Preliminary Investigation in the NACA Low-Turbulence Tunnel of Low-drag Airfoil Sections Suitable for Admitting Air at the Leading Edge

NASA Technical Reports Server (NTRS)

von Doenhoff, Albert E.; Horton, Elmer A.

1942-01-01

An investigation was carried out in the NACA low-turbulence tunnel to develop low-drag airfoil sections suitable for admitting air at the leading edge. A thickness distribution having the desired type of pressure distribution was found from tests of a flexible model. Other airfoil shapes were derived from this original shape by varying the thickness, the camper, the leading-edge radius, and the size of the leading-edge opening. Data are presented giving the characteristics of the airfoil shapes in the range of lift coefficients for high-speed and cruising flight. Shapes have been developed which show no substantial increases in drag over that of the same position along the chord. Many of these shapes appear to have higher critical compressibility speeds than plain airfoils of the same thickness. Low-drag airfoil sections have been developed with openings in the leading edge as large as 41.5 percent of the maximum thickness. The range of lift coefficients for low drag in several cases is nearly as large as that of the corresponding plain airfoil sections. Preliminary measurements of maximum lift characteristics indicate that nose-opening sections of the type herein considered may not produce any marked effects on the maximum lift coefficient.

Numerical evaluations of the effect of leading-edge protuberances on the static and dynamic stall characteristics of an airfoil

NASA Astrophysics Data System (ADS)

Cai, C.; Zuo, Z. G.; Liu, S. H.; Wu, Y. L.; Wang, F. B.

2013-12-01

Wavy leading edge modifications of airfoils through imitating humpback whale flippers has been considered as a viable passive way to control flow separation. In this paper, flows around a baseline 634-021 airfoil and one with leading-edge sinusoidal protuberances were simulated using S-A turbulence model. When studying the static stall characteristics, it is found that the modified airfoil does not stall in the traditional manner, with increasing poststall lift coefficients. At high angles of attack, the flows past the wavy leading edge stayed attached for a distance, while the baseline foil is in a totally separated flow condition. On this basis, the simulations of pitch characteristic were carried out for both foils. At high angles of attack mild variations in lift and drag coefficients of the modified foil can be found, leading to a smaller area of hysteresis loop. The special structure of wavy leading edge can help maintain high consistency of the flow field in dynamic pitching station within a particular range of angles of attack.

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

NASA Astrophysics Data System (ADS)

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

2014-04-01

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

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.

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.

CFD study on NACA 4415 airfoil implementing spherical and sinusoidal Tubercle Leading Edge

PubMed Central

2017-01-01

The Humpback whale tubercles have been studied for more than a decade. Tubercle Leading Edge (TLE) effectively reduces the separation bubble size and helps in delaying stall. They are very effective in case of low Reynolds number flows. The current Computational Fluid Dynamics (CFD) study is on NACA 4415 airfoil, at a Reynolds number 120,000. Two TLE shapes are tested on NACA 4415 airfoil. The tubercle designs implemented on the airfoil are sinusoidal and spherical. A parametric study is also carried out considering three amplitudes (0.025c, 0.05c and 0.075c), the wavelength (0.25c) is fixed. Structured mesh is utilized to generate grid and Transition SST turbulence model is used to capture the flow physics. Results clearly show spherical tubercles outperform sinusoidal tubercles. Furthermore experimental study considering spherical TLE is carried out at Reynolds number 200,000. The experimental results show that spherical TLE improve performance compared to clean airfoil. PMID:28850622

CFD study on NACA 4415 airfoil implementing spherical and sinusoidal Tubercle Leading Edge.

PubMed

Aftab, S M A; Ahmad, K A

2017-01-01

The Humpback whale tubercles have been studied for more than a decade. Tubercle Leading Edge (TLE) effectively reduces the separation bubble size and helps in delaying stall. They are very effective in case of low Reynolds number flows. The current Computational Fluid Dynamics (CFD) study is on NACA 4415 airfoil, at a Reynolds number 120,000. Two TLE shapes are tested on NACA 4415 airfoil. The tubercle designs implemented on the airfoil are sinusoidal and spherical. A parametric study is also carried out considering three amplitudes (0.025c, 0.05c and 0.075c), the wavelength (0.25c) is fixed. Structured mesh is utilized to generate grid and Transition SST turbulence model is used to capture the flow physics. Results clearly show spherical tubercles outperform sinusoidal tubercles. Furthermore experimental study considering spherical TLE is carried out at Reynolds number 200,000. The experimental results show that spherical TLE improve performance compared to clean airfoil.

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.

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.

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

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.

Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partical-span Leading-edge Slat for Various Modes of Thermal Ice Protection

NASA Technical Reports Server (NTRS)

Von Glahn, Uwe H; Gray, Vernon H

1954-01-01

Studies were made to determine the effect of ice formations on the section drag of a 6.9-foot-chord 36 degree swept NACA 63A-009 airfoil with partial-span leading-edge slat. In general, the icing of a thin swept airfoil will result in greater aerodynamic penalties than for a thick unswept airfoil. Glaze-ice formations at the leading edge of the airfoil caused large increases in section drag even at liquid-water content of 0.39 gram per cubic meter. The use of an ice-free parting strip in the stagnation region caused a negligible change in drag compared with a completely unheated airfoil. Cyclic de-icing when properly applied caused the drag to decrease almost to the bare-airfoil drag value.

Study on Trailing Edge Ramp of Supercritical Airfoil

DTIC Science & Technology

2016-03-30

7 th Asia-Pacific International Symposium on Aerospace Technology, 25 – 27 November 2015, Cairns Study on Trailing Edge Ramp of Supercritical...China Abstract Trailing edge flow control method could improve the performance of supercritical airfoil with a small modification on the original...airfoil. In this paper, a ramp of 2%~7% chord length is sliced near the trailing edge to improve airfoil performance. The trailing edge ramp is

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

RANS Based Methodology for Predicting the Influence of Leading Edge Erosion on Airfoil Performance

DOE Office of Scientific and Technical Information (OSTI.GOV)

Langel, Christopher M.; Chow, Raymond C.; van Dam, C. P.

The impact of surface roughness on flows over aerodynamically designed surfaces is of interested in a number of different fields. It has long been known the surface roughness will likely accelerate the laminar- turbulent transition process by creating additional disturbances in the boundary layer. However, there are very few tools available to predict the effects surface roughness will have on boundary layer flow. There are numerous implications of the premature appearance of a turbulent boundary layer. Increases in local skin friction, boundary layer thickness, and turbulent mixing can impact global flow properties compounding the effects of surface roughness. With thismore » motivation, an investigation into the effects of surface roughness on boundary layer transition has been conducted. The effort involved both an extensive experimental campaign, and the development of a high fidelity roughness model implemented in a R ANS solver. Vast a mounts of experimental data was generated at the Texas A&M Oran W. Nicks Low Speed Wind Tunnel for the calibration and validation of the roughness model described in this work, as well as future efforts. The present work focuses on the development of the computational model including a description of the calibration process. The primary methodology presented introduces a scalar field variable and associated transport equation that interacts with a correlation based transition model. The additional equation allows for non-local effects of surface roughness to be accounted for downstream of rough wall sections while maintaining a "local" formulation. The scalar field is determined through a boundary condition function that has been calibrated to flat plate cases with sand grain roughness. The model was initially tested on a NACA 0012 airfoil with roughness strips applied to the leading edge. Further calibration of the roughness model was performed using results from the companion experimental study on a NACA 63 3 -418 airfoil

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.

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

NASA Astrophysics Data System (ADS)

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

2013-11-01

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

Mechanism of Supercooled Water Droplet Breakup near the Leading Edge of an Airfoil

NASA Technical Reports Server (NTRS)

Veras-Alba, Belen; Palacios, Jose; Vargas, Mario; Ruggeri, Charles; Bartkus, Tadas P.

2017-01-01

This work presents the results of an experimental study on supercooled droplet deformation and breakup near the leading edge of an airfoil. The results are compared to prior room temperature droplet deformation results to explore the effects of droplet supercooling. The experiments were conducted in the Adverse Environment Rotor Test Stand (AERTS) at The Pennsylvania State University. An airfoil model placed at the end of the rotor blades mounted onto the hub in the AERTS chamber was moved at speeds ranging between 50 and 80 m/sec. The temperature of the chamber was set at -20°C. A monotonic droplet generator was used to produce droplets that fell from above, perpendicular to the path of the airfoil. The supercooled state of the droplets was determined by measurement of the temperature of the drops at various locations below the droplet generator exit. A temperature prediction code was also used to estimate the temperature of the droplets based on vertical velocity and the distance traveled by droplets from the droplet generator to the airfoil stagnation line. High speed imaging was employed to observe the interaction between the droplets and the airfoil. The high speed imaging provided droplet deformation information as the droplet approached the airfoil near the stagnation line. A tracking software program was used to measure the horizontal and vertical displacement of the droplet against time. It was demonstrated that to compare the effects of water supercooling on droplet deformation, the ratio of the slip velocity and the initial droplet velocity must be equal. A case with equal slip velocity to initial velocity ratios was selected for room temperature and supercooled droplet conditions. The airfoil velocity was 60 m/s and the slip velocity for both sets of data was 40 m/s. In these cases, the deformation of the weakly supercooled and warm droplets did not present different trends. The similar behavior for both environmental conditions indicates that water

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

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

Wind Tunnel and Numerical Analysis of Thick Blunt Trailing Edge Airfoils

NASA Astrophysics Data System (ADS)

McLennan, Anthony William

Two-dimensional aerodynamic characteristics of several thick blunt trailing edge airfoils are presented. These airfoils are not only directly applicable to the root section of wind turbine blades, where they provide the required structural strength at a fraction of the material and weight of an equivalent sharp trailing edge airfoil, but are also applicable to the root sections of UAVs having high aspect ratios, that also encounter heavy root bending forces. The Reynolds averaged Navier-Stokes code, ARC2D, was the primary numerical tool used to analyze each airfoil. The UCD-38-095, referred to as the Pareto B airfoil in this thesis, was also tested in the University of California, Davis Aeronautical Wind Tunnel. The Pareto B has an experimentally determined maximum lift coefficient of 1.64 at 14 degrees incidence, minimum drag coefficient of 0.0385, and maximum lift over drag ratio of 35.9 at a lift coefficient of 1.38, 10 degrees incidence at a Reynolds number of 666,000. Zig-zag tape at 2% and 5% of the chord was placed on the leading edge pressure and suction side of the Pareto B model in order to determine the aerodynamic performance characteristics at turbulent flow conditions. Experimental Pareto B wind tunnel data and previous FB-3500-0875 data is also presented and used to validate the ARC2D results obtained in this study. Additionally MBFLO, a detached eddy simulation Navier-Stokes code, was used to analyze the Pareto B airfoil for comparison and validation purposes.

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.

Flow visualization over a thick blunt trailing-edge airfoil with base cavity at low Reynolds numbers using PIV technique.

PubMed

Taherian, Gholamhossein; Nili-Ahmadabadi, Mahdi; Karimi, Mohammad Hassan; Tavakoli, Mohammad Reza

2017-01-01

In this study, the effect of cutting the end of a thick airfoil and adding a cavity on its flow pattern is studied experimentally using PIV technique. First, by cutting 30% chord length of the Riso airfoil, a thick blunt trialing-edge airfoil is generated. The velocity field around the original airfoil and the new airfoil is measured by PIV technique and compared with each other. Then, adding two parallel plates to the end of the new airfoil forms the desired cavity. Continuous measurement of unsteady flow velocity over the Riso airfoil with thick blunt trailing edge and base cavity is the most important innovation of this research. The results show that cutting off the end of the airfoil decreases the wake region behind the airfoil, when separation occurs. Moreover, adding a cavity to the end of the thickened airfoil causes an increase in momentum and a further decrease in the wake behind the trailing edge that leads to a drag reduction in comparison with the thickened airfoil without cavity. Furthermore, using cavity decreases the Strouhal number and vortex shedding frequency.

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.

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.

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

NASA Astrophysics Data System (ADS)

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

2015-07-01

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

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.

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.

Development of a design model for airfoil leading edge film cooling

NASA Astrophysics Data System (ADS)

Wadia, A. R.; Nealy, D. A.

1985-03-01

A series of experiments on scaled cylinder models having injection through holes inclined at 20, 30, 45, and 90 degrees are presented. The experiments were conducted in a wind tunnel on several stainless steel test specimens in which flow and heat transfer parameters were measured over simulated airfoil leading edge surfaces. On the basis of the experimental results, an engineering design model is proposed that treats the gas-to-surface heat transfer coefficient with film cooling in a manner suggested by Luckey and L'Ecuyer (1981). It is shown that the main factor influencing the averaged film cooling effectiveness in the showerhead region is the inclination of the injection holes. The effectiveness parameter was not affected by variations in the coolant-to-gas stream pressure ratio, the freestream Mach number, the gas to coolant temperature ratio, or the gas stream Reynolds number. Experience in the wind tunnel tests is reflected in the design of the model in which the coolant side heat transfer coefficient is offset by a simultaneous increase in the gas side film coefficient. The design applications of the analytical model are discussed, with emphasis given to high temperature first stage turbine vanes and rotor blades.

Experimental investigation of trailing edge noise from stationary and rotating airfoils.

PubMed

Zajamsek, Branko; Doolan, Con J; Moreau, Danielle J; Fischer, Jeoffrey; Prime, Zebb

2017-05-01

Trailing edge noise from stationary and rotating NACA 0012 airfoils is characterised and compared with a noise prediction based on the semi-empirical Brooks, Pope, and Marcolini (BPM) model. The NACA 0012 is symmetrical airfoil with no camber and 12% thickness to chord length ratio. Acoustic measurements were conducted in an anechoic wind tunnel using a stationary NACA 0012 airfoil at 0° pitch angle. Airfoil self-noise emissions from rotating NACA 0012 airfoils mounted at 0° and 10° pitch angles on a rotor-rig are studied in an anechoic room. The measurements were carried out using microphone arrays for noise localisation and magnitude estimation using beamforming post-processing. Results show good agreement between peak radiating trailing edge noise emissions of stationary and rotating NACA 0012 airfoils in terms of the Strouhal number. Furthermore, it is shown that noise predictions based on the BPM model considering only two dimensional flow effects, are in good agreement with measurements for rotating airfoils, at these particular conditions.

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.

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.

Airfoil

DOEpatents

Ristau, Neil; Siden, Gunnar Leif

2015-07-21

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

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.

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.

Experimental investigation of trailing edge noise from stationary and rotating airfoils

PubMed Central

Zajamsek, Branko; Doolan, Con J.; Moreau, Danielle J.; Fischer, Jeoffrey; Prime, Zebb

2017-01-01

Trailing edge noise from stationary and rotating NACA 0012 airfoils is characterised and compared with a noise prediction based on the semi-empirical Brooks, Pope, and Marcolini (BPM) model. The NACA 0012 is symmetrical airfoil with no camber and 12% thickness to chord length ratio. Acoustic measurements were conducted in an anechoic wind tunnel using a stationary NACA 0012 airfoil at 0° pitch angle. Airfoil self-noise emissions from rotating NACA 0012 airfoils mounted at 0° and 10° pitch angles on a rotor-rig are studied in an anechoic room. The measurements were carried out using microphone arrays for noise localisation and magnitude estimation using beamforming post-processing. Results show good agreement between peak radiating trailing edge noise emissions of stationary and rotating NACA 0012 airfoils in terms of the Strouhal number. Furthermore, it is shown that noise predictions based on the BPM model considering only two dimensional flow effects, are in good agreement with measurements for rotating airfoils, at these particular conditions. PMID:28599535

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.

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.

Airfoil noise reductions through leading edge serrations

NASA Astrophysics Data System (ADS)

Narayanan, S.; Chaitanya, P.; Haeri, S.; Joseph, P.; Kim, J. W.; Polacsek, C.

2015-02-01

This paper provides an experimental investigation into the use of leading edge (LE) serrations as a means of reducing the broadband noise generated due to the interaction between the aerofoil's LE and impinging turbulence. Experiments are performed on a flat plate in an open jet wind tunnel. Grids are used to generate isotropic homogeneous turbulence. The leading edge serrations are in the form of sinusoidal profiles of wavelengths, λ, and amplitudes, 2h. The frequency and amplitude characteristics are studied in detail in order to understand the effect of LE serrations on noise reduction characteristics and are compared with straight edge baseline flat plates. Noise reductions are found to be insignificant at low frequencies but significant in the mid frequency range (500 Hz-8 kHz) for all the cases studied. The flat plate results are also compared to the noise reductions obtained on a serrated NACA-65 aerofoil with the same serration profile. Noise reductions are found to be significantly higher for the flat plates with a maximum noise reduction of around 9 dB compared with about 7 dB for the aerofoil. In general, it is observed that the sound power reduction level (ΔPWL) is sensitive to the amplitude, 2h of the LE serrations but less sensitive to the serration wavelength, λ. Thus, this paper sufficiently demonstrates that the LE amplitude acts as a key parameter for enhancing the noise reduction levels in flat plates and aerofoils.

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

An experimental study of an airfoil with a bio-inspired leading edge device at high angles of attack

NASA Astrophysics Data System (ADS)

Mandadzhiev, Boris A.; Lynch, Michael K.; Chamorro, Leonardo P.; Wissa, Aimy A.

2017-09-01

Robust and predictable aerodynamic performance of unmanned aerial vehicles at the limits of their design envelope is critical for safety and mission adaptability. Deployable aerodynamic surfaces from the wing leading or trailing edges are often used to extend the aerodynamic envelope (e.g. slats and flaps). Birds have also evolved feathers at the leading edge (LE) of their wings, known as the alula, which enables them to perform high angles of attack maneuvers. In this study, a series of wind tunnel experiments are performed to quantify the effect of various deployment parameters of an alula-like LE device on the aerodynamic performance of a cambered airfoil (S1223) at stall and post stall conditions. The alula relative angle of attack, measured from the mean chord of the airfoil, is varied to modulate tip-vortex strength, while the alula deflection angle is varied to modulate the distance between the tip vortex and the wing surface. Integrated lift force measurements were collected at various alula-inspired device configurations. The effect of the alula-inspired device on the boundary layer velocity profile and turbulence intensity were investigated through hot-wire anemometer measurements. Results show that as alula deflection angle increases, the lift coefficient also increase especially at lower alula relative angles of attack. Moreover, at post stall wing angles of attack, the wake velocity deficit is reduced in the presence of alula device, confirming the mitigation of the wing adverse pressure gradient. The results are in strong agreement with measurements taken on bird wings showing delayed flow reversal and extended range of operational angles of attack. An engineered alula-inspired device has the potential to improve mission adaptability in small unmanned air vehicles during low Reynolds number flight.

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.

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.

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.

Leading-Edge Flow Sensing for Aerodynamic Parameter Estimation

NASA Astrophysics Data System (ADS)

Saini, Aditya

The identification of inflow air data quantities such as airspeed, angle of attack, and local lift coefficient on various sections of a wing or rotor blade provides the capability for load monitoring, aerodynamic diagnostics, and control on devices ranging from air vehicles to wind turbines. Real-time measurement of aerodynamic parameters during flight provides the ability to enhance aircraft operating capabilities while preventing dangerous stall situations. This thesis presents a novel Leading-Edge Flow Sensing (LEFS) algorithm for the determination of the air -data parameters using discrete surface pressures measured at a few ports in the vicinity of the leading edge of a wing or blade section. The approach approximates the leading-edge region of the airfoil as a parabola and uses pressure distribution from the exact potential-ow solution for the parabola to _t the pressures measured from the ports. Pressures sensed at five discrete locations near the leading edge of an airfoil are given as input to the algorithm to solve the model using a simple nonlinear regression. The algorithm directly computes the inflow velocity, the stagnation-point location, section angle of attack and lift coefficient. The performance of the algorithm is assessed using computational and experimental data in the literature for airfoils under different ow conditions. The results show good correlation between the actual and predicted aerodynamic quantities within the pre-stall regime, even for a rotating blade section. Sensing the deviation of the aerodynamic behavior from the linear regime requires additional information on the location of ow separation on the airfoil surface. Bio-inspired artificial hair sensors were explored as a part of the current research for stall detection. The response of such artificial micro-structures can identify critical ow characteristics, which relate directly to the stall behavior. The response of the microfences was recorded via an optical microscope for

Performance of hydrofoils with humpback whale-like leading edge protuberances.

