performance, but this is very difficult even with advanced CFD models. In this study, wind tunnel testing was used to evaluate a range of complex erosion stages. Erosion patterns were transferred to thin films that were applied to 18% thick commercial wind turbine aerofoils and full lift and drag polars were measured in a wind tunnel. Tests were conducted up to a Reynolds number of 2.20 × 106 scaling based on the local roughness Reynolds number was used in combination with different film thicknesses to simulate a variety of erosion depths. The results will be very useful for conducting cost/benefit analyses of different methods of blade protection and repair, as well as for defining the appropriate timescales for these processes.

  8. Evaluation of Navier-Stokes and Euler solutions for leading-edge separation vortices

    NASA Technical Reports Server (NTRS)

    Fujii, K.; Gavali, S.; Holst, T. L.

    1987-01-01

    Extensive study on the numerical simulation of the vortical flow over a double delta wing is carried out using the thin layer Navier-Stokes and Euler equations. Two important flow characteristics, vortex interaction and vortex breakdown, are successfully simulated. Grid resolution is one of the most important factors associated with the vortex problem. Computations were performed on a series of grids with various levels of refinement, coarse, medium, and fine. Computations using either the coarse or medium grids fail to capture the proper physical phenomena. The computed result using a fine grid shows flow unsteadiness once the vortex breakdown takes place. The C sub L - alpha characteristics are well predicted up to the breakdown angle of attack for all the grid distributions. The Euler solutions show fairly good agreement with the experiment on the C sub L - alpha characteristics. However, other aspects of the solution at each angle of attack, such as the locus of the leading edge separation vortex, are not consistent with the experiment. Even for the fine grid Navier-Stokes computations, further grid resolution is required to obtain good quantitative agreement with the experiment.

  9. 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.

  10. 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.

  11. 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.

  12. 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.

  13. 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.

  14. Clathrin regulates lymphocyte migration by driving actin accumulation at the cellular leading edge.

    PubMed

    Ramírez-Santiago, Guillermo; Robles-Valero, Javier; Morlino, Giulia; Cruz-Adalia, Aranzazu; Pérez-Martínez, Manuel; Zaldivar, Airen; Torres-Torresano, Mónica; Chichón, Francisco Javier; Sorrentino, Andrea; Pereiro, Eva; Carrascosa, José L; Megías, Diego; Sorzano, Carlos Oscar S; Sánchez-Madrid, Francisco; Veiga, Esteban

    2016-10-01

    Lymphocyte migration, which is essential for effective immune responses, belongs to the so-called amoeboid migration. The lymphocyte migration is up to 100 times faster than between mesenchymal and epithelial cell types. Migrating lymphocytes are highly polarized in three well-defined structural and functional zones: uropod, medial zone, and leading edge (LE). The actiomyosin-dependent driving force moves forward the uropod, whereas massive actin rearrangements protruding the cell membrane are observed at the LE. These actin rearrangements resemble those observed at the immunological synapse driven by clathrin, a protein normally involved in endocytic processes. Here, we used cell lines as well as primary lymphocytes to demonstrate that clathrin and clathrin adaptors colocalize with actin at the LE of migrating lymphocytes, but not in other cellular zones that accumulate both clathrin and actin. Moreover, clathrin and clathrin adaptors, including Hrs, the clathrin adaptor for multivesicular bodies, drive local actin accumulation at the LE. Clathrin recruitment at the LE resulted necessary for a complete cell polarization and further lymphocyte migration in both 2D and 3D migration models. Therefore, clathrin, including the clathrin population associated to internal vesicles, controls lymphocyte migration by regulating actin rearrangements occurring at the LE.

  15. The effect of butterfly-scale inspired patterning on leading-edge vortex growth

    NASA Astrophysics Data System (ADS)

    Wilroy, Jacob; Lang, Amy

    2015-11-01

    Leading edge vortices (LEVs) are important for generating thrust and lift in flapping flight, and the surface patterning (scales) on butterfly wings is hypothesized to play a role in the vortex formation of the LEV. To simplify this complex flow problem, an experiment was designed to focus on the alteration of 2-D vortex development with a variation in surface patterning. Specifically, the secondary vorticity generated by the LEV interacting at the patterned surface was studied and the subsequent affect on the growth rate of the circulation in the LEV. For this experiment we used butterfly inspired grooves attached to a flat plate and compared the vortex formation to a smooth plate case as the plate moved vertically. The plate is impulsively started in quiescent water and flow fields at Re = 1500, 3000, and 6000 are examined using Digital Particle Image Velocimetry (DPIV). The vortex formation time is 3.0 and is based on the flat plate travel length and chord length. We would like to thank the National Science Foundation REU Site Award 1358991 for funding this research.