NASA Astrophysics Data System (ADS)

Levshin, Alexandra; Henoch, Charles; Johari, Hamid

2005-11-01

The humpback whale (Megaptera novaeangliae) is extremely maneuverable, compared to other whale species, despite its large size and rigid body. Turning maneuvers are especially evident during pursuit of prey. The agility of humpback whale has been attributed to their use of pectoral flippers. The thick flippers have large aspect ratios, and large scale protuberances are present on the leading edge. The flippers do not flap during turning maneuvers. The cross-section of the flipper has a profile similar to a NACA 634-021 airfoil. The amplitude of leading edge protuberances ranges from 2.5 to 12% of the chord, with a spanwise extent of 10 to 50% the chord depending on the location along the span. It has been hypothesized that the `bumpy' leading edge is used for flow control. To examine the effects of protuberances on the leading edge of hydrofoils, a series of rectangular foils with bumpy leading edges were manufactured. The leading edge is sinusoidal in the spanwise direction with amplitudes and wavelengths comparable to that of humpback whale's flippers. The forces and moments on these bumpy foils were measured in a water tunnel and compared with a smooth leading edge foil.

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.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading edge thrust are possible at supersonic speeds for certain planforms which match the theoretical thrust distribution potential with the supporting airfoil geometry. The analytical process employed spanwise distribution of both it and/or that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

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.

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.

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.

The aerodynamic design of an advanced rotor airfoil

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

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.

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.

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.

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.

Unsteady aerodynamic behavior of an airfoil with and without a slat

NASA Technical Reports Server (NTRS)

Tung, Chee; Mcalister, Kenneth W.; Wang, Clin M.

1993-01-01

Unsteady flow behavior and load characteristics of a 2D VR-7 airfoil with and without a leading-edge slat were studied in the water tunnel of the Aeroflightdynamics Directorate, NASA Ames Research Center. Both airfoils were oscillated sinusoidally between 5 and 25 deg at Re = 200,000 to obtain the unsteady lift, drag, and pitching moment data. A fluorescent dye was released from an orifice located at the leading edge of the airfoil for the purpose of visualizing the boundary layer and wake flow. The flowfield and load predictions of an incompressible Navier-Stokes code based on a velocity-vorticity formulation were compared with the test data. The test and predictions both confirm that the slatted VR-7 airfoil delays both static and dynamic stall as compared to the VR-7 airfoil alone.

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.

Vortex scale of unsteady separation on a pitching airfoil.

PubMed

Fuchiwaki, Masaki; Tanaka, Kazuhiro

2002-10-01

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

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.

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.

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)

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.

Application of the Program Profile for the Design of Low-Speed, Low- Observable Configuration Airfoils

DTIC Science & Technology

1992-12-01

112 61 . Airfoil T503 - t/c = 3.79% .... ........... .. 113 62. Airfoil T503 Leading-Edge - t/c = 3.79% ..... ... 114 63. Pressure...points on C unit circle, 6 slope of airfoil surface near trailing edge 61 boundary-layer displacement thickness 62 boundary-layer momentum thickness 63...equivalent thickness NACA 4-digit airfoils . 4 II. Theory Potential-Flow Design Method This section will overview the basic theory used in PROFILE. Eppler

Investigation of a Novel Turbulence Model and Using Leading-Edge Slots for Improving the Aerodynamic Performance of Airfoils and Wind Turbines

NASA Astrophysics Data System (ADS)

Beyhaghi, Saman

as compared to the baseline DES. In the second part of this study, the focus is on improving the aerodynamic performance of airfoils and wind turbines in terms of lift and drag coefficients and power generation. One special type of add-on feature for wind turbines and airfoils, i.e., leading-edge slots are investigated through numerical simulation and laboratory experiments. Although similar slots are designed and employed for aircrafts, a special slot with a reversed flow direction is drilled in the leading edge of a sample wind turbine airfoil to study its influence on the aerodynamic performance. The objective is to vary the five main geometrical parameters of slot and characterize the performance improvement of the new design under different operating conditions. A number of Design of Experiment and optimization studies are conducted to determine the most suitable slot configuration to maximize the lift or lift-over-drag ratio. Results indicate that proper sizing and placement of slot can improve the lift coefficient, while it has negligible negative impact on the drag. Some recommendations for future investigation on slot are proposed at the end. The performance of a horizontal axis wind turbine blade equipped with leading-edge slot is also studied, and it is concluded that slotted blades can generate about 10% more power than solid blades, for the two operating conditions investigated. The good agreement between the CFD predictions and experimental data confirms the validity of the model and results.

Reduced-order aeroelastic model for limit-cycle oscillations in vortex-dominated unsteady airfoil flows

NASA Astrophysics Data System (ADS)

Suresh Babu, Arun Vishnu; Ramesh, Kiran; Gopalarathnam, Ashok

2017-11-01

In previous research, Ramesh et al. (JFM,2014) developed a low-order discrete vortex method for modeling unsteady airfoil flows with intermittent leading edge vortex (LEV) shedding using a leading edge suction parameter (LESP). LEV shedding is initiated using discrete vortices (DVs) whenever the Leading Edge Suction Parameter (LESP) exceeds a critical value. In subsequent research, the method was successfully employed by Ramesh et al. (JFS, 2015) to predict aeroelastic limit-cycle oscillations in airfoil flows dominated by intermittent LEV shedding. When applied to flows that require large number of time steps, the computational cost increases due to the increasing vortex count. In this research, we apply an amalgamation strategy to actively control the DV count, and thereby reduce simulation time. A pair each of LEVs and TEVs are amalgamated at every time step. The ideal pairs for amalgamation are identified based on the requirement that the flowfield in the vicinity of the airfoil is least affected (Spalart, 1988). Instead of placing the amalgamated vortex at the centroid, we place it at an optimal location to ensure that the leading-edge suction and the airfoil bound circulation are conserved. Results of the initial study are promising.

Unsteady flow past an airfoil pitched at constant rate

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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

NASA Technical Reports Server (NTRS)

1989-01-01

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

A comparative analysis between NACA 4412 airfoil and it's modified form with tubercles

NASA Astrophysics Data System (ADS)

Hasan, Md. Jonayed; Islam, Md. Tazul; Hassan, Md. Mehedi

2017-06-01

The effect of tubercles on the leading edge of an airfoil become more vivid at high angle of attacks. The effect of tubercles with large wavelength and small amplitude on the leading edge of a NACA 4412 airfoil section was investigated numerically and experimentally. The phenomena of improving the airfoil performance by modifying the contours drove our interest to do this analysis. The models were developed & numerical simulations were carried out with both NACA 4412 airfoil and modified airfoil model at Re=1.03×106 and angles of attack ranging from 0° to 20°. Flow separation was analyzed with vector profiles. CL, CD at different angle of attacks was developed and it gave down noticeable pre-stall & post-stall behavior. The airfoils were studied experimentally in a low speed wind tunnel. Pressure distribution over the two airfoils was obtained. It was evident from the pressure distributions that the modified airfoil exhibits significant aerodynamic performance at high angles of attack. We can infer that these effects will be advantageous for maneuverability and post-stall behavior.

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.

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

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.

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.

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.

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.

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.

Optimization of multi-element airfoils for maximum lift

NASA Technical Reports Server (NTRS)

Olsen, L. E.

1979-01-01

Two theoretical methods are presented for optimizing multi-element airfoils to obtain maximum lift. The analyses assume that the shapes of the various high lift elements are fixed. The objective of the design procedures is then to determine the optimum location and/or deflection of the leading and trailing edge devices. The first analysis determines the optimum horizontal and vertical location and the deflection of a leading edge slat. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This design procedure does not require that flow separation effects be modeled. The second analysis determines the slat and flap deflection required to maximize the lift of a three element airfoil. This approach requires that the effects of flow separation from one or more of the airfoil elements be taken into account. The theoretical results are in good agreement with results of a wind tunnel test used to corroborate the predicted optimum slat and flap positions.

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.

Near-wall serpentine cooled turbine airfoil

DOEpatents

Lee, Ching-Pang

2013-09-17

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

Close-loop Dynamic Stall Control on a Pitching Airfoil

NASA Astrophysics Data System (ADS)

Giles, Ian; Corke, Thomas

2017-11-01

A closed-loop control scheme utilizing a plasma actuator to control dynamic stall is presented. The plasma actuator is located at the leading-edge of a pitching airfoil. It initially pulses at an unsteady frequency that perturbs the boundary layer flow over the suction surface of the airfoil. As the airfoil approaches and enters stall, the amplification of the unsteady disturbance is detected by an onboard pressure sensor also located near the leading edge. Once detected, the actuator is switched to a higher voltage control state that in static airfoil experiments would reattach the flow. The threshold level of the detection is a parameter in the control scheme. Three stall regimes were examined: light, medium, and deep stall, that were defined by their stall penetration angles. The results showed that in general, the closed-loop control scheme was effective at controlling dynamic stall. The cycle-integrated lift improved in all cases, and increased by as much as 15% at the lowest stall penetration angle. As important, the cycle-integrated aerodynamic damping coefficient also increased in all cases, and was made to be positive at the light stall regime where it traditionally is negative. The latter is important in applications where negative damping can lead to stall flutter.

Turbine airfoil to shroud attachment method

DOE Office of Scientific and Technical Information (OSTI.GOV)

Campbell, Christian X; Kulkarni, Anand A; James, Allister W

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) ofmore » 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.« less

A Factor Affecting Transonic Leading-edge Flow Separation

NASA Technical Reports Server (NTRS)

Wood, George P; Gooderum, Paul B

1956-01-01

A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock-wave - boundary layer interaction are proposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. Some experimental evidence in support of the importance of this factor is presented.

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

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.

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.

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.

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

This paper investigates the feasibility of locating a flush air-data sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered. 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 air-data calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush air-data systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

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.

Dynamic stall study of a multi-element airfoil

NASA Technical Reports Server (NTRS)

Tung, Chee; Mcalister, Kenneth W.; Wang, Clin M.

1992-01-01

Unsteady flow behavior and load characteristics of a VR-7 airfoil with and without a slat were studied in the water tunnel of the Aeroflightdynamics Directorate, NASA Ames Research Center. Both airfoils were oscillated sinusoidally between 5 and 25 degrees at a Reynolds number of 200,000 to obtain the unsteady lift, drag and pitching moment data. A fluorescing dye was released from an orifice located at the leading edge of the airfoil for the purpose of visualizing the boundary layer and wake flow. The flow field and load predictions of an incompressible Navier-Stokes code based on a velocity-vorticity formulation were compared with the test data. The test and predictions both confirm that the slatted VR-7 airfoil delays both static and dynamic stall as compared to the VR-7 airfoil alone.

Wall Modeled Large Eddy Simulation of Airfoil Trailing Edge Noise

NASA Astrophysics Data System (ADS)

Kocheemoolayil, Joseph; Lele, Sanjiva

2014-11-01

Large eddy simulation (LES) of airfoil trailing edge noise has largely been restricted to low Reynolds numbers due to prohibitive computational cost. Wall modeled LES (WMLES) is a computationally cheaper alternative that makes full-scale Reynolds numbers relevant to large wind turbines accessible. A systematic investigation of trailing edge noise prediction using WMLES is conducted. Detailed comparisons are made with experimental data. The stress boundary condition from a wall model does not constrain the fluctuating velocity to vanish at the wall. This limitation has profound implications for trailing edge noise prediction. The simulation over-predicts the intensity of fluctuating wall pressure and far-field noise. An improved wall model formulation that minimizes the over-prediction of fluctuating wall pressure is proposed and carefully validated. The flow configurations chosen for the study are from the workshop on benchmark problems for airframe noise computations. The large eddy simulation database is used to examine the adequacy of scaling laws that quantify the dependence of trailing edge noise on Mach number, Reynolds number and angle of attack. Simplifying assumptions invoked in engineering approaches towards predicting trailing edge noise are critically evaluated. We gratefully acknowledge financial support from GE Global Research and thank Cascade Technologies Inc. for providing access to their massively-parallel large eddy simulation framework.

Pressure distributions from subsonic tests of an advanced laminar-flow-control wing with leading- and trailing-edge flaps

NASA Technical Reports Server (NTRS)

Applin, Zachary T.; Gentry, Garl L., Jr.

1988-01-01

An unswept, semispan wing model equipped with full-span leading- and trailing-edge flaps was tested in the Langley 14- by 22-Foot Subsonic Tunnel to determine the effect of high-lift components on the aerodynamics of an advanced laminar-flow-control (LFC) airfoil section. Chordwise pressure distributions near the midsemispan were measured for four configurations: cruise, trailing-edge flap only, and trailing-edge flap with a leading-edge Krueger flap of either 0.10 or 0.12 chord. Part 1 of this report (under separate cover) presents a representative sample of the plotted pressure distribution data for each configuration tested. Part 2 presents the entire set of plotted and tabulated pressure distribution data. The data are presented without analysis.

Unsteady loading on an airfoil of arbitrary thickness

NASA Astrophysics Data System (ADS)

Glegg, Stewart A. L.; Devenport, William

2009-01-01

The unsteady loading on an airfoil of arbitrary thickness is evaluated by using the generalized form of Blasius theorem and a conformal mapping that maps the airfoil surface onto a circle. For a blade vortex interaction the results show that the time history of the unsteady loading is determined by the passage of the vortex relative to the leading edge singularity in the circle plane. The singularity lies inside the circle and moves to a smaller radius as the thickness is increased, causing the unsteady loading pulse to be smoothed. The effect of angle of attack is to move the stagnation point relative to the leading edge singularity and this significantly increases the unsteady lift if the vortex passes on the suction side of the airfoil. These characteristics are different for a step upwash gust, which is considered as a simplified model of a large scale turbulent gust. It is shown that the time history of the magnitude of the unsteady loading is almost completely unaltered by angle of attack for the step gust, but it's direction of action rotates forward by an angle equal to the angle of attack, extending an earlier result by Howe for a flat plate in a turbulent flow to airfoils of arbitrary thickness. However spectral analysis of the gust shows that the high frequency blade response is reduced as the thickness of the airfoil is increased.

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.

Computational analysis of blunt, thin airfoil sections at supersonic and subsonic speeds

NASA Astrophysics Data System (ADS)

Goodsell, Aga Myung

The past decade has brought renewed interest in commercial supersonic aircraft design. Recent wing designs have included regions of low sweep resulting in supersonic leading edges at cruise. Thin biconvex sections are used in those regions to minimize wave drag and skin-friction drag. However, airfoil sections with sharp leading edges exhibit poor aerodynamic behavior at subsonic flight conditions. Blunt leading edges may improve performance by delaying the onset of separation at subsonic and transonic speeds. Their disadvantage is that they increase both wave drag, due to the formation of a detached bow wave, and skin-friction drag, from a loss of laminar flow. The effect of adding bluntness to a 4%-thick biconvex section was investigated using computational analysis tools. The aerodynamic performance of biconvex sections with circular leading edges was computed at supersonic, transonic, and takeoff conditions. At supersonic cruise, the increase in wave drag due to bluntness is a function of Mach number and leading-edge diameter. Some of the drag penalty is offset by the suction created downstream of the circular leading edge. The possibility of further drag reduction was explored with the development of a semi-analytical method to design blunt airfoil shapes which minimize wave drag. The effect on the transition location was evaluated using linear stability analyses of laminar boundary-layer profiles and the eN method. The analysis showed that laminar boundary layers on blunt airfoil sections are considerably less stable to Tollmien-Schlichting waves than that on a sharp biconvex. At transonic speeds, the results suggest a possible improvement in the lift-to-drag ratio over a limited range of angles of attack. At the takeoff condition, slight blunting of the leading edge does improve the lift-to-drag ratio at low angles of attack, but has little effect on maximum lift. It is concluded that the benefit of a blunt leading edge at off-design conditions is not

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.

On the acoustic signature of tandem airfoils: The sound of an elastic airfoil in the wake of a vortex generator

DOE Office of Scientific and Technical Information (OSTI.GOV)

Manela, A.

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 calculationsmore » 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.« less

Airfoil System for Cruising Flight

NASA Technical Reports Server (NTRS)

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

2014-01-01

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

Development of an experimental setup for analyzing the influence of Magnus effect on the performance of airfoil

NASA Astrophysics Data System (ADS)

Aktharuzzaman, Md; Sarker, Md. Samad; Safa, Wasiul; Sharah, Nahreen; Salam, Md. Abdus

2017-12-01

Magnus effect is a phenomenon where pressure difference is created according to Bernoulli's effect due to induced velocity changes caused by a rotating object in a fluid. Using this concept, the idea of delaying boundary layer separation on airfoil by providing moving surface boundary layer control has been developed. In order to analyze the influence of Magnus effect on the aerodynamic performance of an airfoil, there is no alternative of developing an experimental setup. This paper aims to develop such an experimental setup which will be capable of analyzing the influence of Magnus effect on both symmetric and asymmetric airfoils by placing a cylinder at the leading edge. To provide arrangements for a rotating cylinder at the leading edge of airfoil, necessary modifications and additions have been done in the test section of an AF100 subsonic wind tunnel.

Integral Textile Structure for 3-D CMC Turbine Airfoils

NASA Technical Reports Server (NTRS)

Marshall, David B. (Inventor); Cox, Brian N. (Inventor); Sudre, Olivier H. (Inventor)

2017-01-01

An integral textile structure for 3-D CMC turbine airfoils includes top and bottom walls made from an angle-interlock weave, each of the walls comprising warp and weft fiber tows. The top and bottom walls are merged on a first side parallel to the warp fiber tows into a single wall along a portion of their widths, with the weft fiber tows making up the single wall interlocked through the wall's thickness such that delamination of the wall is inhibited. The single wall suitably forms the trailing edge of an airfoil; the top and bottom walls are preferably joined along a second side opposite the first side and parallel to the radial fiber tows by a continuously curved section in which the weave structure remains continuous with the weave structure in the top and bottom walls, the continuously curved section being the leading edge of the airfoil.

Airfoil for a gas turbine engine

DOEpatents

Liang, George [Palm City, FL

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.

Pressure Distribution Over a Symmetrical Airfoil Section with Trailing Edge Flap

NASA Technical Reports Server (NTRS)

Jacobs, Eastman N; Pinkerton, Robert M

1931-01-01

Measurements were made to determine the distribution of pressure over one section of an R. A. F. 30 (symmetrical) airfoil with trailing edge flaps. In order to study the effect of scale measurements were made with air densities of approximately 1 and 20 atmospheres. Isometric diagrams of pressure distribution are given to show the effect of change in incidence, flap displacement, and scale upon the distribution. Plots of normal force coefficient versus angle of attack for different flap displacements are given to show the effect of a displaced flap. Plots are given of both the experimental and theoretical characteristic coefficients versus flap angle, in order to provide a comparison with the theory. It is concluded that for small flap displacements the agreement for the pitching and hinge moments is such that it warrants the use of the theoretical parameters. However, the agreement for the lift is not as good, particularly for the smaller flaps. In an appendix, an example is given of the calculation of the load and moments on an airfoil with hinged flap from these parameters.

Low-speed aerodynamic characteristics of a 13.1-percent-thick, high-lift airfoil

NASA Technical Reports Server (NTRS)

Sivier, K. R.; Ormsbee, A. I.; Awker, R. W.

1974-01-01

Experimental study of the low-speed, sectional characteristics of a high-lift airfoil, and comparison of these characteristics with the predictions of the theoretical methods used in the airfoil's design. The 13.1% thick UI-1720 airfoil was found to achieve the predicted maximum lift coefficient of nearly 2.0. No upper-surface flow separation was found below the stall angle of attack of 16 deg; it appeared that stall was due to an abrupt leading-edge flow separation.

Reynolds number effect on airfoil wake structures under pitching and heaving motion

NASA Astrophysics Data System (ADS)

Kim, Kyung Chun; Karbasian, Hamidreza; ExpTENsys Team

2017-11-01

Detached Eddy Simulation (DES) and particle image velocimetry (PIV) measurements were performed to investigate the wake flow characteristics of an airfoil under pitching and heaving motion. A NACA0012 airfoil was selected for the numerical simulation and experiments were carried out in a wind tunnel and a water tunnel at Reynolds number of 15,000 and 90,000, respectively. The airfoil oscillated around an axis located 1/4 distance from the leading edge chord. Two different angles of attack, 20° and 30°, were selected with +/-10° maximum amplitude of oscillation. In order to extract the coherent flow structures from time-resolved PIV data, proper orthogonal decomposition (POD) analysis was performed on 1,000 instantaneous realisations for each condition using the method of snapshots. Vorticity contour and velocity profiles for both PIV and DES results are in good agreement for pitching and heaving motion. At high Reynolds number, 3D stream-wise vortices appeared after generating span-wise vortices. The higher maximum angle of attack allows the leading edge vortex to grow stronger and that the angle of attack appears to be more important in influencing the growth of the leading edge vortex structure than the reduced frequency. National Research Foundation of Korea (No. 2011-0030013).

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.

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.

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

Vortex developments over steady and accelerated airfoils incorporating a trailing edge jet

NASA Technical Reports Server (NTRS)

Finaish, F.; Okong'o, N.; Frigerio, J.

1993-01-01

Computational and experimental studies are conducted to investigate the influence of a trailing edge jet on flow separation and subsequent vortex formation over steady and accelerated airfoils at high angles of attack. A computer code, employing the stream function-vorticity approach, is developed and utilized to conduct numerical experiments on the flow problem. To verify and economize such efforts, an experimental system is developed and incorporated into a subsonic wind tunnel where streamline and vortex flow visualization experiments are conducted. The study demonstrates the role of the trailing edge jet in controlling flow separation and subsequent vortex development for steady and accelerating flow at angles past the static stall angle of attack. The results suggest that the concept of the trailing edge jet may be utilized to control the characteristics of unsteady separated flows over lifting surfaces. This control possibility seems to be quite effective and could have a significant role in controlling unsteady separated flows.

Direct Numerical Simulation of Flows over an NACA-0012 Airfoil at Low and Moderate Reynolds Numbers

NASA Technical Reports Server (NTRS)

Balakumar, P.

2017-01-01

Direct numerical simulations (DNS) of flow over an NACA-0012 airfoil are performed at a low and a moderate Reynolds numbers of Re(sub c)=50 times10(exp 3) and 1times 10(exp 6). The angles of attack are 5 and 15 degrees at the low and the moderate Reynolds number cases respectively. The three-dimensional unsteady compressible Navier-Stokes equations are solved using higher order compact schemes. The flow field in the low Reynolds number case consists of a long separation bubble near the leading-edge region and an attached boundary layer on the aft part of the airfoil. The shear layer that formed in the separated region persisted up to the end of the airfoil. The roles of the turbulent diffusion, advection, and dissipation terms in the turbulent kinetic-energy balance equation change as the boundary layer evolves over the airfoil. In the higher Reynolds number case, the leading-edge separation bubble is very small in length and in height. A fully developed turbulent boundary layer is observed in a short distance downstream of the reattachment point. The boundary layer velocity near the wall gradually decreases along the airfoil. Eventually, the boundary layer separates near the trailing edge. The Reynolds stresses peak in the outer part of the boundary layer and the maximum amplitude also gradually increases along the chord.

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.

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.

Active Control of Flow Separation Over an Airfoil

NASA Technical Reports Server (NTRS)

Ravindran, S. S.

1999-01-01

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

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.

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.

Computation of oscillating airfoil flows with one- and two-equation turbulence models

NASA Technical Reports Server (NTRS)

Ekaterinaris, J. A.; Menter, F. R.

1994-01-01

The ability of one- and two-equation turbulence models to predict unsteady separated flows over airfoils is evaluated. An implicit, factorized, upwind-biased numerical scheme is used for the integration of the compressible, Reynolds-averaged Navier-Stokes equations. The turbulent eddy viscosity is obtained from the computed mean flowfield by integration of the turbulent field equations. One- and two-equation turbulence models are first tested for a separated airfoil flow at fixed angle of incidence. The same models are then applied to compute the unsteady flowfields about airfoils undergoing oscillatory motion at low subsonic Mach numbers. Experimental cases where the flow has been tripped at the leading-edge and where natural transition was allowed to occur naturally are considered. The more recently developed turbulence models capture the physics of unsteady separated flow significantly better than the standard kappa-epsilon and kappa-omega models. However, certain differences in the hysteresis effects are observed. For an untripped high-Reynolds-number flow, it was found necessary to take into account the leading-edge transitional flow region to capture the correct physical mechanism that leads to dynamic stall.

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.

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

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.

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.

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.

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.

PIV Study on Flow around Leading-Edge Slat of 30P30N Airfoil

NASA Astrophysics Data System (ADS)

Ando, Ryosuke; Onishi, Yusaku; Sakakibara, Jun

2017-11-01

We measured flow velocity distribution around leading-edge slat using PIV. Simultaneously, noise measurement using microphone was also performed. A leading-edge slat and main wing model having a chord length of 160 mm was placed in the tunnel with free stream velocity of about 26m/s and chord Reynolds number of 2.8 x 105. Angle of attack was changed from 4 degrees to 10 degrees at two degree intervals. In this experiment, we investigated the relationship between the unsteady flow condition and the noise. At 4 degrees in the angle of attack, vortices shedding from the slat cusp were moved to the downstream. At 6 degrees or more, flow velocity distributions show that vortices were reattached on the slat lower surface and the flow in the slat cove recirculated. In FFT analysis of noise measurement, at 6 degrees in the angle of attack, there were some peaks on low frequency area and dominant peak on high frequency area was found. At 8 degrees or more, there were also some peaks on low frequency area. But dominant peak on high frequency area disappeared.

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.

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.

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.

Schlieren visualization of flow-field modification over an airfoil by near-surface gas-density perturbations generated by a nanosecond-pulse-driven plasma actuator

NASA Astrophysics Data System (ADS)

Komuro, Atsushi; Takashima, Keisuke; Konno, Kaiki; Tanaka, Naoki; Nonomura, Taku; Kaneko, Toshiro; Ando, Akira; Asai, Keisuke

2017-06-01

Gas-density perturbations near an airfoil surface generated by a nanosecond dielectric-barrier-discharge plasma actuator (ns-DBDPA) are visualized using a high-speed Schlieren imaging method. Wind-tunnel experiments are conducted for a wind speed of 20 m s-1 with an NACA0015 airfoil whose chord length is 100 mm. The results show that the ns-DBDPA first generates a pressure wave and then stochastic perturbations of the gas density near the leading edge of the airfoil. Two structures with different characteristics are observed in the stochastic perturbations. One structure propagates along the boundary between the shear layer and the main flow at a speed close to that of the main flow. The other propagates more slowly on the surface of the airfoil and causes mixing between the main and shear flows. It is observed that these two heated structures interact with each other, resulting in a recovery in the negative pressure coefficient at the leading edge of the airfoil.

Effect of leading-edge roughness on stability and transition of wind turbine blades

NASA Astrophysics Data System (ADS)

Kutz, Douglas; Freels, Justin; Hidore, John; White, Edward

2011-11-01

Over time, wind turbine blades erode due to impacts with sand and other debris. The resulting surface roughness degrades the blades' aerodynamic performance. Experimental studies conducted at the Texas A&M University Low-Speed Wind Tunnel examine roughness effects using a 2D NACA 63-418 airfoil with interchangeable leading edges of varying roughness at chord Reynolds numbers up to 3 . 0 ×106 . These data reveal decreased CL , max and increased CD , min as roughness increases. At very high roughness levels, even the lift curve slope is reduced. To better understand these findings and improve modeling of roughness effects, extensive boundary layer measurements including surface-mounted hotfilms and boundary-layer velocity profiles are used to assess how laminar-to-turbulent transition is promoted by roughness. As expected, roughness accelerates transition. Tollmien-Schlichting (TS) transition is observed only for a smooth leading edge while bypass transition is observed for the moderate and high roughness levels. Results are compared to N-factor transition predictions generated with software used by the wind industry. Predictions are successful for the smooth leading edge but even the low roughness level prevents correct transition prediction using TS-based methods. Support for this work by Vestas Technology Americas, Inc., is gratefully acknowledged as is the support of the wind-energy research group and the Low-Speed Wind Tunnel staff.

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.

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

NASA Technical Reports Server (NTRS)

Whitcomb, R. T. (Inventor)

1976-01-01

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

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.

Cooled airfoil in a turbine engine

DOEpatents

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.

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.

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.

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

NASA Technical Reports Server (NTRS)

Nirmalan, V.; Hylton, L. D.

1989-01-01

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

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.

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.

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.

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

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

Theoretical and experimental study of a new method for prediction of profile drag of airfoil sections

NASA Technical Reports Server (NTRS)

Goradia, S. H.; Lilley, D. E.

1975-01-01

Theoretical and experimental studies are described which were conducted for the purpose of developing a new generalized method for the prediction of profile drag of single component airfoil sections with sharp trailing edges. This method aims at solution for the flow in the wake from the airfoil trailing edge to the large distance in the downstream direction; the profile drag of the given airfoil section can then easily be obtained from the momentum balance once the shape of velocity profile at a large distance from the airfoil trailing edge has been computed. Computer program subroutines have been developed for the computation of the profile drag and flow in the airfoil wake on CDC6600 computer. The required inputs to the computer program consist of free stream conditions and the characteristics of the boundary layers at the airfoil trailing edge or at the point of incipient separation in the neighborhood of airfoil trailing edge. The method described is quite generalized and hence can be extended to the solution of the profile drag for multi-component airfoil sections.

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.

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

NASA Technical Reports Server (NTRS)

Emery, James C

1956-01-01

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

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

NASA Technical Reports Server (NTRS)

Lattanzi, Bernardino; Bellante, Erno

1949-01-01

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

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.

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

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.

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.

An improved viscid/inviscid interaction procedure for transonic flow over airfoils

NASA Technical Reports Server (NTRS)

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

1985-01-01

A new interacting boundary layer approach for computing the viscous transonic flow over airfoils is described. The theory includes a complete treatment of viscous interaction effects induced by the wake and accounts for normal pressure gradient effects across the boundary layer near trailing edges. The method is based on systematic expansions of the full Reynolds equation of turbulent flow in the limit of Reynolds numbers, Reynolds infinity. Procedures are developed for incorporating the local trailing edge solution into the numerical solution of the coupled full potential and integral boundary layer equations. Although the theory is strictly applicable to airfoils with cusped or nearly cusped trailing edges and to turbulent boundary layers that remain fully attached to the airfoil surface, the method was successfully applied to more general airfoils and to flows with small separation zones. Comparisons of theoretical solutions with wind tunnel data indicate the present method can accurately predict the section characteristics of airfoils including the absolute levels of drag.

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.

The Effectiveness at High Speeds of a 20-Percent-Chord Plain Trailing-Edge Flap on the NACA 65-210 Airfoil

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 flap on the NACA 65-210 airfoil section. The analysis indicates that the plain trailing-edge flap employed on the 10-percent-thick airfoil at Mach numbers as high as 0.875 retains at least 50-percent of its low-speed lift-control effectiveness, and is sufficiently effective in lateral control application, assuming a rigid wing, to provide adequate airplane rolling characteristics. The plain trailing-edge flap, as compared to the spoiler and the dive-recovery flap, appears to afford the most favorable characteristics as a device for controlling lift continuously throughout the range of Mach numbers from 0.3 to 0.875. At Mach numbers above those for lift divergence of the wing, either a plain flap or a dive-recovery flap may be used on a thin airplane wing to provide auxiliary wing lift when the airplane is to be controlled in flight, other than in dives, at these Mach numbers. The choice of a lift-control device for this use, however, should include the consideration of other factors such as the increments of drag and pitching moment accompanying the use of the device, and the structural and high-speed aerodynamic characteristics of the airplane which is to employ the device.

Lift-Enhancing Tabs on Multielement Airfoils

NASA Technical Reports Server (NTRS)

Ross, James C.; Storms, Bruce L.; Carrannanto, Paul G.

1995-01-01

The use of flat-plate tabs (similar to Gurney flaps) to enhance the lift of multielement airfoils is extended here by placing them on the pressure side and near the trailing edge of the main element rather than just on the furthest downstream wing element. The tabs studied range in height from 0.125 to 1.25% of the airfoil reference chord. In practice, such tabs would be retracted when the high-lift system is stowed. The effectiveness of the concept was demonstrated experimentally and computationally on a two-dimensional NACA 63(sub 2)-215 Mod B airfoil with a single-slotted, 30%-chord flap. Both the experiments and computations showed that the tabs significantly increase the lift at a given angle of attack and the maximum lift coefficient of the airfoil. The computational results showed that the increased lift was a result of additional turning of the flow by the tab that reduced or eliminated now separation on the flap. The best configuration tested, a 0.5%-chord tab placed 0.5% chord upstream of the trailing edge of the main element, increased the maximum lift coefficient of the airfoil by 12% and the maximum lift-to-drag ratio by 40%.

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.

Design and Experimental Results for the S414 Airfoil

DTIC Science & Technology

2010-08-01

EXECUTION The Eppler Airfoil Design and Analysis Code (refs. 15 and 16), a subcritical, single- element code, was used to design the initial fore- and...1965. 14. Maughmer, Mark D.: Trailing Edge Conditions as a Factor in Airfoil Design. Ph.D. Dis- sertation, Univ. of Illinois, 1983.14 15. Eppler ...Richard: Airfoil Design and Data. Springer-Verlag (Berlin), 1990. 16. Eppler , Richard: Airfoil Program System “PROFIL07.” User’s Guide. Richard

Enhancement of aerodynamic performance of a heaving airfoil using synthetic-jet based active flow control.

PubMed

Wang, Chenglei; Tang, Hui

2018-05-25

In this study, we explore the use of synthetic jet (SJ) in manipulating the vortices around a rigid heaving airfoil, so as to enhance its aerodynamic performance. The airfoil heaves at two fixed pitching angles, with the Strouhal number, reduced frequency and Reynolds number chosen as St  =  0.3, k  =  0.25 and Re  =  100, respectively, all falling in the ranges for natural flyers. As such, the vortex force plays a dominant role in determining the airfoil's aerodynamic performance. A pair of in-phase SJs is implemented on the airfoil's upper and lower surfaces, operating with the same strength but in opposite directions. Such a fluid-structure interaction problem is numerically solved using a lattice Boltzmann method based numerical framework. It is found that, as the airfoil heaves with zero pitching angle, its lift and drag can be improved concurrently when the SJ phase angle [Formula: see text] relative to the heave motion varies between [Formula: see text] and [Formula: see text]. But this concurrent improvement does not occur as the airfoil heaves with [Formula: see text] pitching angle. Detailed inspection of the vortex evolution and fluid stress over the airfoil surface reveals that, if at good timing, the suction and blowing strokes of the SJ pair can effectively delay or promote the shedding of leading edge vortices, and mitigate or even eliminate the generation of trailing edge vortices, so as to enhance the airfoil's aerodynamic performance. Based on these understandings, an intermittent operation of the SJ pair is then proposed to realize concurrent lift and drag improvement for the heaving airfoil with [Formula: see text] pitching angle.

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.

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.

Unsteady Navier-Stokes computations over airfoils using both fixed and dynamic meshes

NASA Technical Reports Server (NTRS)

Rumsey, Christopher L.; Anderson, W. Kyle

1989-01-01

A finite volume implicit approximate factorization method which solves the thin layer Navier-Stokes equations was used to predict unsteady turbulent flow airfoil behavior. At a constant angle of attack of 16 deg, the NACA 0012 airfoil exhibits an unsteady periodic flow field with the lift coefficient oscillating between 0.89 and 1.60. The Strouhal number is 0.028. Results are similar at 18 deg, with a Strouhal number of 0.033. A leading edge vortex is shed periodically near maximum lift. Dynamic mesh solutions for unstalled airfoil flows show general agreement with experimental pressure coefficients. However, moment coefficients and the maximum lift value are underpredicted. The deep stall case shows some agreement with experiment for increasing angle of attack, but is only qualitatively comparable past stall and for decreasing angle of attack.

Performance of flapping airfoil propulsion with LBM method and DMD analysis

NASA Astrophysics Data System (ADS)

Li, Bing-Hua; Huang, Xian-Wen; Zheng, Yao; Xie, Fang-Fang; Wang, Jing; Zou, Jian-Feng

2018-05-01

In this work, the performance of flapping airfoil propulsion at low Reynolds number of Re = 100-400 is studied numerically with the lattice Boltzmann method (LBM). Combined with immersed boundary method (IBM), the LBM has been widely used to simulate moving boundary problems. The influences of the reduced frequency on the plunging and pitching airfoil are explored. It is found that the leading-edge vertex separation and inverted wake structures are two main coherent structures, which dominate the flapping airfoil propulsion. However, the two structures play different roles in the flow and the combination effects on the propulsion need to be clarified. To do so, we adopt the dynamic mode decomposition (DMD) algorithm to reveal the underlying physics. The DMD has been proven to be very suitable for analyzing the complex transient systems like the vortex structure of flapping flight.

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

PubMed

Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

2015-01-01

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

Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics

PubMed Central

Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

2015-01-01

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

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.

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.

Numerical Investigation of an Oscillating Flat Plate Airfoil

NASA Astrophysics Data System (ADS)

Mohaghegh, Fazlolah; Janechek, Matthew; Buchholz, James; Udaykumar, Hs

2017-11-01

This research investigates the vortex dynamics of a plunging flat plate airfoil by analyzing the vorticity transport in 2D simulations. A horizontal airfoil is subject to a freestream flow at Re =10000. A prescribed vertical sinusoidal motion is applied to the airfoil. Smoothed Profile Method (SPM) models the fluid-structure interaction. SPM as a diffuse interface model considers a thickness for the interface and applies a smooth transition from solid to fluid. As the forces on the airfoil are highly affected by the interaction of the generated vortices from the surface, it is very important to find out whether a diffuse interface solver can model a flow dominated by vorticities. The results show that variation of lift coefficient with time agrees well with the experiment. Study of vortex evolution shows that similar to experiments, when the plate starts moving downward from top, the boundary layer is attached to the surface and the leading-edge vortex (LEV) is very small. By time, LEV grows and rolls up and a secondary vortex emerges. Meanwhile, the boundary layer starts to separate and finally LEV detaches from the surface. In overall, SPM as a diffuse interface model can predict the lift force and vortex pattern accurately.

Consideration of Unsteady Aerodynamics and Boundary-Layer Transition in Rotorcraft Airfoil Design

NASA Astrophysics Data System (ADS)

Oliveira Vieira, Bernardo Augusto de

Traditional rotorcraft airfoil design is based on steady-flow aerodynamic requirements. The approach assumes a strong correlation between steady and unsteady aerodynamic characteristics, which is often not observed in practice. This is particularly relevant at high speed and high thrust conditions, when the rotor is susceptible to dynamic stall and its many negative consequences. Given the abrupt nature of the phenomena, large margins are typically established to prevent fatigue loads on the blades and pitch links; thus, limiting operation under high altitudes, high payloads, high temperatures, as well as during maneuvers. This work addresses the problem from the perspective of passive airfoil design. Typical design requirements are revisited to include metrics for improved dynamic stall and new ways to qualifying rotorcraft airfoils are proposed. A number of design studies are conducted to better understand the relation between airfoil shape and dynamic stall behavior. The design manipulations are handled by an inversedesign, conformal mapping method, and unsteady Reynolds-averaged Navier-Stokes equations are used to predict the aerodynamic performance under pitch motion. In unsteady flow, the occurrence of aerodynamic lags in the development of pressures, boundary-layer separation, and viscous-inviscid interactions suggest more strict requirements than in steady flow. In order to postpone the onset of dynamic stall, the design needs to handle competing leading- and trailing-edge separation mechanisms, which are heavily influenced by local supersonic flow, strong shock waves, and laminar-turbulent transition effects. It is found that a particular tailoring of the trailing-edge separation development can provide adequate dynamic stall characteristics and minimize penalties in drag and nose-down pitching moment. At the same time, a proper design of the nose shape is required to avoid strong shock waves and prevent premature leading-edge stall. A proof

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

NASA Technical Reports Server (NTRS)

Allison, Dennis O.; Sewall, William G.