  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. Effects of forward contour modification on the aerodynamic characteristics of the NACA 641-212 airfoil section

    NASA Technical Reports Server (NTRS)

    Hicks, R. M.; Mendoza, J. P.; Bandettini, A.

    1975-01-01

    Two different forward contour modifications designed to increase the maximum lift coefficient of the NACA 64 sub 1-212 airfoil section were evaluated experimentally at low speeds. One modification consisted of a slight droop of the leading edge with an increased leading-edge radius; the other modification incorporated increased thickness over the forward 35 percent of the upper surface of the profile. Both modified airfoil sections were found to provide substantially higher maximum lift coefficients than the 64 sub 1-212 section. The drooped leading-edge modification incurred a drag penalty of approximately 10 percent at low and moderate lift coefficients and exhibited a greater nosedown pitching moment than the 64 sub 1-212 profile. The upper surface modification produced about the same drag level as the 64 sub 1-212 section at low and moderate lift coefficients and less nosedown pitching moment than the 64 sub 1-212 profile. Both modified airfoil sections had lower drag coefficients than the 64 sub 1-212 section at high lift coefficients.

  18. Signalling crosstalk at the leading edge controls tissue closure dynamics in the Drosophila embryo

    PubMed Central

    Carballès, Fabrice; Parassol, Nadège; Schaub, Sébastien; Cérézo, Delphine; Noselli, Stéphane

    2017-01-01

    Tissue morphogenesis relies on proper differentiation of morphogenetic domains, adopting specific cell behaviours. Yet, how signalling pathways interact to determine and coordinate these domains remains poorly understood. Dorsal closure (DC) of the Drosophila embryo represents a powerful model to study epithelial cell sheet sealing. In this process, JNK (JUN N-terminal Kinase) signalling controls leading edge (LE) differentiation generating local forces and cell shape changes essential for DC. The LE represents a key morphogenetic domain in which, in addition to JNK, a number of signalling pathways converges and interacts (anterior/posterior -AP- determination; segmentation genes, such as Wnt/Wingless; TGFβ/Decapentaplegic). To better characterize properties of the LE morphogenetic domain, we sought out new JNK target genes through a genomic approach: 25 were identified of which 8 are specifically expressed in the LE, similarly to decapentaplegic or puckered. Quantitative in situ gene profiling of this new set of LE genes reveals complex patterning of the LE along the AP axis, involving a three-way interplay between the JNK pathway, segmentation and HOX genes. Patterning of the LE into discrete domains appears essential for coordination of tissue sealing dynamics. Loss of anterior or posterior HOX gene function leads to strongly delayed and asymmetric DC, due to incorrect zipping in their respective functional domain. Therefore, in addition to significantly increasing the number of JNK target genes identified so far, our results reveal that the LE is a highly heterogeneous morphogenetic organizer, sculpted through crosstalk between JNK, segmental and AP signalling. This fine-tuning regulatory mechanism is essential to coordinate morphogenesis and dynamics of tissue sealing. PMID:28231245

  19. Existence of and decay to equilibrium of the filament end density along the leading edge of the lamellipodium.

    PubMed

    Manhart, Angelika; Schmeiser, Christian

    2017-01-01

    A model for the dynamics of actin filament ends along the leading edge of the lamellipodium is analyzed. It contains accounts of nucleation by branching, of deactivation by capping, and of lateral flow along the leading edge by polymerization. A nonlinearity arises from a Michaelis-Menten type modeling of the branching process. For branching rates large enough compared to capping rates, the existence and stability of nontrivial steady states is investigated. The main result is exponential convergence to nontrivial steady states, proven by investigating the decay of an appropriate Lyapunov functional.