1995-01-01

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

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.

Computational Investigations of a NACA 0012 Airfoil in Low Reynolds Number Flows

DTIC Science & Technology

1992-09-01

11 D . RESULTS .................................... 13 1. Eppler E585 Airfoil ............................. 13 2. NACA 0012 Airfoil ...function in FORTRAN should also be used to calculate/3. D. RESULTS 1. Eppler E585 Airfoil The first investigation was conducted for an Eppler E585...The velocities match the given distribution well except for slight deviations at the trailing edge. This Figure 2.3 Eppler E585 Airfoil difference can

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.

Blade-mounted trailing edge flap control for BVI noise reduction

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

Effect of an extendable slat on the stall behavior of a VR-12 airfoil

NASA Technical Reports Server (NTRS)

Dehugues, P. Plantin; Mcalister, K. W.; Tung, C.

1993-01-01

Experimental and computational tests were performed on a VR-12 airfoil to determine if the dynamic-stall behavior that normally accompanies high-angle pitch oscillations could be modified by segmenting the forward portion of the airfoil and extending it ahead of the main element. In the extended position the configuration would appear as an airfoil with a leading-edge slat, and in the retracted position it would appear as a conventional VR-12 airfoil. The calculations were obtained from a numerical code that models the vorticity transport equation for an incompressible fluid. These results were compared with test data from the water tunnel facility of the Aeroflightdynamics Directorate at Ames Research Center. Steady and unsteady flows around both airfoils were examined at angles of attack between 0 and 30 deg. The Reynolds number was fixed at 200,000 and the unsteady pitch oscillations followed a sinusoidal motion described by alpha = alpha(sub m) + 10 deg sin(omega t). The mean angle (alpha(sub m)) was varied from 10 to 20 deg and the reduced frequency from 0.05 to 0.20. The results from the experiment and the calculations show that the extended-slat VR-12 airfoil experiences a delay in both static and dynamic stall not experienced by the basic VR-12 airfoil.

Characteristics of surface roughness associated with leading edge ice accretion

NASA Technical Reports Server (NTRS)

Shin, Jaiwon

1994-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

Leading-edge vortex lifts swifts.

PubMed

Videler, J J; Stamhuis, E J; Povel, G D E

2004-12-10

The current understanding of how birds fly must be revised, because birds use their hand-wings in an unconventional way to generate lift and drag. Physical models of a common swift wing in gliding posture with a 60 degrees sweep of the sharp hand-wing leading edge were tested in a water tunnel. Interactions with the flow were measured quantitatively with digital particle image velocimetry at Reynolds numbers realistic for the gliding flight of a swift between 3750 and 37,500. The results show that gliding swifts can generate stable leading-edge vortices at small (5 degrees to 10 degrees) angles of attack. We suggest that the flow around the arm-wings of most birds can remain conventionally attached, whereas the swept-back hand-wings generate lift with leading-edge vortices.

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.

Computer programs for smoothing and scaling airfoil coordinates

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.

1983-01-01

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

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 1: Analysis and design of stages A, B, and C

NASA Technical Reports Server (NTRS)

Brent, J. A.; Cheatham, J. G.; Nilsen, A. W.

1972-01-01

A conventional rotor and stator, two dual-airfoil tandem rotors, and one dual-airfoil tandem stator were designed. The two tandem rotors were each designed with different percentages of the overall lift produced by the front airfoil. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. Rotor inlet hub/tip ratio was 0.8. Design values of rotor tip velocity and stage pressure ratio were 757 ft/sec and 1.30, respectively.

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.

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.

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.

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.

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.

Leading edge gypsy moth population dynamics

Treesearch

M. R. Carter; F. W. Ravlin; M. L. McManus

1991-01-01

Leading edge gypsy moth populations have been the focus of several intervention programs (MDIPM, AIPM). Knowledge of gypsy moth population dynamics in leading edge area is crucial for effective management. Populations in these areas tend to reach outbreak levels (noticeable defoliation) within three to four years after egg masses are first detected. Pheromone traps...

An investigation on the effect of second-order additional thickness distributions to the upper surface of an NACA 64-206 airfoil. [using flow equations and a CDC 7600 digital 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 an NACA 64-206 airfoil. Additional thickness distributions employed were in the form of two second-order polynomial arcs which have a specified thickness at a given chordwise location. The forward arc disappears at the airfoil leading edge, the aft arc disappears at the airfoil trailing edge. At the juncture of the two arcs, x = x, continuity of slope is maintained. The effect of varying the maximum additional thickness and its chordwise location on airfoil lift coefficient, pitching moment, and pressure distribution was investigated. Results were obtained at a Mach number of 0.2 with an angle-of-attack of 6 degrees on the basic NACA 64-206 airfoil, 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.

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

NASA Technical Reports Server (NTRS)

Bragg, M. B.

1986-01-01

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

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.

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.

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

CFD Study of NACA 0018 Airfoil with Flow Control

NASA Technical Reports Server (NTRS)

Eggert, Christopher A.; Rumsey, Christopher L.

2017-01-01

The abilities of two different Reynolds-Averaged Navier-Stokes codes to predict the effects of an active flow control device are evaluated. The flow control device consists of a blowing slot located on the upper surface of an NACA 0018 airfoil, near the leading edge. A second blowing slot present on the airfoil near mid-chord is not evaluated here. Experimental results from a wind tunnel test show that a slot blowing with high momentum coefficient will increase the lift of the airfoil (compared to no blowing) and delay flow separation. A slot with low momentum coefficient will decrease the lift and induce separation even at low angles of attack. Two codes, CFL3D and FUN3D, are used in two-dimensional computations along with several different turbulence models. Two of these produced reasonable results for this flow, when run fully turbulent. A more advanced transition model failed to predict reasonable results, but warrants further study using different inputs. Including inviscid upper and lower tunnel walls in the simulations was found to be important in obtaining pressure distributions and lift coefficients that best matched experimental data. A limited number of three-dimensional computations were also performed.

The effects of gusts on the fluctuating airloads of airfoils in transonic flow

NASA Technical Reports Server (NTRS)

Mccroskey, W. J.

1984-01-01

Unsteady interactions of distributed and sharp-edged gusts with a stationary airfoil have been analyzed in two-dimensional transonic flow.A simple method of introducing such disturbances has been numerically implemented within the framework of unsteady, transonic small-disturbance theory. Representative solutions for various airfoils subjected to chordwise and transverse gusts show that the strength and unsteady motion of the shock wave on the airfoil significantly affect the flowfield development and, consequently, the dynamic airloads. Also a study was made of the reductions in the unsteady airloads that can be achieved by the proper active control motion of a trailing-edge flap, and a simple gust-alleviation strategy was developed. However, the chordwise pressure distributions associated with gusts are very different from those produced by trailing-edge flap oscillations. Consequently, the fluctuating lift and the unsteady pitching moments cannot both be eliminated simultaneously.

Analysis of high-incidence separated flow past airfoils

NASA Technical Reports Server (NTRS)

Chia, K. N.; Osswald, G. A.; Chia, U.

1989-01-01

An unsteady Navier-Stokes (NS) analysis is developed and used to carefully examine high-incidence aerodynamic separated flows past airfoils. Clustered conformal C-grids are employed for the 12 percent thick symmetric Joukowski airfoil as well as for the NACA 0012 airfoil with a sharp trailing edge. The clustering is controlled by appropriate one-dimensional stretching transformations. An attempt is made to resolve many of the dominant scales of an unsteady flow with massive separation, while maintaining the transformation metrics to be smooth and continuous in the entire flow field. A fully implicit time-marching alternating-direction implicit-block Gaussian elimination (ADI-BGE) method is employed, in which no use is made of any explicit artificial dissipation. Detailed results are obtained for massively separated, unsteady flow past symmetric Joukowski and NACA 0012 airfoils.

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.

An investigation on the effect of second-order additional thickness distributions to the upper surface of an NACA 64 sub 1-212 airfoil. [using flow equations and a CDC 7600 digital computer

NASA Technical Reports Server (NTRS)

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

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 an NACA 64 sub 1 - 212 airfoil. Additional thickness distributions employed were in the form of two second-order polynomial arcs which have a specified thickness at a given chordwise location. The forward arc disappears at the airfoil leading edge, the aft arc disappears at the airfoil trailing edge. At the juncture of the two arcs, x = x, continuity of slope is maintained. The effect of varying the maximum additional thickness and its chordwise location on airfoil lift coefficient, pitching moment, and pressure distribution was investigated. Results were obtained at a Mach number of 0.2 with an angle-of-attack of 6 degrees on the basic NACA 64 sub 1 - 212 airfoil, 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.

Effect of Full-Chord Porosity on Aerodynamic Characteristics of the NACA 0012 Airfoil

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.; Hartwich, Peter M.

1996-01-01

A test was conducted on a model of the NACA 0012 airfoil section with a solid upper surface or a porous upper surface with a cavity beneath for passive venting. The purposes of the test were to investigate the aerodynamic characteristics of an airfoil with full-chord porosity and to assess the ability of porosity to provide a multipoint or self-adaptive design. The tests were conducted in the Langley 8-Foot Transonic Pressure Tunnel over a Mach number range from 0.50 to 0.82 at chord Reynolds numbers of 2 x 10(exp 6), 4 x 10(exp 6), and 6 x 10(exp 6). The angle of attack was varied from -1 deg to 6 deg. At the lower Mach numbers, porosity leads to a dependence of the drag on the normal force. At subcritical conditions, porosity tends to flatten the pressure distribution, which reduces the suction peak near the leading edge and increases the suction over the middle of the chord. At supercritical conditions, the compression region on the porous upper surface is spread over a longer portion of the chord. In all cases, the pressure coefficient in the cavity beneath the porous surface is fairly constant with a very small increase over the rear portion. For the porous upper surface, the trailing edge pressure coefficients exhibit a creep at the lower section normal force coefficients, which suggests that the boundary layer on the rear portion of the airfoil is significantly thickening with increasing normal force coefficient.

Unsteady aerodynamics and vortex-sheet formation of a two-dimensional airfoil

NASA Astrophysics Data System (ADS)

Xia, X.; Mohseni, K.

2017-11-01

Unsteady inviscid flow models of wings and airfoils have been developed to study the aerodynamics of natural and man-made flyers. Vortex methods have been extensively applied to reduce the dimensionality of these aerodynamic models, based on the proper estimation of the strength and distribution of the vortices in the wake. In such modeling approaches, one of the most fundamental questions is how the vortex sheets are generated and released from sharp edges. To determine the formation of the trailing-edge vortex sheet, the classical Kutta condition can be extended to unsteady situations by realizing that a flow cannot turn abruptly around a sharp edge. This condition can be readily applied to a flat plate or an airfoil with cusped trailing edge since the direction of the forming vortex sheet is known to be tangential to the trailing edge. However, for a finite-angle trailing edge, or in the case of flow separation away from a sharp corner, the direction of the forming vortex sheet is ambiguous. To remove any ad-hoc implementation, the unsteady Kutta condition, the conservation of circulation, as well as the conservation laws of mass and momentum are coupled to analytically solve for the angle, strength, and relative velocity of the trailing-edge vortex sheet. The two-dimensional aerodynamic model together with the proposed vortex-sheet formation condition is verified by comparing flow structures and force calculations with experimental results for airfoils in steady and unsteady background flows.

Aerodynamic Characteristics of Airfoils at High Speeds

NASA Technical Reports Server (NTRS)

Briggs, L J; Hull, G F; Dryden, H L

1925-01-01

This report deals with an experimental investigation of the aerodynamical characteristics of airfoils at high speeds. Lift, drag, and center of pressure measurements were made on six airfoils of the type used by the air service in propeller design, at speeds ranging from 550 to 1,000 feet per second. The results show a definite limit to the speed at which airfoils may efficiently be used to produce lift, the lift coefficient decreasing and the drag coefficient increasing as the speed approaches the speed of sound. The change in lift coefficient is large for thick airfoil sections (camber ratio 0.14 to 0.20) and for high angles of attack. The change is not marked for thin sections (camber ratio 0.10) at low angles of attack, for the speed range employed. At high speeds the center of pressure moves back toward the trailing edge of the airfoil as the speed increases. The results indicate that the use of tip speeds approaching the speed of sound for propellers of customary design involves a serious loss in efficiency.

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

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.

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

NASA Astrophysics Data System (ADS)

Mu≁oz, A.; Méndez, B.; Munduate, X.

2016-09-01

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

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.

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

NASA Technical Reports Server (NTRS)

Brent, J. A.; Clemmons, D.

1972-01-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

1987-01-01

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

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.

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.

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.

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

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

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.

3D scanning and printing of airfoils for modular UAS

NASA Astrophysics Data System (ADS)

Dahlgren, Robert P.; Pinsker, Ethan A.; Dary, Omar G.; Ogunbiyi, Joab A.; Mazhari, Arash Alex

2017-02-01

The NASA Ames Research Center has been developing small unmanned airborne systems (UAS) based upon remotecontrolled military aircraft such as the RQ-14 DragonEye and RQ-11 Raven manufactured by AeroVironment. The first step is replacing OEM avionics with COTS avionics that do not use military frequencies for command and control. 3D printing and other rapid prototyping techniques are used to graft RQ-14 components into new "FrankenEye" aircraft and RQ-11 components into new "FrankenRaven" airframes. To that end, it is necessary to design new components to concatenate wing sections into elongated wingspans, construct biplane architectures, attach payload pods, and add control surfaces. When making components such as wing splices it is critical that the curvature and angles of the splice identically match the existing wing at the mating surfaces. The RQ-14 has a thick, simple airfoil with a rectangular planform and no twist or dihedral which make splice development straightforward. On the other hand the RQ-11 has a much thinner sailplane-type airfoil having a tapered polyhedral planform. 3D scanning of the Raven wings with a NextEngine scanner could not capture the complex curvature of the high-performance RQ-11 airfoil, resulting in non-matching and even misshapen splice prototypes. To characterize the airfoil a coordinate measuring machine (CMM) was employed to measure the wing's shape, fiducials and mounting features, enabling capture of the subtle curves of the airfoil and the leading and trailing edges with high fidelity. In conclusion, both rapid and traditional techniques are needed to precisely measure and fabricate wing splice components.

Shock Interaction Control for Scramjet Cowl Leading Edges

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

An Aeroacoustic Characterization of a Multi-Element High-Lift Airfoil

NASA Astrophysics Data System (ADS)

Pascioni, Kyle A.

The leading edge slat of a high-lift system is known to be a large contributor to the overall radiated acoustic field from an aircraft during the approach phase of the flight path. This is due to the unsteady flow field generated in the slat-cove and near the leading edge of the main element. In an effort to understand the characteristics of the flow-induced source mechanisms, a suite of experimental measurements has been performed on a two-dimensional multi-element airfoil, namely, the MD-30P30N. Particle image velocimetry provide mean flow field and turbulence statistics to illustrate the differences associated with a change in angle of attack. Phase-averaged quantities prove shear layer instabilities to be linked to narrowband peaks found in the acoustic spectrum. Unsteady surface pressure are also acquired, displaying strong narrowband peaks and large spanwise coherence at low angles of attack, whereas the spectrum becomes predominately broadband at high angles. Nonlinear frequency interaction is found to occur at low angles of attack, while being negligible at high angles. To localize and quantify the noise sources, phased microphone array measurements are per- formed on the two dimensional high-lift configuration. A Kevlar wall test section is utilized to allow the mean aerodynamic flow field to approach distributions similar to a free-air configuration, while still capable of measuring the far field acoustic signature. However, the inclusion of elastic porous sidewalls alters both aerodynamic and acoustic characteristics. Such effects are considered and accounted for. Integrated spectra from Delay and Sum and DAMAS beamforming effectively suppress background facility noise and additional noise generated at the tunnel wall/airfoil junction. Finally, temporally-resolved estimates of a low-dimensional representation of the velocity vector fields are obtained through the use of proper orthogonal decomposition and spectral linear stochastic estimation. An

Simulation of Acoustic Scattering from a Trailing Edge

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

Experimental and Analytical Investigation of the Coolant Flow Characteristics in Cooled Turbine Airfoils

NASA Technical Reports Server (NTRS)

Damerow, W. P.; Murtaugh, J. P.; Burggraf, F.

1972-01-01

The flow characteristics of turbine airfoil cooling system components were experimentally investigated. Flow models representative of leading edge impingement, impingement with crossflow (midchord cooling), pin fins, feeder supply tube, and a composite model of a complete airfoil flow system were tested. Test conditions were set by varying pressure level to cover the Mach number and Reynolds number range of interest in advanced turbine applications. Selected geometrical variations were studied on each component model to determine these effects. Results of these tests were correlated and compared with data available in the literature. Orifice flow was correlated in terms of discharge coefficients. For the leading edge model this was found to be a weak function of hole Mach number and orifice-to-impinged wall spacing. In the impingement with crossflow tests, the discharge coefficient was found to be constant and thus independent of orifice Mach number, Reynolds number, crossflow rate, and impingement geometry. Crossflow channel pressure drop showed reasonable agreement with a simple one-dimensional momentum balance. Feeder tube orifice discharge coefficients correlated as a function of orifice Mach number and the ratio of the orifice-to-approach velocity heads. Pin fin data was correlated in terms of equivalent friction factor, which was found to be a function of Reynolds number and pin spacing but independent of pin height in the range tested.

Advancements in adaptive aerodynamic technologies for airfoils and wings

NASA Astrophysics Data System (ADS)

Jepson, Jeffrey Keith

Although aircraft operate over a wide range of flight conditions, current fixed-geometry aircraft are optimized for only a few of these conditions. By altering the shape of the aircraft, adaptive aerodynamics can be used to increase the safety and performance of an aircraft by tailoring the aircraft for multiple flight conditions. Of the various shape adaptation concepts currently being studied, the use of multiple trailing-edge flaps along the span of a wing offers a relatively high possibility of being incorporated on aircraft in the near future. Multiple trailing-edge flaps allow for effective spanwise camber adaptation with resulting drag benefits over a large speed range and load alleviation at high-g conditions. The research presented in this dissertation focuses on the development of this concept of using trailing-edge flaps to tailor an aircraft for multiple flight conditions. One of the major tasks involved in implementing trailing-edge flaps is in designing the airfoil to incorporate the flap. The first part of this dissertation presents a design formulation that incorporates aircraft performance considerations in the inverse design of low-speed laminar-flow adaptive airfoils with trailing-edge cruise flaps. The benefit of using adaptive airfoils is that the size of the low-drag region of the drag polar can be effectively increased without increasing the maximum thickness of the airfoil. Two aircraft performance parameters are considered: level-flight maximum speed and maximum range. It is shown that the lift coefficients for the lower and upper corners of the airfoil low-drag range can be appropriately adjusted to tailor the airfoil for these two aircraft performance parameters. The design problem is posed as a part of a multidimensional Newton iteration in an existing conformal-mapping based inverse design code, PROFOIL. This formulation automatically adjusts the lift coefficients for the corners of the low-drag range for a given flap deflection as

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

NASA Technical Reports Server (NTRS)

Ashby, Dale L.

1996-01-01

An experimental and computational investigation of the effect of lift-enhancing tabs on a two-element airfoil has been conducted. The objective of the study was to develop an understanding of the flow physics associated with lift-enhancing tabs on a multi-element airfoil. An NACA 63(2)-215 ModB airfoil with a 30% chord fowler flap was tested in the NASA Ames 7- by 10-Foot Wind Tunnel. Lift-enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. A combination of tabs located at the main element and flap trailing edges increased the airfoil lift coefficient by 11% relative to the highest lift coefficient achieved by any baseline configuration at an angle of attack of 0 deg, and C(sub 1max) was increased by 3%. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computed results predicted all of the trends observed in the experimental data quite well. In addition, a simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift-enhancing tabs work. The tabs were modeled by a point vortex at the air-foil or flap trailing edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift-enhancing tabs on a multi-element airfoil. Results of the modeling effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

Span efficiency of wings with leading edge protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2013-11-01

Past work has shown that sinusoidal leading edge protuberances resembling those found on humpback whale flippers alter the lift and drag coefficients of full- and finite-span foils and wings depending on the angle of attack and leading edge geometry. Although the load characteristics of protuberance modified finite-span wings have been reported for flipper-like geometries at higher Reynolds numbers and for rectangular planforms at lower Reynolds numbers, the effects of leading edge geometry on the span efficiency, which is indicative of the deviation of the spanwise lift distribution from elliptical and the viscous effects, for a range of planforms and Reynolds numbers have not been addressed. The lift and drag coefficients of 7 rectangular, 2 swept, and 2 flipper-like planform models with aspect ratios of 4.3, 4.0, and 8.86, respectively, were used to compute the span efficiency at Reynolds numbers ranging from 0.9 to 4.5 × 105. The span efficiency, based on the data at lower angles of attack, of modified wings was compared with the unmodified models. For the cases considered, the span efficiencies of the leading edge modified models were less than those of the equivalent unmodified models. The dependence of span efficiency on the leading edge geometry, planform, and Reynolds number will be presented. Supported by the ONR-ULI program.

Turbulence Measurements on a Flap-Edge Model

NASA Technical Reports Server (NTRS)

Moriarty, Patrick; Bradshaw, Peter; Cantwell, Brian; Ross, James

1998-01-01

Turbulence measurements have been made on a flap-edge and leading-edge slat model using hot-wire anemometry, and, later, particle image velocimetry. The properties of hot-wire anemometry were studied using facilities at NASA Ames Research Center. Hot-film probes were used because of their durability, but cross-films were limited by non-linear end effects. As a warm-up exercise, hot-film probes were used to measure velocities in the farfield wake of a cylinder with an airfoil in the near-field wake. The airfoil reduced the drag coefficient of the system by 10%. A single-wire hot-film probe was used to measure velocity profiles over the top of a NACA 63(sub 2)-215 Mod. B wing with a Fowler flap and leading,-edge slat. Results showed the size of slat wake was dependent upon the slat deflection angle. Velocity increased through the slat gap with increased deflection. The acoustically modified slat decreased the chance of separation. Measurements were taken at the flap edge with a single hot-film. Trends in the data indicate velocity and turbulence levels increase at the flap edge. The acoustically modified flap modifies the mean flow near the flap edge. Correlations were made between the hot-film signal and the unsteady pressure transducers on the wing which were published in a NASA CDTM. The principles of Particle Image Velocimetry (PIV) were studied at Florida State University. Spectral PIV was used to measure the spectra of a subsonic jet. Measured frequencies were close to the predicted frequency of jet shedding. Spectral PIV will be used to measure the spectra of the slat flow in the second 7 x lO-ft. wind tunnel test. PIV has an advantage that it can measure velocity and spectra of the entire flowfield instantaneously. However, problems arise when trying, to store this massive amount of PIV data. Support for this research has continued through a NASA Graduate Student Program Fellowship which will end in June 1999. The thesis should be completed by this time.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Barone, Matthew Franklin

2011-08-01

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

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

NASA Technical Reports Server (NTRS)

Adair, Desmond; Horne, W. Clifton

1988-01-01

Flow characteristics in the vicinity of the trailing edge of a single-slotted airfoil flap are presented and analyzed. The experimental arrangement consisted of a NACA 4412 airfoil equipped with a NACA 4415 flap whose angle of deflection was 21.8 deg. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The flow was complicated by the presence of a strong, initially inviscid jet emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer. Downstream of the flap, the airfoil and flap wakes fully merged to form an asymmetrical curved wake. The airfoil configuration was tested at an angle of attack of 8.2 deg, at a Mach number of 0.09, and a chord based Reynolds number of 1.8 x 10 to the 6th power in the Ames Research Center 7- by 10-Foot Wind Tunnel. Surface pressure measurements were made on the airfoil and flap and on the wind tunnel roof and floor. It was estimated that the wall interference increased the C sub L by 7 percent and decreased the C sub M by 4.5 percent. Velocity characteristics were quantified using hot-wire anemometry in regions of flow with preferred direction and low turbulence intensity. A 3-D laser velocimeter was used in regions of flow recirculation and relatively high turbulence intensity.

Heat-Pipe-Cooled Leading Edges for Hypersonic Vehicles

NASA Technical Reports Server (NTRS)

Glass, David E.

2006-01-01

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

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

NASA Technical Reports Server (NTRS)

Ashby, Dale

1996-01-01

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

Asymptotic theory of two-dimensional trailing-edge flows

NASA Technical Reports Server (NTRS)

Melnik, R. E.; Chow, R.

1975-01-01

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

Aerodynamic sound of flow past an airfoil

NASA Technical Reports Server (NTRS)

Wang, Meng

1995-01-01

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

The Influence of Chordwise Flexibility on the Flow Structure and Streamwise Force of a Sinusoidally Pitching Airfoil

NASA Astrophysics Data System (ADS)

Olson, David Arthur

Many natural flyers and swimmers need to exploit unsteady mechanisms in order to generate sufficient aerodynamic forces for sustained flight and propulsion. This is, in part, due to the low speed and length scales at which they typically operate. In this low Reynolds number regime, there are many unanswered questions on how existing aerodynamic theory for both steady and unsteady flows can be applied. Additionally, most of these natural flyers and swimmers have deformable wing/fin structures, three dimensional wing planforms, and exhibit complex kinematics during motion. While some biologically-inspired studies seek to replicate these complex structures and kinematics in the laboratory or in numerical simulations, it becomes difficult to draw explicit connections to the existing knowledge base of classical unsteady aerodynamic theory due to the complexity of the problems. In this experimental study, wing kinematics, structure, and planform are greatly simplified to investigate the effect of chordwise flexibility on the streamwise force (thrust) and wake behavior of a sinusoidally pitching airfoil. The study of flexibility in the literature has typically utilized flat plates with varying thicknesses or lengths to change their chordwise flexibility. This choice introduces additional complexities when comparing to the wealth of knowledge originally developed on streamlined aerodynamic shapes. The current study capitalizes on the recent developments in 3D printer technology to create accurate shapes out of materials with varying degrees of flexibility by creating two standard NACA 0009 airfoils: one rigid and one flexible. Each of the two airfoils are sinusoidally pitched about the quarter chord over a range of oscillation amplitudes and frequencies while monitoring the deformation of the airfoil. The oscillation amplitude is selected to be small enough such that leading edge vortices do not form, and the vortical structures in the wake are formed from the trailing

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.

Active Control of Separation From the Flap of a Supercritical Airfoil

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Yao, Chung-Sheng; Seifert, Avi

2006-01-01

Zero-mass-flux periodic excitation was applied at several regions on a simplified high-lift system to delay the occurrence of flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approximately equal to 10) and low frequency amplitude modulation (F(+) sub AM approximately equal to 1) of the high frequency excitation were used for control. It was noted that the same performance gains were obtained with amplitude modulation and required only 30% of the momentum input required by pure sine excitation.

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.

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

NASA Technical Reports Server (NTRS)

Mueller, T. J.

1986-01-01

A semi-empirical method for predicting separation bubble characteristics was evaluated using low Reynolds number test data. On the basis of this data, several observations were made. First, a sizable growth in the momentum thickness can occur in the laminar portion of a separation bubble. This is in direct contrast to the theory and is apparently due to low Reynolds number effects. Secondly, the transition Reynolds number (R sub l sub 1) which governs the extent of a bubble's laminar region, was found to be much lower than that used in the method. At present, there does not seem to be any evidence supporting a single value for R sub l sub 1. Apparently, R sub l sub 1 is affected by the freestream disturbance environment, and airfoil's pressure distribution, and possibly the chord Reynolds number as well. Thirdly, the growth in momentum thickness over a bubble's turbulent region was predicted reasonably well by the method, provided that Roberts' suggested value for the mean dissipation coefficient was used. Finally, the present data does not substantiate the universality of the velocity profile at reattachment. However, measurement error may be responsible for this result.

Steady Aerodynamic Characteristics of Two-Dimensional NACA0012 Airfoil for One Revolution Angle of Attack

NASA Astrophysics Data System (ADS)

Park, Byung Ho; Han, Yong Oun

2018-04-01

Steady variations in aerodynamic forces and flow behaviors of two-dimensional NACA0012 airfoil were investigated using a numerical method for One Revolution Angle of Attack (AOA) at Reynolds number of 105 . The profiles of lift coefficients, drag coefficients, and pressure coefficients were compared with those of the experimental data. The AERODAS model was used to analyze the profiles of lift and drag coefficients. Wake characteristics were given along with the deficit profiles of incoming velocity components. Both the characteristics of normal and reverse airfoil models were compared with the basic aerodynamic data for the same range of AOA. The results show that two peaks of the lift coefficients appeared at 11.5{°} and 42{°} and are in good agreement with the pre-stall and post-stall models, respectively. Counter-rotating vortex flows originated from the leading and trailing edges at a high AOA, which formed an impermeable zone over the suction surface and made reattachments in the wake. Moreover, the acceleration of inflow along the boundary of the vortex wrap appeared in the profile of the wake velocity. The drag profile was found to be independent of the airfoil mode, but the lift profile was quite sensitive to the airfoil mode.

Leading-edge receptivity for blunt-nose bodies

NASA Technical Reports Server (NTRS)

Kerschen, Edward J.

1991-01-01

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

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

NASA Technical Reports Server (NTRS)

Ratnayake, Nalin A.

2009-01-01

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

An experimental investigation of multi-element airfoil ice accretion and resulting performance degradation

NASA Technical Reports Server (NTRS)

Potapczuk, Mark G.; Berkowitz, Brian M.

1989-01-01

An investigation of the ice accretion pattern and performance characteristics of a multi-element airfoil was undertaken in the NASA Lewis 6- by 9-Foot Icing Research Tunnel. Several configurations of main airfoil, slat, and flaps were employed to examine the effects of ice accretion and provide further experimental information for code validation purposes. The text matrix consisted of glaze, rime, and mixed icing conditions. Airflow and icing cloud conditions were set to correspond to those typical of the operating environment anticipated tor a commercial transport vehicle. Results obtained included ice profile tracings, photographs of the ice accretions, and force balance measurements obtained both during the accretion process and in a post-accretion evaluation over a range of angles of attack. The tracings and photographs indicated significant accretions on the slat leading edge, in gaps between slat or flaps and the main wing, on the flap leading-edge surfaces, and on flap lower surfaces. Force measurments indicate the possibility of severe performance degradation, especially near C sub Lmax, for both light and heavy ice accretion and performance analysis codes presently in use. The LEWICE code was used to evaluate the ice accretion shape developed during one of the rime ice tests. The actual ice shape was then evaluated, using a Navier-Strokes code, for changes in performance characteristics. These predicted results were compared to the measured results and indicate very good agreement.

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.

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.

Airfoil gust response and the sound produced by airifoil-vortex interaction

NASA Technical Reports Server (NTRS)

Amiet, R. K.

1986-01-01

This paper contributes to the understanding of the noise generation process of an airfoil encountering an unsteady upwash. By using a fast Fourier transform together with accurate airfoil response functions, the lift-time waveform for an airfoil encountering a delta function gust (the indicial function) is calculated for a flat plate airfoil in a compressible flow. This shows the interesting property that the lift is constant until the generated acoustic wave reaches the trailing edge. Expressions are given for the magnitude of this constant and for the pressure distribution on the airfoil during this time interval. The case of an airfoil cutting through a line vortex is also analyzed. The pressure-time waveform in the far field is closely related to the left-time waveform for the above problem of an airfoil entering a delta function gust. The effects of varying the relevant parameters in the problem are studied, including the observed position, the core diameter of the vortex, the vortex orientation and the airfoil span. The far field sound varies significantly with observer position, illustrating the importance of non-compactness effects. Increasing the viscous core diameter tends to smooth the pressure-time waveform. For small viscous core radius and infinite span, changing the vortex orientation changes only the amplitude of the pressure-time waveform, and not the shape.

Theory of viscous transonic flow over airfoils at high Reynolds number

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

Determination of the Pressure Drag of Airfoils by Integration of Surface Pressures

NASA Technical Reports Server (NTRS)

Phillips, William H.

1990-01-01

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

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

NASA Technical Reports Server (NTRS)

Sorenson, R. L.

1982-01-01

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

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.

Trailing Edge Noise Prediction Based on a New Acoustic Formulation

NASA Technical Reports Server (NTRS)

Casper, J.; Farassat, F.

2002-01-01

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

Broadband Trailing Edge Noise Predictions in the Time Domain. Revised

NASA Technical Reports Server (NTRS)

Casper, Jay; Farassat, Fereidoun

2003-01-01

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

Experimental and simulated control of lift using trailing edge devices

NASA Astrophysics Data System (ADS)

Cooperman, A.; Blaylock, M.; van Dam, C. P.

2014-12-01

Two active aerodynamic load control (AALC) devices coupled with a control algorithm are shown to decrease the change in lift force experienced by an airfoil during a change in freestream velocity. Microtabs are small (1% chord) surfaces deployed perpendicular to an airfoil, while microjets are pneumatic jets with flow perpendicular to the surface of the airfoil near the trailing edge. Both devices are capable of producing a rapid change in an airfoil's lift coefficient. A control algorithm for microtabs has been tested in a wind tunnel using a modified S819 airfoil, and a microjet control algorithm has been simulated for a NACA 0012 airfoil using OVERFLOW. In both cases, the AALC devices have shown the ability to mitigate the changes in lift during a gust.

The leading-edge vortex of yacht sails

NASA Astrophysics Data System (ADS)

Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

We experimentally show, for the first time, that a stable Leading-Edge Vortex (LEV) can be formed on an asymmetric spinnaker, which is a high-lift sail used by yachts to sail downwind. We tested a 3D printed rigid sail in a water flume at a chord-based Reynolds number of ca. 104. We found that on the leeward side of the sail (the suction side), the flow separates at the leading edge reattaching further downstream and forming a stable LEV. The LEV grows in diameter from the root to the tip of the sail, where it merges with the tip vortex. We detected the LEV using the γ criterion, and we verified its stability over time. The lift contribution provided by the LEV was computed solving a complex potential model of each sail section. This analysis indicated that the LEV provides a substantial contribution to the total sail's lift. These findings suggest that the maximum lift of low-aspect-ratio wings with a sharp leading edge, such as spinnakers, can be enhanced by promoting a stable LEV. This work was funded by the Consejo Nacional de Ciencia y Tecnologia (CONACYT).

Measuremants in the wake of an infinite swept airfoil

NASA Technical Reports Server (NTRS)

Novak, C. J.; Ramaprian, B. R.

1982-01-01

This is a report of the measurements in the trailing edge region as well as in the report of the developing wake behind a swept NACA 0012 airfoil at zero incidence and a sweep angle of 30 degrees. The measurements include both the mean and turbulent flow properties. The mean flow velocities, flow inclination and static pressure are measured using a calibrated three-hole yaw probe. The measurements of all the relevant Reynolds stress components in the wake are made using a tri-axial hot-wire probe and a digital data processing technique developed by the authors. The development of the three dimensional near-wake into a nearly two dimensional far-wake is discussed in the light of the experimental data. A complete set of wake data along with the data on the initial boundary layer in the trailing edge region of the airfoil are tabulated in an appendix to the report.

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.

Low-speed single-element airfoil synthesis

NASA Technical Reports Server (NTRS)

Mcmasters, J. H.; Henderson, M. L.

1979-01-01

The use of recently developed airfoil analysis/design computational tools to clarify, enrich and extend the existing experimental data base on low-speed, single element airfoils is demonstrated. A discussion of the problem of tailoring an airfoil for a specific application at its appropriate Reynolds number is presented. This problem is approached by use of inverse (or synthesis) techniques, wherein a desirable set of boundary layer characteristics, performance objectives, and constraints are specified, which then leads to derivation of a corresponding viscous flow pressure distribution. Examples are presented which demonstrate the synthesis approach, following presentation of some historical information and background data which motivate the basic synthesis process.

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.

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.

Study of 3-D Dynamic Roughness Effects on Flow Over a NACA 0012 Airfoil Using Large Eddy Simulations at Low Reynolds Numbers

NASA Astrophysics Data System (ADS)

Guda, Venkata Subba Sai Satish

There have been several advancements in the aerospace industry in areas of design such as aerodynamics, designs, controls and propulsion; all aimed at one common goal i.e. increasing efficiency --range and scope of operation with lesser fuel consumption. Several methods of flow control have been tried. Some were successful, some failed and many were termed as impractical. The low Reynolds number regime of 104 - 105 is a very interesting range. Flow physics in this range are quite different than those of higher Reynolds number range. Mid and high altitude UAV's, MAV's, sailplanes, jet engine fan blades, inboard helicopter rotor blades and wind turbine rotors are some of the aerodynamic applications that fall in this range. The current study deals with using dynamic roughness as a means of flow control over a NACA 0012 airfoil at low Reynolds numbers. Dynamic 3-D surface roughness elements on an airfoil placed near the leading edge aim at increasing the efficiency by suppressing the effects of leading edge separation like leading edge stall by delaying or totally eliminating flow separation. A numerical study of the above method has been carried out by means of a Large Eddy Simulation, a mathematical model for turbulence in Computational Fluid Dynamics, owing to the highly unsteady nature of the flow. A user defined function has been developed for the 3-D dynamic roughness element motion. Results from simulations have been compared to those from experimental PIV data. Large eddy simulations have relatively well captured the leading edge stall. For the clean cases, i.e. with the DR not actuated, the LES was able to reproduce experimental results in a reasonable fashion. However DR simulation results show that it fails to reattach the flow and suppress flow separation compared to experiments. Several novel techniques of grid design and hump creation are introduced through this study.

Composite airfoil assembly

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garcia-Crespo, Andres Jose

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.

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2008-11-01

The agility of the humpback whale has been attributed to the use of its pectoral flippers, on which protuberances are present along the leading edge. The forces and moments on hydrofoils with leading edge protuberances were measured in a water tunnel and were compared to a baseline NACA 63(4)-021 hydrofoil revealing significant performance differences. Three protuberance amplitudes and two spanwise wavelengths, closely resembling the morphology found in nature, were examined. Qualitative flow visualization techniques were used to examine flow patterns surrounding the hydrofoils, and Particle Image Velocimetry (PIV) was used to quantify these patterns. Flow visualizations have revealed counter-rotating vortices stemming from the shoulders of the protuberances. These streamwise vortices are a result of the spanwise pressure gradient brought about by the varying leading edge curvature. PIV was used to quantify the strength of these vortices as a function of angle of attack and leading edge geometry. At low angles of attack, these vortices are symmetric with respect to the protuberances; however, the symmetry is lost at high angles of attack. The loss of symmetry can be correlated with the separation point location on the hydrofoil.

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.

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

NASA Technical Reports Server (NTRS)

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

2010-01-01

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

The computation of the post-stall behavior of a circulation controlled airfoil

NASA Technical Reports Server (NTRS)

Linton, Samuel W.

1993-01-01

The physics of the circulation controlled airfoil is complex and poorly understood, particularly with regards to jet stall, which is the eventual breakdown of lift augmentation by the jet at some sufficiently high blowing rate. The present paper describes the numerical simulation of stalled and unstalled flows over a two-dimensional circulation controlled airfoil using a fully implicit Navier-Stokes code, and the comparison with experimental results. Mach numbers of 0.3 and 0.5 and jet total to freestream pressure ratios of 1.4 and 1.8 are investigated. The Baldwin-Lomax and k-epsilon turbulence models are used, each modified to include the effect of strong streamline curvature. The numerical solutions of the post-stall circulation controlled airfoil show a highly regular unsteady periodic flowfield. This is the result of an alternation between adverse pressure gradient and shock induced separation of the boundary layer on the airfoil trailing edge.