  20. Mass loss of TEOS-coated RCC subjected to the environment at the shuttle wing leading edge

    NASA Technical Reports Server (NTRS)

    Stroud, C. W.; Rummler, D. R.

    1981-01-01

    Coated, reinforced carbon-carbon (RCC) is used for the leading edges of the Space Shuttle. The mass loss characteristics of RCC specimens coated with tetra-ethyl-ortho-silicate (TEOS) were determined for conditions which simulated the entry environment expected at the stagnation area of the wing leading edge. Maximum specimen temperature was 1632 K. Specimens were exposed for up to 100 missions. Stress levels up to 8.274 MPa caused an average increase in oxidation of 6 percent over unstressed specimens. Experimentally determined mass losses were compared with those predicted by an existing empirical analysis.

  1. Evaluation of leading- and trailing-edge flaps on flat and cambered delta wings at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Hernandez, Gloria; Wood, Richard M.; Collins, Robert E.

    1989-01-01

    An experimental investigation has been conducted to evaluate the effectiveness of leading- and trailing-edge flaps on a flat and cambered wing at superconic speeds. Results from the experimental tests showed that highly complex and three-dimensional flow can occur over the wings with leading- and/or trailing-edge flaps deflected. An analysis of the data also showed that flap effectiveness varies significantly between a cambered and flat wing of identical planform and flap geometry. Mach number effects are similar for both flat and cambered wings for all aerodynamic parameters.

  2. 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.

  3. 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.

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. Petiolate wings: effects on the leading-edge vortex in flapping flight

    PubMed Central

    2017-01-01

    The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1–3. The wings were driven using a mechanical device, the ‘Flapperatus’, to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ* (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested. PMID:28163876

  9. 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.

  10. Petiolate wings: effects on the leading-edge vortex in flapping flight.

    PubMed

    Phillips, Nathan; Knowles, Kevin; Bomphrey, Richard J

    2017-02-06

    The wings of many insect species including crane flies and damselflies are petiolate (on stalks), with the wing planform beginning some distance away from the wing hinge, rather than at the hinge. The aerodynamic impact of flapping petiolate wings is relatively unknown, particularly on the formation of the lift-augmenting leading-edge vortex (LEV): a key flow structure exploited by many insects, birds and bats to enhance their lift coefficient. We investigated the aerodynamic implications of petiolation P using particle image velocimetry flow field measurements on an array of rectangular wings of aspect ratio 3 and petiolation values of P = 1-3. The wings were driven using a mechanical device, the 'Flapperatus', to produce highly repeatable insect-like kinematics. The wings maintained a constant Reynolds number of 1400 and dimensionless stroke amplitude Λ* (number of chords traversed by the wingtip) of 6.5 across all test cases. Our results showed that for more petiolate wings the LEV is generally larger, stronger in circulation, and covers a greater area of the wing surface, particularly at the mid-span and inboard locations early in the wing stroke cycle. In each case, the LEV was initially arch-like in form with its outboard end terminating in a focus-sink on the wing surface, before transitioning to become continuous with the tip vortex thereafter. In the second half of the wing stroke, more petiolate wings exhibit a more detached LEV, with detachment initiating at approximately 70% and 50% span for P = 1 and 3, respectively. As a consequence, lift coefficients based on the LEV are higher in the first half of the wing stroke for petiolate wings, but more comparable in the second half. Time-averaged LEV lift coefficients show a general rise with petiolation over the range tested.

  11. Characterization of multifunctional skin-material for morphing leading-edge applications

    NASA Astrophysics Data System (ADS)

    Geier, Sebastian; Kintscher, Markus; Mahrholz, Thorsten; Wierach, Peter; Monner, Hans-Peter; Wiedemann, Martin