Cavitation on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

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

2012-11-01

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

Effect of controlled spanwise bending on the stability of the leading-edge vortex

NASA Astrophysics Data System (ADS)

Bhattacharya, Samik; Scofield, Tyler

2017-11-01

When an airfoil is accelerated from rest at a high angle of attack, a leading-edge vortex (LEV) forms, which soon gets destabilized and convects downstream. In this work, we control the spanwise bending of a flat plate wing to actively influence the vorticity transfer from the LEV. Our aim is to investigate the effect of spanwise curvature variation on the geometry, growth, and stability of the LEV during the acceleration phase. A 3D printed flat-plate with a chord of 5 cm and span of 15 cm is towed in a small fish tank at different angles of attack greater than 15°. The plate starts from rest and reaches a Reynolds number of 5000 after travelling different multiples and submultiples of chord-length. We carry out dye-flow visualization and measure the circulation build up and the convection velocity of the LEV with the help of particle image velocimetry (PIV). The unsteady loads coming on to the wing is measured with a force sensor. An analytical scheme for computing the load from the measured displacement of the plate is presented and compared with the force sensor data. Preliminary results indicate that controlled curvature variation can influence the formation and stability of an LEV.

Turbine blade squealer tip rail with fence members

DOE Office of Scientific and Technical Information (OSTI.GOV)

Little, David A

2012-11-20

A turbine blade includes an airfoil, a blade tip section, a squealer tip rail, and a plurality of chordally spaced fence members. The blade tip section includes a blade tip floor located at an end of the airfoil distal from the root. The blade tip floor includes a pressure side and a suction side joined together at chordally spaced apart leading and trailing edges of the airfoil. The squealer tip rail extends radially outwardly from the blade tip floor adjacent to the suction side and extends from a first location adjacent to the airfoil trailing edge to a second locationmore » adjacent to the airfoil leading edge. The fence members are located between the airfoil leading and trailing edges and extend radially outwardly from the blade tip floor and axially from the squealer tip rail toward the pressure side.« less

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

NASA Astrophysics Data System (ADS)

Modlin, James Michael

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

Thermal management of tungsten leading edges in DIII-D

DOE PAGES

Nygren, Richard E.; Rudakov, Dmitry L.; Murphy, Christopher; ...

2017-04-29

The DiMES materials probe exposed tungsten blocks with 0.3 and 1 mm high leading edges to DIII-D He plasmas in 2015 and 2016 viewed with high resolution IRTV. The 1-mm edge may have reached >2400° C in a 3-s shot with a (parallel) heat load of ~50 MW/m 2 and ~10 MW/m 2 on the surface based on modeling. The experiments support ITER. Leading edges were also a concern in the DIII-D Metal Tile Experiment in 2016. Two toroidal rings of divertor tiles had W-coated molybdenum inserts 50 mm wide radially. This study presents data and thermal analyses.

Thermal management of tungsten leading edges in DIII-D

DOE Office of Scientific and Technical Information (OSTI.GOV)

Nygren, Richard E.; Rudakov, Dmitry L.; Murphy, Christopher

The DiMES materials probe exposed tungsten blocks with 0.3 and 1 mm high leading edges to DIII-D He plasmas in 2015 and 2016 viewed with high resolution IRTV. The 1-mm edge may have reached >2400° C in a 3-s shot with a (parallel) heat load of ~50 MW/m 2 and ~10 MW/m 2 on the surface based on modeling. The experiments support ITER. Leading edges were also a concern in the DIII-D Metal Tile Experiment in 2016. Two toroidal rings of divertor tiles had W-coated molybdenum inserts 50 mm wide radially. This study presents data and thermal analyses.

An investigation of unsteady 3D effects on trailing edge flaps

NASA Astrophysics Data System (ADS)

Jost, E.; Fischer, A.; Lutz, T.; Krämer, E.

2016-09-01

The present study investigates the impact of unsteady and viscous three-dimensional aerodynamic effects on a wind turbine blade with trailing edge flap by means of CFD. Harmonic oscillations are simulated on the DTU 10 MW rotor with a flap of 10% chord extent ranging from 70% to 80% blade radius. The deflection frequency is varied in the range between 1p and 6p. To quantify 3D effects, rotor simulations are compared to 2D airfoil computations. A significant influence of trailing and shed vortex structures has been found which leads to a reduction of the lift amplitude and hysteresis effects in the lift response with regard to the flap deflection. In the 3D rotor results greater amplitude reductions and less hystereses have been found compared to the 2D airfoil simulations.

On the study of wavy leading-edge vanes to achieve low fan interaction noise

NASA Astrophysics Data System (ADS)

Tong, Fan; Qiao, Weiyang; Xu, Kunbo; Wang, Liangfeng; Chen, Weijie; Wang, Xunnian

2018-04-01

The application of wavy leading-edge vanes to reduce a single-stage axial fan noise is numerically studied. The aerodynamic and acoustic performance of the fan is numerically investigated using a hybrid unsteady Reynolds averaged Navier-Stokes (URANS)/acoustic analogy method (Goldstein equations). First, the hybrid URANS/Goldstein method is developed and successfully validated against experiment results. Next, numerical simulations are performed to investigate the noise reduction effects of the wavy leading-edge vanes. The aerodynamic and acoustic performance is assessed for a fan with vanes equipped with two different wavy leading-edge profiles and compared with the performance of conventional straight leading-edge vanes. Results indicate that a fan with wavy leading-edge vanes produces lower interaction noise than the baseline fan without a significant loss in aerodynamic performance. In fact, it is demonstrated that wavy leading-edge vanes have the potential to lead to both aerodynamic and acoustic improvements. The two different wavy leading-edge profiles are shown to successfully reduce the fan tone sound power level by 1.2 dB and 4.3 dB, respectively. Fan efficiency is also improved by about 1% with one of the tested wavy leading-edge profiles. Large eddy simulation (LES) is also performed for a simplified fan stage model to assess the effects of wavy leading-edge vanes on the broadband fan noise. Results indicate that the overall sound power level of a fan can be reduced by about 4 dB with the larger wavy leading-edge profile. Finally, the noise reduction mechanisms are investigated and analysed. It is found that the wavy leading-edge profiles can induce significant streamwise vorticity around the leading-edge protuberances and reduce pressure fluctuations (especially at locations of wavy leading-edge hills) and unsteady forces on the stator vanes. The underlying mechanism of the reduced pressure fluctuations is also discussed by examining the magnitude

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Somers, D. M.

2005-01-01

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

Vortex interaction with a leading-edge of finite thickness

NASA Technical Reports Server (NTRS)

Sohn, D.; Rockwell, Donald

1987-01-01

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

Theory of step on leading edge of negative corona current pulse

NASA Astrophysics Data System (ADS)

Gupta, Deepak K.; Mahajan, Sangeeta; John, P. I.

2000-03-01

Theoretical models taking into account different feedback source terms (e.g., ion-impact electron emission, photo-electron emission, field emission, etc) have been proposed for the existence and explanation of the shape of negative corona current pulse, including the step on the leading edge. In the present work, a negative corona current pulse with the step on the leading edge is obtained in the presence of ion-impact electron emission feedback source only. The step on the leading edge is explained in terms of the plasma formation process and enhancement of the feedback source. Ionization wave-like movement toward the cathode is observed after the step. The conditions for the existence of current pulse, with and without the step on the leading edge, are also described. A qualitative comparison with earlier theoretical and experimental work is also included.

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

PubMed

Johnston, Joe; Gopalarathnam, Ashok

2012-09-01

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

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.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading-edge thrust are possible at supersonic speeds for certain planforms having the geometry to support the theoretical thrust-distribution potential. The new analytical process employed provides not only the level of leading-edge thrust attainable but also the spanwise distribution of both it and that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

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.

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

NASA Technical Reports Server (NTRS)

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

1959-01-01

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

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

NASA Technical Reports Server (NTRS)

Shrewsbury, George D.

1987-01-01

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

Aeroacoustic interaction of a distributed vortex with a lifting Joukowski airfoil

NASA Technical Reports Server (NTRS)

Hardin, J. C.; Lamkin, S. L.

1984-01-01

A first principles computational aeroacoustics calculation of the flow and noise fields produced by the interaction of a distributed vortex with a lifting Joukowski airfoil is accomplished at the Reynolds number of 200. The case considered is that where the circulations of the vortex and the airfoil are of opposite sign, corresponding to blade vortex interaction on the retreating side of a single helicopter rotor. The results show that the flow is unsteady, even in the absence of the incoming vortex, resulting in trailing edge noise generation. After the vortex is input, it initially experiences a quite rapid apparent diffusion rate produced by stretching in the airfoil velocity gradients. Consideration of the effects of finite vortex size and viscosity causes the noise radiation during the encounter to be much less impulsive than predicted by previous analyses.

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Aerodynamic characteristics of two rotorcraft airfoils designed for application to the inboard region of a main rotor blade

NASA Technical Reports Server (NTRS)

Noonan, Kevin W.

1990-01-01

A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of two new rotorcraft airfoils designed especially for application to the inboard region of a helicopter main rotor blade. The two new airfoils, the RC(4)-10 and RC(5)-10, and a baseline airfoil, the VR-7, were all studied in the Langley Transonic Tunnel at Mach nos. from about 0.34 to 0.84 and at Reynolds nos. from about 4.7 to 9.3 x 10 (exp 6). The VR-7 airfoil had a trailing edge tab which is deflected upwards 4.6 degs. In addition, the RC(4)-10 airfoil was studied in the Langley Low Turbulence Pressure Tunnel at Mach nos. from 0.10 to 0.44 and at Reynolds nos. from 1.4 to 5.4 x 10 (exp 6) respectively. Some comparisons were made of the experimental data for the new airfoils and the predictions of two different theories. The results of this study indicates that both of the new airfoils offer advantages over the baseline airfoil. These advantages are discussed.

NASA supercritical airfoils: A matrix of family-related airfoils

NASA Technical Reports Server (NTRS)

Harris, Charles D.

1990-01-01

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

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

NASA Technical Reports Server (NTRS)

Graham, Donald J

1948-01-01

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

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.

A Wind-Tunnel Investigation of the Application of the NASA Supercritical Airfoil to a Variable-Wing-Sweep Fighter Airplane

NASA Technical Reports Server (NTRS)

Ayers, T. G.

1973-01-01

An investigation was conducted in the Langley 8 foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel to evaluate the effectiveness of three variations of the NASA supercritical airfoil as applied to a model of a variable wing sweep fighter airplane. Wing panels incorporating conventional NACA 64A series airfoil with 0.20 and 0.40 camber were used as bases of reference for this evaluation. Static force and moment measurements were obtained for wing leading edge sweep angles of 26, 33, 39, and 72.5 degrees. Fluctuating wing root bending moment data were obtained at subsonic speeds to determine buffet characteristics. Subsonic data were also obtained for determining the effects of wing transition location and spoiler deflection. Limited lateral directional data are included for the conventional 0.20 cambered wing and the supercritical wing.

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.

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.

Leading edge flap system for aircraft control augmentation

NASA Technical Reports Server (NTRS)

Rao, D. M. (Inventor)

1984-01-01

Traditional roll control systems such as ailerons, elevons or spoilers are least effective at high angles of attack due to boundary layer separation over the wing. This invention uses independently deployed leading edge flaps on the upper surfaces of vortex stabilized wings to shift the center of lift outboard. A rolling moment is created that is used to control roll in flight at high angles of attack. The effectiveness of the rolling moment increases linearly with angle of attack. No adverse yaw effects are induced. In an alternate mode of operation, both leading edge flaps are deployed together at cruise speeds to create a very effective airbrake without appreciable modification in pitching moment. Little trim change is required.

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

NASA Astrophysics Data System (ADS)

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

2015-11-01

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

A direct-inverse method for transonic and separated flows about airfoils

NASA Technical Reports Server (NTRS)

Carlson, K. D.

1985-01-01

A direct-inverse technique and computer program called TAMSEP that can be sued for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicing the flowfield about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.

A direct-inverse method for transonic and separated flows about airfoils

NASA Technical Reports Server (NTRS)

Carlson, Leland A.

1990-01-01

A direct-inverse technique and computer program called TAMSEP that can be used for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicting the flow field about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.

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.

Flight Tests of a Supersonic Natural Laminar Flow Airfoil

NASA Technical Reports Server (NTRS)

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

2014-01-01

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

Experiments with an Airfoil Model on which the Boundary Layers are Controlled Without the Use of Supplementary Equipment

NASA Technical Reports Server (NTRS)

Abbott, I H

1931-01-01

This report describes test made in the Variable Density Wind Tunnel of the NACA to determine the possibility of controlling the boundary layer on the upper surface of an airfoil by use of the low pressure existing near the leading edge. The low pressure was used to induce flow through slots in the upper surface of the wing. The tests showed that the angle of attack for maximum lift was increased at the expense of a reduction in the maximum lift coefficient and an increase in the drag coefficient.

The fish tail motion forms an attached leading edge vortex

PubMed Central

Borazjani, Iman; Daghooghi, Mohsen

2013-01-01

The tail (caudal fin) is one of the most prominent characteristics of fishes, and the analysis of the flow pattern it creates is fundamental to understanding how its motion generates locomotor forces. A mechanism that is known to greatly enhance locomotor forces in insect and bird flight is the leading edge vortex (LEV) reattachment, i.e. a vortex (separation bubble) that stays attached at the leading edge of a wing. However, this mechanism has not been reported in fish-like swimming probably owing to the overemphasis on the trailing wake, and the fact that the flow does not separate along the body of undulating swimmers. We provide, to our knowledge, the first evidence of the vortex reattachment at the leading edge of the fish tail using three-dimensional high-resolution numerical simulations of self-propelled virtual swimmers with different tail shapes. We show that at Strouhal numbers (a measure of lateral velocity to the axial velocity) at which most fish swim in nature (approx. 0.25) an attached LEV is formed, whereas at a higher Strouhal number of approximately 0.6 the LEV does not reattach. We show that the evolution of the LEV drastically alters the pressure distribution on the tail and the force it generates. We also show that the tail's delta shape is not necessary for the LEV reattachment and fish-like kinematics is capable of stabilising the LEV. Our results suggest the need for a paradigm shift in fish-like swimming research to turn the focus from the trailing edge to the leading edge of the tail. PMID:23407826

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

NASA Technical Reports Server (NTRS)

Kaltenbach, Hans-Jakob; Choi, Haecheon

1995-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

Tests of N.A.C.A. airfoils in the variable-density wind tunnel Series 24

NASA Technical Reports Server (NTRS)

Jacobs, Eastman N; WARD KENNETH E

1932-01-01

This note is the fifth of a series covering an investigation of a number of related airfoils. It presents the results obtained from tests of a group of six low-cambered airfoils in the variable-density wind tunnel. The mean camber lines are identical for the six airfoils and are of such a form that the maximum mean camber is 2 per cent of the chord and is at a position 0.4 of the chord behind the loading edge. The airfoils differ in thickness only, the maximum-thickness/chord ratios being 0.06, 0.09, 0.12, 0.15, 0.18, and 0.21. The results have been presented in the form of both infinite and finite aspect-ratio characteristics. The values of C(sub L) max/C(sub d) degrees min for this group of airfoils are among the highest thus far obtained, the minimum profile drags being approximately equal to those for the symmetrical series of corresponding thickness, while the maximum lift coefficients are considerably higher.

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

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

1998-01-01

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

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

NASA Astrophysics Data System (ADS)

Wang, Yufeng; Cai, Le; Wang, Songtao; Zhou, Xun

2018-04-01

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.

Natural laminar flow and airplane stability and control

NASA Technical Reports Server (NTRS)

Vandam, Cornelis P.

1986-01-01

Location and mode of transition from laminar to turbulent boundary layer flow have a dominant effect on the aerodynamic characteristics of an airfoil section. The influences of these parameters on the sectional lift and drag characteristics of three airfoils are examined. Both analytical and experimental results demonstrate that when the boundary layer transitions near the leading edge as a result of surface roughness, extensive trailing-edge separation of the turbulent boundary layer may occur. If the airfoil has a relatively sharp leading-edge, leading-edge stall due to laminar separation can occur after the leading-edge suction peak is formed. These two-dimensional results are used to examine the effects of boundary layer transition behavior on airplane longitudinal and lateral-directional stability and control.

The leading-edge vortex of swift-wing shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-11-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the Leading-Edge Vortex (LEV) for lift generation in a variety of flight conditions. In this investigation, a model non-slender delta shaped wing with a sharp leading-edge is tested at low Reynolds Number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the un-modified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift-wing shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds Number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta shaped wing. This work received funding from the Engineering and Physical Sciences Research Council [EP/M506515/1] and the Consejo Nacional de Ciencia y Tecnología (CONACYT).

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.

A flight test of laminar flow control leading-edge systems

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Crescentic ramp turbine stage

NASA Technical Reports Server (NTRS)

Lee, Ching-Pang (Inventor); Tam, Anna (Inventor); Kirtley, Kevin Richard (Inventor); Lamson, Scott Henry (Inventor)

2007-01-01

A turbine stage includes a row of airfoils joined to corresponding platforms to define flow passages therebetween. Each airfoil includes opposite pressure and suction sides and extends in chord between opposite leading and trailing edges. Each platform includes a crescentic ramp increasing in height from the leading and trailing edges toward the midchord of the airfoil along the pressure side thereof.

Closed loop steam cooled airfoil

DOEpatents

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

2006-04-18

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

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

NASA Technical Reports Server (NTRS)

Lawing, P. L.

1985-01-01

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

Computational fluid dynamics of airfoils and wings

NASA Technical Reports Server (NTRS)

Garabedian, P.; Mcfadden, G.

1982-01-01

It is pointed out that transonic flow is one of the fields where computational fluid dynamics turns out to be most effective. Codes for the design and analysis of supercritical airfoils and wings have become standard tools of the aircraft industry. The present investigation is concerned with mathematical models and theorems which account for some of the progress that has been made. The most successful aerodynamics codes are those for the analysis of flow at off-design conditions where weak shock waves appear. A major breakthrough was achieved by Murman and Cole (1971), who conceived of a retarded difference scheme which incorporates artificial viscosity to capture shocks in the supersonic zone. This concept has been used to develop codes for the analysis of transonic flow past a swept wing. Attention is given to the trailing edge and the boundary layer, entropy inequalities and wave drag, shockless airfoils, and the inverse swept wing code.

Navier-Stokes computations for circulation control airfoils

NASA Technical Reports Server (NTRS)

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

1987-01-01

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

Navier-Stokes computations for circulation controlled airfoils

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

Porous plug for reducing orifice induced pressure error in airfoils

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

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.

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.

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

Aerothermal Performance Constraint Analysis of Sharp Nosecaps and Leading Edges

NASA Technical Reports Server (NTRS)

Rizk, Yehia; Gee, Ken

2004-01-01

The main objective of this work is to predict the Aerothermal Performance Constraint (APC) for a class of Crew Transfer Vehicles (CTV) with shap noses and wing leading edges made out of UHTC which is a family of Ultra High Temperature Ceramics materials developed at NASA Ames. The APC is based on the theoretical temperature limit of the material which is usually encountered at the CTV nose or wing leading edge. The APC places a lower limit on the trajectory of the CTV in the altitude velocity space. The APC is used as one of the constraints in developing reentry and abort trajectories for the CTV. The trajectories are then used to generate transient thermal response of the nosecaps and wing leading edges which are represented as either a one piece of UHTC or two piece (UHTC + RCC) with perfect axial contact. The final paper will include more details about the analysis procedure and will also include results for reentry and abort design trajectories.

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.

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.

A unified viscous theory of lift and drag of 2-D thin airfoils and 3-D thin wings

NASA Technical Reports Server (NTRS)

Yates, John E.

1991-01-01

A unified viscous theory of 2-D thin airfoils and 3-D thin wings is developed with numerical examples. The viscous theory of the load distribution is unique and tends to the classical inviscid result with Kutta condition in the high Reynolds number limit. A new theory of 2-D section induced drag is introduced with specific applications to three cases of interest: (1) constant angle of attack; (2) parabolic camber; and (3) a flapped airfoil. The first case is also extended to a profiled leading edge foil. The well-known drag due to absence of leading edge suction is derived from the viscous theory. It is independent of Reynolds number for zero thickness and varies inversely with the square root of the Reynolds number based on the leading edge radius for profiled sections. The role of turbulence in the section induced drag problem is discussed. A theory of minimum section induced drag is derived and applied. For low Reynolds number the minimum drag load tends to the constant angle of attack solution and for high Reynolds number to an approximation of the parabolic camber solution. The parabolic camber section induced drag is about 4 percent greater than the ideal minimum at high Reynolds number. Two new concepts, the viscous induced drag angle and the viscous induced separation potential are introduced. The separation potential is calculated for three 2-D cases and for a 3-D rectangular wing. The potential is calculated with input from a standard doublet lattice wing code without recourse to any boundary layer calculations. Separation is indicated in regions where it is observed experimentally. The classical induced drag is recovered in the 3-D high Reynolds number limit with an additional contribution that is Reynold number dependent. The 3-D viscous theory of minimum induced drag yields an equation for the optimal spanwise and chordwise load distribution. The design of optimal wing tip planforms and camber distributions is possible with the viscous 3-D wing theory.

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

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

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

NASA Technical Reports Server (NTRS)

Howe, M. S.

1981-01-01

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

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

DTIC Science & Technology

1992-06-01

with published empirical data for tangential velocity and/or pressure coefficient distributions for the NACA 0012 and Eppler E64 airfoils before its use...tangential velocity distribution for the Eppler airfoil . No difference from the NACA 0012 Cp data could be identified. 5. Flight Regime Selection It was...37 1. P-3 Airfoil Section ...... ............ .. 37 2. Program Inputs and Outputs .. ........ .. 37 3. Program Operation

Membrane tension controls adhesion positioning at the leading edge of cells

PubMed Central

Pontes, Bruno; Gole, Laurent; Kosmalska, Anita Joanna; Tam, Zhi Yang; Luo, Weiwei; Kan, Sophie; Viasnoff, Virgile; Roca-Cusachs, Pere; Tucker-Kellogg, Lisa

2017-01-01

Cell migration is dependent on adhesion dynamics and actin cytoskeleton remodeling at the leading edge. These events may be physically constrained by the plasma membrane. Here, we show that the mechanical signal produced by an increase in plasma membrane tension triggers the positioning of new rows of adhesions at the leading edge. During protrusion, as membrane tension increases, velocity slows, and the lamellipodium buckles upward in a myosin II–independent manner. The buckling occurs between the front of the lamellipodium, where nascent adhesions are positioned in rows, and the base of the lamellipodium, where a vinculin-dependent clutch couples actin to previously positioned adhesions. As membrane tension decreases, protrusion resumes and buckling disappears, until the next cycle. We propose that the mechanical signal of membrane tension exerts upstream control in mechanotransduction by periodically compressing and relaxing the lamellipodium, leading to the positioning of adhesions at the leading edge of cells. PMID:28687667

A Numerical Evaluation of Icing Effects on a Natural Laminar Flow Airfoil

NASA Technical Reports Server (NTRS)

Chung, James J.; Addy, Harold E., Jr.

2000-01-01

As a part of CFD code validation efforts within the Icing Branch of NASA Glenn Research Center, computations were performed for natural laminar flow (NLF) airfoil, NLF-0414. with 6 and 22.5 minute ice accretions. Both 3-D ice castings and 2-D machine-generated ice shapes were used in wind tunnel tests to study the effects of natural ice is well as simulated ice. They were mounted in the test section of the Low Turbulence Pressure Tunnel (LTPT) at NASA Langley that the 2-dimensionality of the flow can be maintained. Aerodynamic properties predicted by computations were compared to data obtained through the experiment by the authors at the LTPT. Computations were performed only in 2-D and in the case of 3-D ice, the digitized ice shape obtained at one spanwise location was used. The comparisons were mainly concentrated on the lift characteristics over Reynolds numbers ranging from 3 to 10 million and Mach numbers ranging from 0.12 to 0.29. WIND code computations indicated that the predicted stall angles were in agreement with experiment within one or two degrees. The maximum lift values obtained by computations were in good agreement with those of the experiment for the 6 minute ice shapes and the minute 3-D ice, but were somewhat lower in the case of the 22.5 minute 2-D ice. In general, the Reynolds number variation did not cause much change in the lift values while the variation of Mach number showed more change in the lift. The Spalart-Allmaras (S-A) turbulence model was the best performing model for the airfoil with the 22.5 minute ice and the Shear Stress Turbulence (SST) turbulence model was the best for the airfoil with the 6 minute ice and also for the clean airfoil. The pressure distribution on the surface of the iced airfoil showed good agreement for the 6 minute ice. However, relatively poor agreement of the pressure distribution on the upper surface aft of the leading edge horn for the 22.5 minute ice suggests that improvements are needed in the grid or

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.

The Role of Separation Bubbles on the Aerodynamic Characteristics of Airfoils, Including Stall and Post-Stall, at Low Reynolds Numbers

NASA Technical Reports Server (NTRS)

Chen, Hsun H.; Cebeci, Tuncer

2007-01-01

Airfoils at high Reynolds numbers, in general, have small separation bubbles that are usually confined to the leading edge. Since the Reynolds number is large, the turbulence model for the transition region between the laminar and turbulent flow is not important. Furthermore, the onset of transition occurs either at separation or prior to separation and can be predicted satisfactorily by empirical correlations when the incident angle is small and can be assumed to correspond to laminar separation when the correlations do not apply, i.e., at high incidence angles.

The effects of leading edge modifications on the post-stall characteristics of wings

NASA Technical Reports Server (NTRS)

Winkelmann, A. E.; Barlow, J. B.; Saini, J. K.; Anderson, J. D., Jr.; Jones, E.

1980-01-01

An investigation of the effects of leading edge modifications on the post-stall characteristics of two rectangular planform wings in a series of low speed wind tunnel tests is presented. Abrupt discontinuities in the leading edge shape of the wings were produced by placing a nose glove over a portion of the span or by deflecting sections of a segmented leading edge flap. Six component balance data, oil flow visualization photographs, and pressure distribution measurements were obtained, and tests made to study the development of flow separation at stall on small scale planform wing models. Results of oil flow visualization tests at and beyond stall showed the formation of counter-rotating swirl patterns on the upper surface of the '2-D' and '3-D' wings, and results of a numerical lifting line technique applied to wings with leading edge modifications are included.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

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

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

The leading-edge vortex of swift wing-shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

The leading-edge vortex of swift wing-shaped delta wings

PubMed Central

Muir, Rowan Eveline; Arredondo-Galeana, Abel

2017-01-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing. PMID:28878968

The leading-edge vortex of swift wing-shaped delta wings.

PubMed

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus . The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

A numerical investigation of the airfoil-gust interaction noise in transonic flows: Acoustic processes

NASA Astrophysics Data System (ADS)

Zhong, Siyang; Zhang, Xin; Gill, James; Fattah, Ryu; Sun, Yuhao

2018-07-01

The sound produced by airfoil-gust interaction is a significant source of broadband noise in turbofan engines or contra-rotating open rotors (CRORs). There are competing mechanisms in this regime because of the presence of shocks that were seldom considered in the previous subsonic studies. A numerical investigation of airfoil-gust interaction noise at transonic speeds is undertaken in this work. By introducing vortical gust/synthetic turbulence to specified regions in the computational domain to interact with different elements in the flow field, it is shown that the dominant sound source is caused by leading edge-gust interaction. It is demonstrated that both streamwise and transverse disturbances interact with the near-field non-uniform mean flow and shocks can produce sound using a local gust injection method. The propagation of sound is significantly influenced by the presence of the shocks, and the far field radiation pattern is changed. We also study the effect of gust strength on the near and far field properties. The linearity is maintained for gust strength smaller than 1.0% of the mean flow velocity. Otherwise, the shocks may experience oscillations that will alter the near-field aerodynamics and far-field radiation.

Flight Tests of a Supersonic Natural Laminar Flow Airfoil

NASA Technical Reports Server (NTRS)

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

2015-01-01

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

Assessment of PIV-based unsteady load determination of an airfoil with actuated flap

NASA Astrophysics Data System (ADS)

Sterenborg, J. J. H. M.; Lindeboom, R. C. J.; Simão Ferreira, C. J.; van Zuijlen, A. H.; Bijl, H.

2014-02-01

For complex experimental setups involving movable structures it is not trivial to directly measure unsteady loads. An alternative is to deduce unsteady loads indirectly from measured velocity fields using Noca's method. The ultimate aim is to use this method in future work to determine unsteady loads for fluid-structure interaction problems. The focus in this paper is first on the application and assessment of Noca's method for an airfoil with an oscillating trailing edge flap. To our best knowledge Noca's method has not been applied yet to airfoils with moving control surfaces or fluid-structure interaction problems. In addition, wind tunnel corrections for this type of unsteady flow problem are considered.

Spanwise visualization of the flow around a three-dimensional foil with leading edge protuberances

NASA Astrophysics Data System (ADS)

Stanway, M. J.; Techet, A. H.

2006-11-01

Studies of model humpback whale fins have shown that leading edge protuberances, or tubercles, can lead to delayed stall and increased lift at higher angles of attack, compared to foils with geometrically smooth leading edges. Such enhanced performance characteristics could prove highly useful in underwater vehicles such as gliders or long range AUVs (autonomous underwater vehicles). In this work, Particle Imaging Velocimetry (PIV) is performed on two static wings in a water tunnel over a range of angles of attack. These three- dimensional, finite-aspect ratio wings are modeled after a humpback whale flipper and are identical in shape, tapered from root to tip, except for the leading edge. In one of the foils the leading edge is smooth, whereas in the other, regularly spaced leading edge bumps are machined to simulate the whale’s fin tubercles. Results from these PIV tests reveal distinct cells where coherent flow structures are destroyed as a result of the leading edge perturbations. Tests are performed at Reynolds numbers Re ˜ O(10^5), based on chordlength, in a recirculating water tunnel. An inline six-axis load cell is mounted to measure the forces on the foil over a range of static pitch angles. It is hypothesized that this spanwise breakup of coherent vortical structures is responsible for the delayed angle of stall. These quantitative experiments complement exiting qualitative studies with two dimensional foils.

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.

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

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

Modeling of Electron Transpiration Cooling for Leading Edges of Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Hanquist, Kyle Matthew

The development of aeronautics has been largely driven by the passion to fly faster. From the flight of the Wright Flyer that flew 48 km/hr to the recent advances in hypersonic flight, most notably NASA's X-43A that flew at over 3 km/s, the velocity of flight has steadily increased. However, as these hypersonic speeds are reached and increased, contradicting aerothermodynamic design requirements present themselves. For example, a hypersonic cruise vehicle requires sharp leading edges to decrease the drag in order to maximize the range. However, the aerodynamic performance gains obtained by having a sharp leading edge come at the cost of very high, localized heating rates. There is currently no ideal way to manage these heating loads for sustained hypersonic flight, especially as flight velocities continue to increase. An approach that has been recently proposed involves using thermo-electric materials on these sharp leading edges to manage the heating loads. When exposed to high convective heating rates, these materials emit a current of electrons that leads to a cooling effect of the surface of the vehicle called electron transpiration cooling (ETC). This dissertation focuses on developing a modeling approach to investigate this phenomenon. The research includes developing and implementing an approach for ETC into a computational fluid dynamics code for simulation of hypersonic flow that accounts for electron emission from the surface. Models for space-charge-limited emission are also developed and implemented in order to accurately determine the level of emission from the surface. This work involves developing analytic models and assessing them using a direct-kinetic plasma sheath solver. Electric field effects are also implemented in the modeling approach, which accounts for forced diffusion and Joule heating. Finally, the modeling approach is coupled to a material response code in order to model the heat transfer into the material surface. Using this modeling

Unsteady lift and thrust of a 2D flapping thin airfoil in the presence of additional leading edge vortices

NASA Astrophysics Data System (ADS)

Alaminos-Quesada, Javier; Fernandez-Feria, Ramon

2017-11-01

The effect of leading-edge vortices (LEVs) on the lift, thrust and moment of a two-dimensional heaving and pitching foil is analyzed from the unsteady, linear potential theory. General expressions taking into account the effect of unsteady point vortices interacting with the oscillatory trailing wake are first derived. Then, simplified expressions for the initial stages of the growing LEV on each half-stroke are used to obtain analytical closed expressions for the main contribution of these vortices to the lift, thrust and moment. It is found that, within the linear potential framework and the Brown-Michael model, the LEV contributes to the aerodynamic forces and moment only for combined pitching and heaving motions of the foil, being a relevant contribution for sufficiently large values of the product of the reduced frequency and the amplitude of the heaving and/or pitching motions. The results are compared with available experimental data and numerical simulations. Supported by the Ministerio de Economia y Competitividad of Spain Grants No. DPI2013-40479-P and DPI2016-76151-C2-1-R.

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.

Nozzle airfoil having movable nozzle ribs

DOEpatents

Yu, Yufeng Phillip; Itzel, Gary Michael

2002-01-01

A nozzle vane or airfoil structure is provided in which the nozzle ribs are connected to the side walls of the vane or airfoil in such a way that the ribs provide the requisite mechanical support between the concave side and convex side of the airfoil but are not locked in the radial direction of the assembly, longitudinally of the airfoil. The ribs may be bi-cast onto a preformed airfoil side wall structure or fastened to the airfoil by an interlocking slide connection and/or welding. By attaching the nozzle ribs to the nozzle airfoil metal in such a way that allows play longitudinally of the airfoil, the temperature difference induced radial thermal stresses at the nozzle airfoil/rib joint area are reduced while maintaining proper mechanical support of the nozzle side walls.

Experimental evaluation of joint designs for a space-shuttle orbiter ablative leading edge

NASA Technical Reports Server (NTRS)

Tompkins, S. S.; Kabana, W. P.

1975-01-01

The thermal performance of two types of ablative leading-edge joints for a space-shuttle orbiter were tested and evaluated. Chordwise joints between ablative leading-edge segments, and spanwise joints between ablative leading-edge segments and reusable surface insulation tiles were exposed to simulated shuttle heating environments. The data show that the thermal performance of models with chordwise joints to be as good as jointless models in simulated ascent-heating and orbital cold-soak environments. The suggestion is made for additional work on the joint seals, and, in particular, on the effects of heat-induced seal-material surface irregularities on the local flow.

Darrieus wind-turbine airfoil configurations

NASA Astrophysics Data System (ADS)

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

1982-06-01

The purpose was to determine what aerodynamic performance improvement, if any, could be achieved by judiciously choosing the airfoil sections for Darrieus wind turbine blades. Ten different airfoils, having thickness to chord ratios of twelve, fifteen and eighteen percent, were investigated. Performance calculations indicated that the NACA 6-series airfoils yield peak power coefficients at least as great as the NACA. Furthermore, the power coefficient-tip speed ratio curves were broader and flatter for the 6-series airfoils. Sample calculations for an NACA 63 sub 2-015 airfoil showed an annual energy output increase of 17 to 27% depending upon rotor solidity, compared to an NACA 0015 airfoil. An attempt was made to account for the flow curvature effects associated with Darrieus turbines by transforming the NACA 63 sub 2-015 airfoil to an appropriate shape.

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

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

PubMed Central

Laplante, Caroline

2011-01-01

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

Aerodynamic robustness in owl-inspired leading-edge serrations: a computational wind-gust model.

PubMed

Rao, Chen; Liu, Hao

2018-06-08

Owls are a master to achieve silent flight in gliding and flapping flights under natural turbulent environments owing to their unique wing morphologies. While the leading-edge serrations are recently revealed, as a passive flow control micro-device, to play a crucial role in aerodynamic force production and sound suppression [25], the characteristics of wind-gust rejection associated with leading-edge serrations remain unclear. Here we address a large-eddy simulation (LES)-based study of aerodynamic robustness in owl-inspired leading-edge serrations, which is conducted with clean and serrated wing models through mimicking wind-gusts under a longitudinal fluctuation in free-stream inflow and a lateral fluctuation in pitch angle over a broad range of angles of attack (AoAs) over 0° ≤ Φ ≤ 20°. Our results show that the leading-edge serration-based passive flow control mechanisms associated with laminar-turbulent transition work effectively under fluctuated inflow and wing pitch, indicating that the leading-edge serrations are of potential gust fluctuation rejection or robustness in aerodynamic performance. Moreover, it is revealed that the tradeoff between turbulent flow control (i.e., aero-acoustic suppression) and force production in the serrated model holds independently to the wind-gust environments: poor at lower AoAs but capable of achieving equivalent aerodynamic performance at higher AoAs > 15o compared to the clean model. Our results reveal that the owl-inspired leading-edge serrations can be a robust micro-device for aero-acoustic control coping with unsteady and complex wind environments in biomimetic rotor designs for various fluid machineries. © 2018 IOP Publishing Ltd.

An approach to constrained aerodynamic design with application to airfoils

NASA Technical Reports Server (NTRS)

Campbell, Richard L.

1992-01-01

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

Modern Airfoil Ice Accretions

NASA Technical Reports Server (NTRS)

Addy, Harold E., Jr.; Potapczuk, Mark G.; Sheldon, David W.

1997-01-01

This report presents results from the first icing tests performed in the Modem Airfoils program. Two airfoils have been subjected to icing tests in the NASA Lewis Icing Research Tunnel (IRT). Both airfoils were two dimensional airfoils; one was representative of a commercial transport airfoil while the other was representative of a business jet airfoil. The icing test conditions were selected from the FAR Appendix C envelopes. Effects on aerodynamic performance are presented including the effects of varying amounts of glaze ice as well as the effects of approximately the same amounts of glaze, mixed, and rime ice. Actual ice shapes obtained in these tests are also presented for these cases. In addition, comparisons are shown between ice shapes from the tests and ice shapes predicted by the computer code, LEWICE for similar conditions. Significant results from the tests are that relatively small amounts of ice can have nearly as much effect on airfoil lift coefficient as much greater amounts of ice and that glaze ice usually has a more detrimental effect than either rime or mixed ice. LEWICE predictions of ice shapes, in general, compared reasonably well with ice shapes obtained in the IRT, although differences in details of the ice shapes were observed.

Experimental investigation of the flowfield of an oscillating airfoil

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

Experimental investigation of the flowfield of an oscillating airfoil

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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.

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.

Shape optimization for aerodynamic efficiency and low observability

NASA Technical Reports Server (NTRS)

Vinh, Hoang; Van Dam, C. P.; Dwyer, Harry A.

1993-01-01

Field methods based on the finite-difference approximations of the time-domain Maxwell's equations and the potential-flow equation have been developed to solve the multidisciplinary problem of airfoil shaping for aerodynamic efficiency and low radar cross section (RCS). A parametric study and an optimization study employing the two analysis methods are presented to illustrate their combined capabilities. The parametric study shows that for frontal radar illumination, the RCS of an airfoil is independent of the chordwise location of maximum thickness but depends strongly on the maximum thickness, leading-edge radius, and leadingedge shape. In addition, this study shows that the RCS of an airfoil can be reduced without significant effects on its transonic aerodynamic efficiency by reducing the leading-edge radius and/or modifying the shape of the leading edge. The optimization study involves the minimization of wave drag for a non-lifting, symmetrical airfoil with constraints on the airfoil maximum thickness and monostatic RCS. This optimization study shows that the two analysis methods can be used effectively to design aerodynamically efficient airfoils with certain desired RCS characteristics.

Control of unsteady separated flow associated with the dynamic pitching of airfoils

NASA Technical Reports Server (NTRS)

Ahmed, Sajeer

1991-01-01

Although studies have been done to understand the dependence of parameters for the occurrence of deep stall, studies to control the flow for sustaining lift for a longer time has been little. To sustain the lift for a longer time, an understanding of the development of the flow over the airfoil is essential. Studies at high speed are required to study how the flow behavior is dictated by the effects of compressibility. When the airfoil is pitched up in ramp motion or during the upstroke of an oscillatory cycle, the flow development on the upper surface of the airfoil and the formation of the vortex dictates the increase in lift behavior. Vortex shedding past the training edge decreases the lift. It is not clear what is the mechanism associated with the unsteady separation and vortex formation in present unsteady environment. To develop any flow control device, to suppress the vortex formation or delay separation, it is important that this mechanism be properly understood. The research activities directed toward understanding these questions are presented and the results are summarized.

Pressure distribution over an NACA 23012 airfoil with an NACA 23012 external-airfoil flap

NASA Technical Reports Server (NTRS)

Wenzinger, Carl J

1938-01-01

Report presents the results of pressure-distribution tests of an NACA 23012 airfoil with an NACA 23012 external airfoil flap made in the 7 by 10-foot wind tunnel. The pressures were measured on the upper and lower surfaces at one chord section on both the main airfoil and on the flap for several different flap deflections and at several angles of attack. A test installation was used in which the airfoil was mounted horizontally in the wind tunnel between vertical end planes so that two-dimensional flow was approximated. The data are presented in the form of pressure-distribution diagrams and as graphs of calculated coefficients for the airfoil-and-flap combination and for the flap alone.