    2013-04-01

    Former research on morphing droop-nose applications revealed great economical and social ecological advantages in terms of providing gapless surfaces for long areas of laminar flow. Furthermore a droop-nose for laminar flow applications provides a low noise exposing high-lift system at the leading-edge. Various kinematic concepts for the active deployment of such devices are already published but the major challenge is still an open issue: a skin material which meets the compromise of needed stiffness and flexibility. Moreover additional functions have to be added to keep up with standard systems. As a result of several national and European projects the DLR developed a gapless 3D smart droop-nose concept, which was successfully analyzed in a low speed wind tunnel test under relevant loads to prove the functionality and efficiency. The main structure of this concept is made of commercial available glass fiber reinforced plastics (GRFP). This paper presents elementary tests to characterize material lay-ups and their integrity by applying different loads under extreme thermal conditions using aged specimens. On the one hand the presented work is focused on the integrity of material-interfaces and on the other hand the efficiency and feasibility of embedded functions. It can be concluded that different preparations, different adhesives and used materials have their significant influence to the interface stability and mechanical property of the whole lay-up. Especially the laminate design can be optimized due to the e. g. mechanical exploitation of the added systems beyond their main function in order to reduce structural mass.

  12. LOW SUBSONIC PRESSURE DISTRIBUTIONS ON THREE RIGID WINGS SIMULATING PARAGLIDERS WITH VARIED CANOPY CURVATURE AND LEADING-EDGE SWEEP

    DTIC Science & Technology

    An investigation was made in the Langley 7- by 10-foot transonic tunnel to determine the subsonic pressure distribution of three paraglider models...through an angle-of-attack range from 0 to 74 degrees. Three rigid meta models simulated a 45 degrees basic flat planform paraglider with leading-edge

  13. Heat transfer characteristics of hypersonic waveriders with an emphasis on the leading edge effects. M.S. Thesis, 1991

    NASA Technical Reports Server (NTRS)

    Vanmol, Denis O.; Anderson, John D., Jr.

    1992-01-01

    The heat transfer characteristics in surface radiative equilibrium and the aerodynamic performance of blunted hypersonic waveriders are studied along two constant dynamic pressure trajectories for four different Mach numbers. The inviscid leading edge drag was found to be a small (4 to 8 percent) but not negligible fraction of the inviscid drag of the vehicle. Although the viscous drag at the leading edge can be neglected, the presence of the leading edge will influence the transition pattern of the upper and the lower surfaces and therefore affect the viscous drag of the entire vehicle. For an application similar to the National Aerospace Plane (NASP), the present study demonstrates that the waverider remains a valuable concept at high Mach numbers if a state-of-the-art active cooling device is used along the leading edge. At low Mach number (less than 5), the study shows the surface radiative cooling might be sufficient. In all cases, radiative cooling is sufficient for the upper and lower surfaces of the vehicle if ceramic composites are used as thermal protection.

  14. The Effect of Leading-Edge Sweep and Surface Inclination on the Hypersonic Flow Field Over a Blunt Flat Plate

    NASA Technical Reports Server (NTRS)

    Creager, Marcus O.

    1959-01-01

    An investigation of the effects of variation of leading-edge sweep and surface inclination on the flow over blunt flat plates was conducted at Mach numbers of 4 and 5.7 at free-stream Reynolds numbers per inch of 6,600 and 20,000, respectively. Surface pressures were measured on a flat plate blunted by a semicylindrical leading edge over a range of sweep angles from 0 deg to 60 deg and a range of surface inclinations from -10 deg to +10 deg. The surface pressures were predicted within an average error of +/- 8 percent by a combination of blast-wave and boundary-layer theory extended herein to include effects of sweep and surface inclination. This combination applied equally well to similar data of other investigations. The local Reynolds number per inch was found to be lower than the free-stream Reynolds number per inch. The reduction in local Reynolds number was mitigated by increasing the sweep of the leading edge. Boundary-layer thickness and shock-wave shape were changed little by the sweep of the leading edge.

  15. 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.

  16. 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.

  17. "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…

  18. Active aerodynamic control of wake-airfoil interaction noise - Experiment

    NASA Astrophysics Data System (ADS)

    Simonich, J. C.; Lavrich, P. L.; Sofrin, T. G.; Topol, D. A.

    A proof of concept experiment is conducted that shows the potential for active aerodynamic control of rotor wake/stator interaction noise in a simplified manner. A single airfoil model representing the stator was fitted with a moveable trailing edge flap controlled by a servo motor. The control system moves the motor driven flap in the correct angular displacement phase and rate to reduce the unsteady load on the airfoil during the wake interaction.

  19. 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.

  20. 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.