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Covell, Peter F.; Wood, Richard M.; Miller, David S.

1987-01-01

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

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.

Ablative overlays for Space Shuttle leading edge ascent heat protection

NASA Technical Reports Server (NTRS)

Strauss, E. L.

1975-01-01

Ablative overlays were evaluated via a plasma-arc simulation of the ascent pulse on the leading edge of the Space Shuttle Orbiter. Overlay concepts included corkboard, polyisocyanurate foam, low-density Teflon, epoxy, and subliming salts. Their densities ranged from 4.9 to 81 lb per cu ft, and the thicknesses varied from 0.107 to 0.330 in. Swept-leading-edge models were fabricated from 30-lb per cu ft silicone-based ablators. The overlays were bonded to maintain the surface temperature of the base ablator below 500 F during ascent. Foams provided minimum-weight overlays, and subliming salts provided minimum-thickness overlays. Teflon left the most uniform surface after ascent heating.

Airfoil profiles for minimum pressure drag at supersonic velocities -- general analysis with application to linearized supersonic flow

NASA Technical Reports Server (NTRS)

Chapman, Dean R

1952-01-01

A theoretical investigation is made of the airfoil profile for minimum pressure drag at zero lift in supersonic flow. In the first part of the report a general method is developed for calculating the profile having the least pressure drag for a given auxiliary condition, such as a given structural requirement or a given thickness ratio. The various structural requirements considered include bending strength, bending stiffness, torsional strength, and torsional stiffness. No assumption is made regarding the trailing-edge thickness; the optimum value is determined in the calculations as a function of the base pressure. To illustrate the general method, the optimum airfoil, defined as the airfoil having minimum pressure drag for a given auxiliary condition, is calculated in a second part of the report using the equations of linearized supersonic flow.

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.

Local heat-transfer measurements on a large, scale-model turbine blade airfoil using a composite of a heater element and liquid crystals

NASA Technical Reports Server (NTRS)

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

1985-01-01

Local heat transfer coefficients were experimentally mapped along the midchord of a five-time-size turbine blade airfoil in a static cascade operated at room temperature over a range of Reynolds numbers. The test surface consisted of a composite of commercially available materials: a mylar sheet with a layer of cholesteric liquid crystals, that change color with temperature, and a heater sheet made of a carbon-impregnated paper, that produces uniform heat flux. After the initial selection and calibration of the composite sheet, accurate, quantitative, and continuous heat transfer coefficients were mapped over the airfoil surface. The local heat transfer coefficients are presented for Reynolds numbers from 2.8 x 10 to the 5th power to 7.6 x 10 to the 5th power. Comparisons are made with analytical values of heat transfer coefficients obtained from the STAN5 boundary layer code. Also, a leading edge separation bubble was revealed by thermal and flow visualization.

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.

Boundary layer relaminarization device

NASA Technical Reports Server (NTRS)

Creel, Theodore R. (Inventor)

1992-01-01

Relamination of a boundary layer formed in supersonic flow over the leading edge of a swept airfoil is accomplished by means of at least one band, especially a quadrangular band, and most preferably a square band. Each band conforms to the leading edge and the upper and lower surfaces of the airfoil as an integral part thereof and extends perpendicularly from the leading edge. Each band has a height of about two times the thickness of the maximum expected boundary layer.

Active flow control of the laminar separation bubble on a plunging airfoil near stall

NASA Astrophysics Data System (ADS)

Pande, Arth; Agate, Mark; Little, Jesse; Fasel, Hermann

2017-11-01

The effects of small amplitude (A/c = 0.048) high frequency (πfc/U∞ = 0.70) plunging motion on the X-56A airfoil are examined experimentally at Re = 200,000 for 12° angle of attack (CL,MAX = 12.25°) . The purpose of this research is to study the aerodynamic influence of structural motion when the wing is vibrating close to its eigenfrequency near static stall. Specific focus is placed on the laminar separation bubble (LSB) near the leading edge and its control via plasma actuation. In the baseline case, the leading edge bubble bursts during the oscillation cycle causing moment stall. A collaborative computational effort has shown that small amplitude forcing at a frequency that is most amplified by the primary instability of the LSB (FLSB+= 1, Fc+= 52) generates coherent spanwise vortices that entrain freestream momentum, thus reducing separation all while maintaining a laminar flow state. Results (PIV and surface pressure) indicate that a similar control mechanism is effective in the experiments. This is significant given the existence of freestream turbulence in the wind tunnel which has been shown to limit the efficacy of this active flow control technique in a model problem using Direct Numerical Simulation. The implications of these results are discussed.

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Martín-Alcántara, A.; Fernandez-Feria, R.; Sanmiguel-Rojas, E.

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

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.

Non-linear unsteady wing theory, part 1. Quasi two-dimensional behavior: Airfoils and slender wings

NASA Technical Reports Server (NTRS)

Mccune, J. E.

1987-01-01

The initial phases of a study of the large-amplitude unsteady aerodynamics of wings in severe maneuver are reported. The research centers on vortex flows, their initiation at wing surfaces, their subsequent convection, and interaction dynamically with wings and control surfaces. The focus is on 2D and quasi-2D aspects of the problem and features the development of an exact nonlinear unsteady airfoil theory as well as an approach to the crossflow problem for slender wing applications including leading-edge separation. The effective use of interactive on-line computing in quantifying and visualizing the nonsteady effects of severe maneuver is demonstrated. Interactive computational work is now possible, in which a maneuver can be initiated and its effects observed and analyzed immediately.

Airfoil Drag Reduction using Controlled Trapped Vorticity Concentrations

NASA Astrophysics Data System (ADS)

Desalvo, Michael; Glezer, Ari

2017-11-01

The aerodynamic performance of a lifting surface at low angles of attack (when the base flow is fully attached) is improved through fluidic modification of its ``apparent'' shape by superposition of near-surface trapped vorticity concentrations. In the present wind tunnel investigations, a controlled trapped vorticity concentration is formed on the pressure surface of an airfoil (NACA 4415) using a hybrid actuator comprising a passive obstruction of scale O(0.01c) and an integral synthetic jet actuator. The jet actuation frequency [Stact O(10)] is selected to be at least an order of magnitude higher than the characteristic unstable frequency of the airfoil wake, thereby decoupling the actuation from the global instabilities of the base flow. Regulation of vorticity accumulation in the vicinity of the actuator by the jet effects changes in the local pressure, leading in turn to changes in the airfoil's drag and lift. Trapped vorticity can lead to a significant reduction in drag and reduced lift (owing to the sense of the vorticity), e.g. at α =4° and Re = 6.7 .105 the drag and lift reductions are 14% and 2%, respectively. PIV measurements show the spatial variation in the distribution of vorticity concentrations and yield estimates of the corresponding changes in circulation.

Optimization of Wind Turbine Airfoils/Blades and Wind Farm Layouts

NASA Astrophysics Data System (ADS)

Chen, Xiaomin

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

The influence of sweep on the aerodynamic loading of an oscillating NACA 0012 airfoil. Volume 1: Technical report

NASA Technical Reports Server (NTRS)

St.hilaire, A. O.; Carta, F. O.; Fink, M. R.; Jepson, W. D.

1979-01-01

Aerodynamic experiments were performed on an oscillating NACA 0012 airfoil utilizing a tunnel-spanning wing in both unswept and 30 degree swept configurations. The airfoil was tested in steady state and in oscillatory pitch about the quarter chord. The unsteady aerodynamic loading was measured using pressure transducers along the chord. Numerical integrations of the unsteady pressure transducer responses were used to compute the normal force, chord force, and moment components of the induced loading. The effects of sweep on the induced aerodynamic load response was examined. For the range of parameters tested, it was found that sweeping the airfoil tends to delay the onset of dynamic stall. Sweeping was also found to reduce the magnitude of the unsteady load variation about the mean response. It was determined that at mean incidence angles greater than 9 degrees, sweep tends to reduce the stability margin of the NACA 0012 airfoil; however, for all cases tested, the airfoil was found to be stable in pure pitch. Turbulent eddies were found to convect downstream above the upper surface and generate forward-moving acoustic waves at the trailing edge which move upstream along the lower surface.

An experimental study of a supercritical trailing-edge flow

NASA Technical Reports Server (NTRS)

Brown, J. L.; Viswanath, P. R.

1984-01-01

An experimental study has been conducted of a transonic, turbulent, high-Reynolds-number blunt trailing-edge flow. The model shape and the surface pressure distribution are characteristics of a modern supercritical airfoil under shock-free conditions. Reynolds number and pressure gradient scaling of the boundary layer are relevant to airfoil applications. The data set is exceptionally accurate and consistent, with the momentum balance accounting for the flux of momentum to within 1 percent, except in the immediate vicinity of the blunt trailing edge. The experimental flow exhibits strong viscous-inviscid interaction and higher-order boundary-layer effects including strong adverse streamwise pressure gradient, significant normal pressure gradients associated with surface and streamline curvature, and significant wake curvature. Navier-Stokes calculations with a two-equation K-epsilon turbulence model predict the correct pressure distribution which demonstrates the utility of these engineering tools. The experiment approaches separation at the strailing edge. However, in comparison to the experiment, the calculations predict too high skin friction and insufficient displacement thickness growth. An analysis of the turbulent and mean flow fields reveals the turbulence model defects are likely in modeling the dissipation source and sink terms, and in the eddy viscosity relation.

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

DTIC Science & Technology

2007-06-01

3rd International Symposium on Integrating CFD and Experiments in Aerodynamics 20-21 June 2007 U.S. Air Force Academy, CO, USA Static Extended...is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge...basic configuration, has a good potential to improve the cruise flight efficiency. Key words: trailing edge, airfoil, wing, lift, drag, aerodynamics

Effects of Angle of Attack and Velocity on Trailing Edge Noise

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, Thomas F.

2006-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1982-01-01

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

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

NASA Technical Reports Server (NTRS)

Krothapalli, A.; Leopold, D.

1984-01-01

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

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).

The Flow Field on Hydrofoils with Leading Edge Protuberances

NASA Astrophysics Data System (ADS)

Custodio, Derrick; Henoch, Charles; Johari, Hamid

2009-11-01

The exceptional mobility of the humpback whale has been linked to the use of its unique pectoral flippers. Biologists speculate that the flippers leading edge protuberances are a form of passive flow control. Force measurements on 2D hydrofoils with spanwise uniform leading edge protuberances, resembling those seen on the humpback whale flipper, were taken in a water tunnel and have revealed performance modifications when compared to a baseline NACA 63(4)-021 hydrofoil model. Qualitative flow visualization techniques and Particle Image Velocimetry (PIV) flow field measurements on the modified hydrofoils have shown that streamwise vortices originating from the shoulders of the protuberances are the likely cause of performance changes. Varying levels of interaction among adjacent streamwise vortices have been observed as a function of angle of attack and chord location. The circulation of these vortices as a function of angle of attack and spatial location was measured and an analysis of the vortex interactions will be presented.

Detailed comparison between DNS and wind tunnel experiment for an airfoil at Re = 20,000 with a view towards control

NASA Astrophysics Data System (ADS)

Tank, Joseph; Jacobs, Gustaaf; Spedding, Geoffrey

2017-11-01

The reduction in size and weight of electronic devices in recent years has enabled the use of small flying devices that operate at Re <1.5 x 105 for a variety of applications. At these low Re, the boundary layer often separates before the trailing edge, even at low angles of attack, leading to aerodynamic behaviors that are not predicted by classical inviscid theories. There is currently no comprehensive database of airfoil data in this Re regime, where the sensitivity of the boundary layer behavior to small disturbances in the free stream often leads to discrepancies between results generated in different facilities. Here we provide experimental results generated in a wind tunnel with a low turbulence intensity for a NACA 65(1)-412 airfoil at Re = 2 x 104. Several unexpected phenomena are observed in force balance results and explanations are proposed based on PIV flow visualization. Qualitative and quantitative comparisons are made with results from a DNS code using higher-order discontinuous Galerkin methods. Internal acoustic forcing at locations dictated by Lagrangian Coherent Structure behavior is explored as a potential closed loop flow control strategy. Support from AFOSR Grant# FA9550-16-1-0392 under Dr Doug Smith is most gratefully acknowledged.

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.

Potential environmental effects of the leading edge hydrokinetic energy technology.

DOT National Transportation Integrated Search

2017-05-01

The Volpe Center evaluated potential environmental challenges and benefits of the ARPA-E funded research project, Marine Hydrokinetic Energy Harvesting Using Cyber-Physical Systems, led by Brown University. The Leading Edge research team develo...

Airfoil shape for a turbine bucket

DOEpatents

Hyde, Susan Marie; By, Robert Romany; Tressler, Judd Dodge; Schaeffer, Jon Conrad; Sims, Calvin Levy

2005-06-28

Third stage turbine buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth Table I wherein X and Y values are in inches and the Z values are non-dimensional values from 0 to 0.938 convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape. The X and Y distances may be scalable as a function of the same constant or number to provide a scaled up or scaled down airfoil section for the bucket. The nominal airfoil given by the X, Y and Z distances lies within an envelop of .+-.0.150 inches in directions normal to the surface of the airfoil.

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.

Modifications of W and Mo leading edges under plasma loads in DIII-D divertor

NASA Astrophysics Data System (ADS)

Rudakov, D. L.; Bykov, I.; Moyer, R. A.; Abrams, T.; Chrobak, C. P.; Guo, H. Y.; Stahl, B.; Thomas, D. M.; Barton, J. L.; Nygren, R. E.; Watkins, J. G.; Lasnier, C. J.; Litnovsky, Andrey; Stangeby, P. C.; Unterberg, E. A.

2017-10-01

Cracking and melting of W and Mo leading edges were observed in the lower divertor of DIII-D during experiments with intentionally misaligned W monoblocks (MBs) and in the course of the Metal Rings Campaign involving W-coated Mo tile inserts (TIs). MBs were exposed near the attached outer strike point during deuterium and helium L- and H-mode discharges using DiMES. Two of the MBs were misaligned by 0.3 mm and 1 mm, forming leading edges. Particulate ejection from a 1 mm leading edge was observed during the exposure, and evidence of melting and cracking was found post mortem. Two toroidal rings of TIs were installed in the lower outer divertor, the inner one at the floor and the outer one at the shelf. The floor TIs bowed during plasma exposure forming leading edges up to 1.2 mm high; about 40% of these edges experienced melting. Re-solidified melt layers up to 1 mm thick were observed, their shape being consistent with motion in the jx B direction with j driven by electron emission. Work supported by US DOE under DE-FC02-04ER54698, DE-FG02-07ER54917, DE-AC04-94AL85000, DE-AC52-07NA27344, and DE-AC05-00OR22725.

Chronophin coordinates cell leading edge dynamics by controlling active cofilin levels

PubMed Central

Delorme-Walker, Violaine; Seo, Ji-Yeon; Gohla, Antje; Fowler, Bruce; Bohl, Ben; DerMardirossian, Céline

2015-01-01

Cofilin, a critical player of actin dynamics, is spatially and temporally regulated to control the direction and force of membrane extension required for cell locomotion. In carcinoma cells, although the signaling pathways regulating cofilin activity to control cell direction have been established, the molecular machinery required to generate the force of the protrusion remains unclear. We show that the cofilin phosphatase chronophin (CIN) spatiotemporally regulates cofilin activity at the cell edge to generate persistent membrane extension. We show that CIN translocates to the leading edge in a PI3-kinase–, Rac1-, and cofilin-dependent manner after EGF stimulation to activate cofilin, promotes actin free barbed end formation, accelerates actin turnover, and enhances membrane protrusion. In addition, we establish that CIN is crucial for the balance of protrusion/retraction events during cell migration. Thus, CIN coordinates the leading edge dynamics by controlling active cofilin levels to promote MTLn3 cell protrusion. PMID:26324884

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

DOE Office of Scientific and Technical Information (OSTI.GOV)

Griffith, Daniel; Richards, Phillip William

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

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

NASA Technical Reports Server (NTRS)

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

2008-01-01

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

Numerical design of advanced multi-element airfoils

NASA Technical Reports Server (NTRS)

Mathias, Donovan L.; Cummings, Russell M.

1994-01-01

The current study extends the application of computational fluid dynamics to three-dimensional high-lift systems. Structured, overset grids are used in conjunction with an incompressible Navier-Stokes flow solver to investigate flow over a two-element high-lift configuration. The computations were run in a fully turbulent mode using the one-equation Baldwin-Barth turbulence model. The geometry consisted of an unswept wing which spanned a wind tunnel test section. Flows over full and half-span Fowler flap configurations were computed. Grid resolution issues were investigated in two dimensional studies of the flapped airfoil. Results of the full-span flap wing agreed well with experimental data and verified the method. Flow over the wing with the half-span was computed to investigate the details of the flow at the free edge of the flap. The results illustrated changes in flow streamlines, separation locations, and surface pressures due to the vortex shed from the flap edge.

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.

Robust, Optimal Subsonic Airfoil Shapes

NASA Technical Reports Server (NTRS)

Rai, Man Mohan

2014-01-01

A method has been developed to create an airfoil robust enough to operate satisfactorily in different environments. This method determines a robust, optimal, subsonic airfoil shape, beginning with an arbitrary initial airfoil shape, and imposes the necessary constraints on the design. Also, this method is flexible and extendible to a larger class of requirements and changes in constraints imposed.

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

NASA Technical Reports Server (NTRS)

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

2007-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

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.

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.

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.

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.

Owl-inspired leading-edge serrations play a crucial role in aerodynamic force production and sound suppression.

PubMed

Rao, Chen; Ikeda, Teruaki; Nakata, Toshiyuki; Liu, Hao

2017-07-04

Owls are widely known for silent flight, achieving remarkably low noise gliding and flapping flights owing to their unique wing morphologies, which are normally characterized by leading-edge serrations, trailing-edge fringes and velvet-like surfaces. How these morphological features affect aerodynamic force production and sound suppression or noise reduction, however, is still not well known. Here we address an integrated study of owl-inspired single feather wing models with and without leading-edge serrations by combining large-eddy simulations (LES) with particle-image velocimetry (PIV) and force measurements in a low-speed wind tunnel. With velocity and pressure spectra analysis, we demonstrate that leading-edge serrations can passively control the laminar-turbulent transition over the upper wing surface, i.e. the suction surface at all angles of attack (0°  <  AoA  <  20°), and hence play a crucial role in aerodynamic force and sound production. We find that there exists a tradeoff between force production and sound suppression: serrated leading-edges reduce aerodynamic performance at lower AoAs  <  15° compared to clean leading-edges but are capable of achieving both noise reduction and aerodynamic performance at higher AoAs  >  15° where owl wings often reach in flight. Our results indicate that the owl-inspired leading-edge serrations may be a useful device for aero-acoustic control in biomimetic rotor designs for wind turbines, aircrafts, multi-rotor drones as well as other fluid machinery.

Turbine airfoil to shround attachment

DOEpatents

Campbell, Christian X; Morrison, Jay A; James, Allister W; Snider, Raymond G; Eshak, Daniel M; Marra, John J; Wessell, Brian J

2014-05-06

A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.

Multiple piece turbine airfoil

DOEpatents

Kimmel, Keith D; Wilson, Jr., Jack W.

2010-11-02

A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of dog bone struts each mounted within openings formed within the shell and spar to allow for relative motion between the spar and shell in the airfoil chordwise direction while also forming a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure.

Cooling circuit for a gas turbine bucket and tip shroud

DOEpatents

Willett, Fred Thomas

2004-07-13

An open cooling circuit for a gas turbine airfoil and associated tip shroud includes a first group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a leading edge of the airfoil; a second group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a trailing edge of the airfoil. A common plenum is formed in the tip shroud in direct communication with the first and second group of cooling holes, but a second plenum may be provided for the second group of radial holes. A plurality of exhaust holes extends from the plenum(s), through the tip shroud and opening along a peripheral edge of the tip shroud.

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.

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.